3 1176 00156 1985 - ntrs.nasa.gov

3 1176 00156 1985" NASA-CR-144887

1.9800006814NASA Contractor Report 144887

•3

,: ANALYSES AND TESTS OF THE B-1 AIRCRAFT

STRUCTURAL MODE CONTROL SYSTEM

John H. Wykes, Thomas R. Byar, Cary J. MacMiller,and David C. Greek

Contract NAS4-2519 '-January 1980

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Wgkes, d. H.Analgses and tests o_ the B-1 aircraft structuralmode control system.NASA

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NASA Contractor Report 144887

ANALYSES AND TESTS OF THE B-1 AIRCRAFT

STRUCTURAL MODE CONTROL SYSTEM

John H. Wykes, Thomas R. Byar, Cary J. MaeMiller,and David C. GreekRockwell InternationalEl Segundo, California

Prepared forDryden Flight Research Centerunder Contract NAS4-2519

NIA ivvo- i 73NationalAeronauticsandSpaceAdministration

1980

TABLE OF CONTENTS

Page

SU_tIARY I

IKI'RODUCTION 1

FLEXIBLEAIRCRAFT EQUATIONSOF MOTION 3

FLEXIBLEAIRCRAFTANALYSESMODEL 22

DynamicAnalysis System 22Free-FreeVibrationModal Data 24

AerodynamicData 27Control-SurfacesInertiaReactionForces 38

Active ControlSystems 42

COMPARISONSOF ANALYSESAND FLIGHT-TESTRESULTS 42

FORWARDSMCS SENSOR PACKAGERELOCATION 58

TRUNCATEDANALYTICALMODELS 78

ANALYSISOF SNCS VANEAERODYI_.IICINTERFERENCEEFFECT 90

IMPACTOF SHCSON SELECTEDLOADS 101

Backgroundof Using SMCS in the B-I FatigueAnalysis I01Gust Loads Analysis Description 102Gust Loads StructuralModel 103"

Gust Loads Aerodynamics 104Gust Loads ControlSystems 106Load Method Discussion 106Gust StatisticalLoad Calculations II0

Load Phasing IIi

ConditionMatching 112-_ ExampleConditionLoad Results 113

SMCS VANE EFFECT ON INLET/ENGINECHARACTERISTICS 154

Test Description 155

Flight-TestInstrumentation 159

i

TABLE OF CONTENTS - (Continued)

Page

Test Results 165

OscillatingVanes 167StaticVane Deflections 170

Maneuvers 170

Engine ThrottleTransients 195

Summary 205

SUMMARYOF SMCS FLIGHT TEST RESULTS 205

SMCS Performancein Time-History-DataFormat 206SMCS Performancein PSD-Data Format 213

SMCS and HandlingQualities 213SMCS High-GainTests 218Crew Evaluationsof SMCS Effects 220

OverallRide Quality 220Flight-PathControlTasks 225Non-Flight-PathControlTasks 225Readabilityof Instrumentsand Displays 225Reaching/UsingControls 226Crew Fatigue 227Motion Sickness 227

PhysicalDiscomfort 228AdditionalRide QualityObservations 228Handling Qualities 229

APPENDIX - NOMENCLATURE. 230

General 230

Ride QualityEquationsof Motion Related 234Load Equationsof Motion Related 247Engine/InletRelated 250

REFERENCES 253

ii

LIST OF ILLUSTRATIONS

Figure Title Page

1 B-I aircraftwith wings swept aft. 22 Sign conventionfor rotationaland linear rates and

accelerations. 15

3 Sign conventionfor coefficientsand control surfacedeflections. 16

4 Structuralmode deflections,slopes,and generalizedforces sign convention. 17

5 Angle-of-attack(_) and sideslip QS) definitions. 196 Euler angle definitionsand rotationsequence. 207 Dynamicanalysissystem. 238 Typical elasticaxes and mass point locations. 259 Fuselageelasticaxis refinement. 26i0 Typical symmetricmode vector plot. 29ii Typical antisymmetricmode vector plot. 3012 Typical structuralmode deflectiongrid points. 3113 Panellingand box grid for Doublet Latticeaerodynamics

program. 3214 Typical low-frequencyadjustmentmade to analytical

frequency-dependentaerodynamicdata. 3515 Typical aeroelasticflexible-to-rigidratio data for

aerodynamiccoefficientsas a functionofparticipatingstructuralmodes. 39

16 Typicalcontrolsurface inertiareactiongeneralizedforces. 41

17 Pitch axis SCAS analyticalmodel. 4318 Yaw axis SCAS analyticalmodel. 4419 Roll axis SCAS analyticalmodel. 45

20 Vertical SMCS analyticalmodel. 4621 LateralSMCS analyticalmodel. 4722 Comparisonof flight test and analyticaldata,

frequencyresponseof normal load factor atFS 571.5 (225)due to SMCS vane deflection,SCAS off, SMCS off. 49

23 Comparisonof flight test and analyticaldata,frequencyresponseof normal load factor atFS 571.5 (225)due to SMCS vane deflection,

SCAS on, SMCS off. 5024 Comparisonof flight test and analyticaldata,

frequencyresponseof normal load factor atFS 571.5 (225)due to SMCS deflection,SCASon, SMCS on. 51

iii

LIST OF ILLUSTRATIONS- (Continued)

Figure Title Page

25 Comparisonof analyticalSMCS actuatormodels toflightand simulatortest data. 53

26 Effect of test-derivedSMCS actuatormodel, frequencyresponseof normal load factor at FS 571.5 (225)due to SMCS vane deflection,SCAS on, SMCS on. 54

"'27 Comparisonof flight test and analyticaldata,frequencyresponseof lateralload factor at FS 571.5(225)due to SMCS vane deflection,SCAS off, SMCS off. 55

28 Comparisonof flight test and analyticaldata, frequencyresponseof lateralload factor at FS 571.5 (225)dueto SMCS vane deflection,SCAS on, SMCS off. 56

29 Comparisonof flight test and analyticaldata, frequencyresponseof lateralload factor at FS 571.5 (225)dueto SMCS vane deflection,SCAS on, SMCS on. 57

30 Effect of SMCS lateralgain on power spectraldensityof lateral load factor at pilot stationFS 746.8 (294). 59

31 SMCS sensorpackage locationsand couplingcharacteristics. 60

32 Effect of SMCS sensor package locationon lateralloadfactor at FS 571.5 (225)due to differentialvanedeflection,analyticaldata. 61

33 First fuselageverticalbendingmode dampingversusSMCS gains for originaland relocatedforwardsensorpackage. 63

34 SMCS vertical axis performancewith relocatedforwardsensorpackage, frequencyresponseof normal loadfactor at FS 515.6 (203)due to SMCS vane deflection,case I. 65

35 SMCS lateralaxis performancewith relocatedforwardsensorpackage, frequencyresponseof lateralloadfactor at FS 515.6 (203)due to SMCS differentialvane deflection,case i. 66

36 SMCS lateralaxis performancewith relocatedforwardsensorpackage, frequencyresponseof lateralloadfactor at FS 746.8 (294)due to SMCS differentialvane deflection,case I. 67

37 SMCS lateralaxis performancewith relocatedforwardsensorpackage, frequencyresponseof lateral loadfactor at FS 515.6 (203)due to SMCS differentialvane deflection,case 2. 68

iv

LIST OF ILLUSTRATIONS - (Continued)

Figure Title Page

38 SMCS lateralaxis performancewith relocatedforwardsensor package,frequencyresponseof lateralloadfactorat FS 746.8 (294)due to SMCS differentialvane deflection,case 2. 69

39 SMCS verticalaxis performancewith relocatedforward

sensorpackage, frequencyresponseof normal loadfactor at FS 515.6 (203)due to SMCS vanedeflection,case 2. 70

40 SMCS lateralaxis performancewith relocatedforwardsensorpackage, frequencyresponseof lateralloadfactor at FS 515.6 (203)due to SMCS differential

vane deflection,case 3. 7241 SMCS performancewith relocatedsensor,PSD of vertical

load factor at FS 746.8 (294) - pilot station,case i. 7342 SMCS performancewith relocatedsensor,PSD of lateral

load factor at FS 746.8 (294) - Pilot station,case i. 74

43 SMCS performancewith relocatedsensor PSD of verticalload factorat FS 746.8 (294) - pilot station,case 2. 75

44 SMCS performancewith relocatedsensor,PSD of lateralload factor at FS 515.6 (203) sensor location,case I. 76

45 SMCS performancewith relocatedsensor,PSD of lateralload factor at FS 746.8 (294) pilot station,case 2. 77

46 SMCS performancewith relocatedsensor,PSD of lateralload factor at FS 515.6 (203) - sensor location,case 2. 79

47 SMCS performancewith relocatedsensor,PSD of lateralload factor at FS 746.8 (294) - pilot station,case 3. 80

48 SMCS performancewith relocatedsensor,PSD of lateralload factorat FS 515.6 (203) - sensor location,case 3. 81

49 SMCS performancewith relocatedsensor,PSD of lateralload factorat FS 746.8 (294) - pilot station,case 4. 82

50 Normal load factordue to horizontaltail frequency

response comparisonsof full and truncateddynamicanalyticalmodel, SCAS on. 91

51 Typicalvane-inducedinterferenceforces. 9352 Effect of SMCS vane aerodynamicinterferences,

frequencyresponse of normal load factor at FS 571.5(225) due to SMCS vane deflection,SCAS off, SMCS off. 95

V .

LIST OF ILLUSTRATIONS (Continued)

Figure Title Page

53 Effect of SMCS vane aerodynamicinterferences,frequencyresponseof normal load factor at FS 571.5 (225)dueto SMCS vane deflection,SCAS on, SMCS off. 96

54 Effect of SMCS vane aerodynamicinterferences,frequencyresponseof normal load factor at FS 571.5 (225)dueto SMCS vane deflection,SCAS on, SMCS on. 97

55 Effect of SMCS vane aerodynamicinterferences,frequency

responseof lateralload factorat FS 571.5 (225)dueto SMCS vane deflection,SCAS off, SMCS off. 98

_6 Effect of SMCS vane aerodynamicinterferences,frequency

responseof lateralload factorat FS 571.5 (225)dueto SMCS vane deflection,SCAS on, SMCS off. 99

57 Effect of SMCS vane aerodynamicinterferences,frequencyresponseof lateralload factorat FS 571.5 (225)dueto SMCS vane deflection,SCAS on, SMCS on. I00

58 Loads analysis SIC point locationsand DoubletLatticegeometry. 105

59 Controlsystemsfor gust loads. 10760 Structuralstationsfor gust loads analysis. 10861 Load calculationequations. 10962 SMCS off, wing bendingmoment frequencyresponse -

WS 985 cm (387.6in). 11563 SMCS off, wing bendingmoment responsepower spectrum

WS 985 cm (387.6in.). 11664 SMCS off, wing bendingmoment exceedances-

WS 985 cm (387.6in.). 11765 SMCS off, forebodybendingmoment frequencyresponse

FS 1377 cm (542in.). 11866 SMCS off, forebodybendingmomentresponse power

spectrum - FS 1377 cm (542 in.). 11967 SMCS off, forebodybendingmoment exceedances

FS 1377 cm (542 in.). 12068 SMCS off, forebodybendingmoment frequencyresponse

FS 2367 cm (932 in.). 12169 SMCS off, forebodybendingmoment responsepower

spectrum - FS 2367 cm (932 in.). 12270 SMCS off, forebodybendingmoment exceedances-

FS 2367 cm (932 in.). 12371 SMCS off, normal load factor frequencyresponse -

aircraftCG. 124

vi

LIST OF ILLUSTRATIONS(Continued)

Figure Title Page

72 SMCS off, normal load factor responsepower spectrum -aircraftCG. 125

73 SMCS off, normal load factor exceedances aircraftCG. 12674 SMCS off, normal load factor frequencyresponse -

pilot station. 12775 SMCS off, normal load factor responsepower spectrum-

pilot station. 12876 SMCS off, normal load factor exceedances pilot station. 12977 SMCS off, delta horizontalstabilizerfrequencyresponse. 13078 SMCS off, delta horizontalstabilizerresponsepower

spectrum. 13179 SMCS off, delta horizontalstabilizerexceedances. 13280 SMCS on, wing bendingmoment frequencyresponse -

WS 985 cm (387.6in.). 133

81 SMCS on, wing bendingmoment responsepower spectrum -WS 985 cm (387.6in.). 134

82 SMCS on, wing bendingmoment exceedances- FS 985 cm(387.6in.). 135

83 SMCS on, forebodybendingmoment frequencyresponse-FS 1377 cm (542 in.). 136

84 SMCS on, forebodybending moment responsepower spectrum -FS 1377 cm (542 in.). 137

85 SMCS on, forebodybendingmoment exceedances FS 1377 cm(542 in.). 138

86 SMCS on,forebody bendingmoment frequencyresponseFS 2367 cm (932in.). 139

87 SMCS on, forebodybendingmoment responsepower spectrum -FS 2367 cm (932in.). 140

88 SMCS on, forebodybendingmoment exceedances-FS 2367 cm (932in.). 141

89 SMCS on, normal load factor frequencyresponse -aircraftCG. 142

90 SMCS on, normal load factor responsepower spectrum-aircraftCG. 143

91 SMCS on, normal load factorexceedances- aircraftCG. 14492 SMCS on, normal load factor frequencyresponse -

pilot station. 14593 SMCS on, normal load factorresponsepower spectrum -

pilot station. 14694 SMCS on, normal load factor exceedancespilot station. 14795 SMCS on, delta horizontalstabilizerfrequencyresponse. 14896 SMCS on, delta horizontalstabilizerresponsepower

spectrum. 149

vii

I,IST OF ILLUSTRATIONS - Continued

Figure Title Page

97 SMCS on, delta horizontalstabilizerexceedances. 15098 SMCS on, delta mode controlvane frequencyresponse. 15199 SMCS on, delta mode controlvane responsepower spectrum. 152i00 SMCS on, delta mode controlvane exceedances. 153I01 B-I flight test aircraft,wings swept 65 degrees. 156I02 B-I air inductionsystem. 157103 B-I inlet subsonicdiffuser flow area distribution. 158

104 B-I inlet/engineaerodynamicinterfaceplaneinstrumentation. 160

105 B-I flight-testAIP data acquisitionand signalconditioningsystem. 161

106 In-flightcalibrationcycle, total pressuresataerodynamicinterfaceplane, flight 1-5, 64 sps. 163

I07 Representativediscriminatoroutput,AIP total-pressureCBW data, flight i-i0. 164

108 AIS/SMCS investigationflight 2-19,Mach 0.83. 168I09 AIS/SMCS investigationflight 2-19,Mach 0.83/16,000feet,

= 3 degrees,vortex ingestionin no. 2 inlet during

sideslipoperation,PLA = max. 169ll0 AIS/SMCS investigation,flight 2-19, _ = 3 degrees,

vortex ingestionin no. 1 inlet during sideslipoperation. 171

Iii AIS/SMCS investigation,flight 2-19, ff= 3 degrees,vortex ingestionin no. 2 inlet during IDLE-INTthrottle transient. 172

If2 Flight 2-33 test matrix and AIP signalsduringMach 0.85operationwith SMCS vanes deflected20 degrees,no. 2 inlet,RB = 7 degrees,RC = 5 degrees 173

113 No. 2 inlet,effect of sideslipangle on steady-stateinlet characteristicsat SMCV= 20 degrees, flight 2-33. 174

114 No. 2 inlet,effect of sideslipangle on steady-stateinlet characteristicsat various SMC vane angles,flight 2-33. 175 .

115 Time historiesof total-pressurerecoveryand stall-margin index,Mach 0.85, with SMCS vane deflectedand 0 and 20 degrees,no. 2 inlet, RB = 7 degrees,RC = 5 degrees,flight 2-33. 177

116 Dynamiccircumferentialand radial distortioncomponents,Mach 0.85, with SMCS vane deflected0 and 20 degrees,no. 2 inlet, RB = 7 degrees,flight 2-33. 178

viii

LIST OF ILLUSTRATIONS - (Continued)

Figure Title Page

117 Dynamic total pressurecontoursduringMach 0.85 179operationWith SMCS Vane deflected0 and 20degrees,no. 2 inlet, RB = 7 degrees,RC = 5de_rees, flight 2-33.........

118 Effectsof SMC vane positionon no. 2 inlet steady-stateand dynamic characteristics,Mach = 0.85 and alpha =

" 2.7 degrees,flight 2-37. 180119 Effects of sideslipangle on steady-stateinlet

characteristicswith SMCV= 0 degreesand alpha = 1

degree,flight 2-37. 181120 Effectsof sideslipangle on steady-stateinlet

characteristicswith SMCV = 20 degreesand alpha =

1 degree, flight= 2-37. 182121 Effectsof SMCVwake on no. 2 inlet dynamictotal-

pressurecontoursduring sideslip,SMCV = 20 degreesand alpha = 1.4 degrees,flight 2-37. 183

122 Effectsof sideslip angle on steady-stateinletcharacteristicswith SMCV = -8 degrees and alpha =1 degree,flight 2-37. 184

123 Effectsof SMCV wake on no. 2 inlet dynamictotal-

pressure contoursduring sideslip,SMCV = -8 degreesand alpha = 0.9 degrees,flight 2-37. 185

124 Effectsof sideslipangle on Steady-stateinletcharacteristicswith SMCV= 0 degreesand alpha =2.6 degrees,flight 2-37. 186

125 Effectsof sideslipangle on steady-stateinletcharacteristicswith S_4CV= 20 degreesand

alpha = 2.6 degrees,flight 2-37. 187126 Effectsof SMCVwake on no. 2 inlet dynamictotal-

pressure contoursduring sideslip,SMCV = 20 degreesand alpha = 2.6 degrees, flight 2-37. _ 188

127 Effects of sideslipangle on steady-stateinlet1 characteristicswith SMCV = 0 degreesand

alpha = 5.8 degrees,flight 2-37. 189_ 128 Effectsof sideslipangie on steady-stateinlet

characteristicswith SMCV = 20 degreesand

alpha = 5.8 degrees,flight 2-37. 190129 IndividualAIP probes, total-pressurerecoveryarray,

SMCS vane deflected20 degrees,Mach 0.85, _ = 6 degrees,flight 2-37. 191

130 Variations in circumferentialand radial distortion

componentsduring sideslipoperationwith the SMCS vanedeflected+20 degrees,Mach 0°85, _ = 6 degrees,

flight 2-37. 192

ix

I,IST OF ILLUSTRATIONS- (Continued)

F i gure Title Page

131 Effectsof SMCV wake on no. 2 inlet dynamictotal-pressure contoursduring sideslip,SMCV= 20 degreesand alpha = 5.8 degrees,flight 2-37. 193

132 Effectsof SMCVwake on no. 1 inlet dynamictotal-pressure contoursduring sideslip,SMCV = 20degreesand alpha = 5.8 degrees,flight 2-37. 194

133 Effectsof angle-of-attackmaneuverson steady-stateinlet characteristicswith SMC vane at 20 degrees,flight 2-42. 196

134 Effectsof an angle-of-attackmaneuver on steady-stateinlet characteristicswith no. 1 engine at intermediate,flight 2-42, SMCS vane deflectionangle = 0 degrees. 197

135 Effectsof an angle-of-attackmaneuveron steady-stateinlet characteristicswith no. 2 engine at intermediate,flight 2-42, SMCS vane deflectionangle = 0 degrees. 198

136 Effectsof an angle-of-attackmaneuver on steady-stateinlet characteristicswith no. 2 engine at intermediate,flight 2-42, SMCS vane deflection angle = 0 degrees, 199

137 Effectsof sideslipangle on steady-stateinletcharacteristicswith SMCV= 13 degreesandalpha = 3.3 degrees,flight 2-38. 200

138 Effectsof SMCVwake on no. 2 inlet dynamictotal-pressure contoursduring sideslip,SMCV = 13 degreesand alpha = 3.3 degrees, flight 2-38. 201

139 Steady-'stateinlet characteristicsduring no. 2 enginetransients,SMCV = 20 degrees, alpha = 5.5 degrees,and beta = +4 degrees, flight 2-38. 202

140 Steady-stateinlet characteristicsduring no. 2 enginetransients,SMCV= 20 degrees, alpha = 3 degrees, andbeta = +4 degrees,flight 2-38. 203

141 Steady-stateinlet characteristicsduring no. 2 enginetransients,SMCV= 20 degreesand -8 degrees,alpha = 1degree with positive sideslip,flight 2-38. 204

142 Typical dynamicresponsenear crew stationdue to

turbulenceduring low-altitude,high-speedflight. 207143 SMCS performancein turbulenceM = 0.70 alt = 305 m

(I,000ft) AGLA= 65°. 208

144 First fuselagesymmetricstructuralmode dampingfrom horizontaltail pulse excitations. 210

X

LIST OF ILLUSTRATIONS- (Concluded)

Figure Title Page

145 Effect of SMCS verticalgain settingon first fuselagesymmetricstructuralmode dampingat variousaircraftweights. 211

.- 146 First fuselageantisymmetricstructuralmode dampingfrom forced SMCS vane oscillations. 212

147 Vertical SMCS performancein turbulenceas shown bypower spectraldensity of vertical load factor atpilot station,FS 747 (294). 214

148 Lateral SMCS performancein turbulenceas shown bypower spectraldensityof lateralload factor atpilot station,FS 747 (294),single-peakresponse. 215

149 Lateral SMCS performancein turbulenceas shown bypower spectraldensityof lateralload factor atpilot station,FS 747 (294),double-peakresponse. 216

150 SMCS impact on short-periodand dutch roll frequencies. 217151 Ride qualityratingsfor varyingdegreesof turbulence. 221152 Effectsof turbulenceon flightpath controltasks. 221153 Effectsof turbulenceon tasks other than flightpath

control. 222

154 Effectsof turbulenceon readabilityof instrumentsand displays. 222

155 Effectsof turbulenceon reachlng/usingcontrols. 223156 Effectsof turbulenceon crew fatigue. 223157 Effectsof turbulenceon tendencyfor motion sickness. 224158 Effectsof turbulenceon physical discomfort. 224

xi

LIST OF TABLES

Table No. Title Page

I GeneralFlexibleVehicle Equationsof Motion, Time Domain 5II Equationsof Motion for LongitudinalRigid-Bodyand

SymmetricStructuralModes, FrequencyDomain I0I[l Equationsof Motion for Lateral-Directional,Rigid-Body

and AntisymmetricStructuralModes, FrequencyDomain 12IV TypicalAnalyticalStructuralMode Data 28V Longitudinal-SymmetricAero CoefficientsFrom Doublet

LatticeProgram 37VI SMCS Vane EffectivenessIncludingInterferenceEffects,

Longitudinal-SymmetricCase, FrequencyDomain 94VII Gust Loads Model StructuralDegreesof Freedom 103VIII Load Comparison,SMCS Off Versus SMCS On 114IX SMCS Vane Effectson Inlet/EngineCharacteristics,

FlightTest Investigation 166X SMCS High-GainTests Summary 219

xii

ANALYSESAND TESTS OF THE B-I AIRCRAFT STRUCrURALMODE CONTROLSYSTIg4

John H. Wykes, Thomas R. Byar, Cary J. MacMiller,and David C. Greek

Rockwell International,North AmericanAircraftDivision

- _y

An 18-monthprogramwas conductedto compileand document for publicationinformationpertainingto analysesand flight tests of the B-I StructuralModeControlSystem (SMCS). This is the secondphase of a continuingeffort;results from the first phase study are documentedin referencei. This reportcovers the followingtopics:

(I) Flexible aircraftequationsof motion

(2) Descriptionof flexibleaircraftanalysesmodel

(3) Comparisonof analysesand flight-testperformanceresults of theSMCS

(4) A stmmaryof the study of the forward_4CS sensorpackagerelocation

(5) Truncatedanalyticalmodels used in simulationeffort

(6) An analysisof the SMCSvane interferenceeffects

(7) ImFactof SMCS on selectedloads

(8) Flight-testresultsof the SMCS vane effectson inlet/enginecharacteristics

" - (9) Stmmaryof SMCS flight-testresults

INTROII/CTION

The B-I aircraftis one of the first vehiclesto includea controlcon-

figuredvehicle (CCV) conceptin the early design phases. The aircrafthas arequirementto providea specifiedlevel of ride quality for the crew. Thisrequirementhas been met on the B-I throughthe use of an automaticcontrolsystem (_CS) whose main externalfeatureis a set of vanes (near the crewstation)which are canted down 30 degreesfrom the horizontal. (See figurei.)

N

Structural modecont ro I vanes

Figure 1. - B-1 aircraft with wings swept aft.

A sM_stantialsavingsinweightwas achievedwith thisapproachas comparedto direct material stiffening. The details of systom re_luiranents had to bedetermined from a production (long-life) point of view, which has not beendone before for a system of this type. Extensive wind tunnel tests of thevane characteristics were conducted. Analytical models of the flexible air-craft and control systems were developed to analyze requirements and toinvestigate stability and performance. Component parts were tested to the

.- requirements in the laboratory. Flight tests of the SMCShave been conducted,and comparisons with analytical predictions have been made. Because of all ofthis, it has been recognized that the B-1 offers an excellent opportunity formuch needed further evaluation of such a system as the SMCSto insure theoptimun use of these systems for future applications.

The overall objective of this research area is to compile and documentinformation about the conceptual design, development, and flight tests of theB-1 SMCSand its impact on ride quality. Since the B-1 is the first aircraftto have a system such as the SMCSdesigned for production and long-service use,it is expected that the reports prepared will add to the technology base fordesign of future large military or civil aircraft. The specific overallobjectives are to:

(1) Investigate the improvements in total dynamic response of a flexibleaircraft and the potential benefits to ride qualities, handling qualities,crew efficiency, and reduced dynamic loads on the primary structures

(2) Evaluate the effectiveness and performance o£ the SMCS,which usessnell aerodynamic surfaces at the vehicle nose to provide damping to thestructural modes

The majoreffortof thephasei study(ref.I) was to compile,edit,andpreparefor publicationas a NASAcontractorreportthe existing-informationon the B-I SMCSconceptualdesignand development.The majoreffortof thepresentphaseII studyis to reporton the analysesand flighttestsof theSMCS;existinginformationhas been augmentedby someadditionallimitedanalysesand flight-testdatareductions.

FLEXIBLE AIRCRAFT EQUATIONS OF MorION

The equationsof motionof the flexibleB-1 formthe foundationof muchthat is to followand so it is appropriateto discussthemfirst. The treat-ment of theseequationsis suchthat theyarenot developedhereinfromfirstprinciples.Thereare manytextbooksthatdo thistask;references2 and 3are typicalof these. The equations,however,are presentedin sufficientdetailto be readilyrelatedto the textbooktreatments.The equationsof

3

motion oi,this sectionof the reportwere developedto serve the purposesof ride quality,terrain-following,and handlingqualitiesevaluations;theequationsof motion relatedto the loads analysesare discussedin a subse-quent section. An attemptwas made to includethe main aircraft charac-ter(stiesimportantto these severaltypes of studies;however, as specificstudieswere conducted,minor modificationswere often made. It is not intended

to catalogall of these iterations;where importantto the studiesreportedherein,they will be discussed.

The discussionstouch upon the key featuresof these equationsalong withexplanationsof the form of the data where it is felt that this would be help-ful. The authorshave electedto stay away frommatrix notation in these _initialdiscussionsin order to show as much informationabout the dynamicmodeling as possible. The appendix containsa list of symbolsused in definingthe equationsof motion of this section.

The equationsof the motion in tables I, II, and III are written in abody axes system (figures2 and 3) where the X-axispasses throughthe centerof gravityand is parallelto the vehicle fuselagereferenceaxis (FRL). Tohelp those who are more familiarwith stabilityaxes notation,it shouldbeobservedthat all of the aerodynamiccoefficientsbut two appear the same ineitherthe stabilityaxes systemor the body axes system. The two that aredifferentare the normal force coefficient,CN (bodyaxes), versus lift coef-ficient,CL (stabilityaxes), and the chord force coefficient,CC (bodyaxes),and drag coefficient,CD (stabilityaxes). For small angles of attack,CNnearly equalsCL and CC nearly equals CD.

All accelerationsand velocitiesare defined positiveas indicatedinfigure 2. One exceptionis noted relativeto the definitionof the verticalload factor. As a concessionto stabilityand controland flight test conven-tion, a positiveload factor,nz, is associatedwith a positiveCN. In con-trast, the definitionof Z-axisacceleration(consistentwith the positivedefinitionsof figure 2) is associatedwith a positiveCZ in figure 3

(Cz = -CN).

Ride quality analyses have been conducted on the simulator using time-domain equations while frequency-domain analyses have been conducted usingdigitalcomputingequipment. The equationsin the time domain are presentedin table I and the frequencydomain in tables II and III.

The structuralflexibilityof the air vehicle is defined in terms offree vibrationmodes of the structure(oftenreferred to as normal modes).Figure 4 definesthe sign conventionsused in associationwith.these normal

TABLE I. - GENERAL FLEXIBLE VEHICLE EQUATIONS OF MOTION, TIME DO~~IN

[Total Vehicle, Body Axes, Units: ft, lb, rad, sec]

FORCE EQUATIONS, RIGID-BODY MODES

- ~~~&H - (rr:ov:cyXco5r)6cv- (:~h.)5A - (t~j[(CN~TF"tJ

i3 - -', + Poc -~~),8 + ~?:,cosesln<p +(~ s..,\{C~ (J + CY.6Ibw)~ + Cyr('b..,')r- + CYf(bw'f~v. V) \m"l."J (0() ~~ ~ z.\i.J'

m n

+ CYp(~)p + Cyp(~fp + [[CY71lni+ CYni(~il + [(CyPn.n~(3 + Cyo..6r- + C'ys..Sr--.-. .=. I •ANrlt.YMI1ET~IC SYMMETRIC

+ Cy~~ +CY6~S: + CY&:v0~v+CY6~~CV +e;,OAoh.+c,-~6A.} +(~,jj&' +(~~(s\nf')~v+ (m:'I..~)8A

+ (~~)~(yv,)v"J

TABLE I. - Continued

STRUCTURAL_DE EQUATIONS

Symmetric modes, i --1 to n

_._..___C_ _nJ=.l

\ MI \ MI

Antisy_etric modes, i = 1 to m

m n

+ _,_vJ__ + +J=l ,J=l

_T "Am 0 "


b,_]B_TEQUATIONS, RIGID-BODY MODES _ Cm_(_--" + \ ly /L (o0 .

n

\- Iy

- +\_,(i,.

_w_.r_,Syl.lrl i_T RI C S'Y_11"t£TRt¢"

.i_ _. \ Ix / lx /


LOAD-FACTOR EQUATIONS (ACCELERO_TER READINGS) _ .xX

ROMi_" 7_I AT ANy | r_ is-_-IN THE SAME"rATION )4..Y.i2! DII::_ECTION AS -I- Cr. I

TATION X,Y, ;z I'l), IS + IN THE SAME

DII_ECTION AS -i- C7ROTARY-RATE EQUATIONS (RATE GYRO READINGS)

n

[_,T ANyTATION _(,y] i=l

_--_L i ._,g_ .

P = P + Le, r_;('._._-,_._R,_.o0_s)rATANy X'Y_]_JLSTAT ION

i,,I

I-- = I_ + _-=-_Y'_ "rAT ANY Z__' 7_i (ANTISYT'II_ETRIC NODES)LSTATI ON _,Y,_] i=l

TABLEI. - Concluded

EULERANGLES

COS_

EARTH-AXESVELOCITIES

_-- N_CoseCOS_+ Vy(sIn_slnocos_- cos_sln_) +V_(co%_slnocos_r. sln_ s,n_)

V..E--_coses,n_. ',_(cos}cos_+s,nc}_nes,mf)+V_(cos}s,nesmg-sln{cos_r)

x4=-_s_ne+vrs,n_cose+v,cos}coseA convenient equivalent expression for Vze is in terms of altitude rate of change and _ and 8.

I_ = V. (s,necos0ccos/s -cosesln{s,rtp - cosecos_cos,_sln0c)

_o

TABLE II. - EQUATIONS OF MOTION FOR LONGITUDINAL RIGID-BODY AND¢_ Slqv_4ETRICSTRUCTURAL MODES, FREQUENCY DOMAIN

[TotalVehicle,Body Axes, Units: ft, Ib, rad, sec]

NORMAL-FORCE EQUATION, RIGID-BODY _K)DE

PITCHING-MOMENT EQUATION, RIGID-BODY MODES

TABLE II. Concluded

SYHMETRICSTRUCTURALMODE EQUATION,i = 1 to n

,<,.+i<...>t,_ I%7 t,i<,U"n H x#

,/NORMAL LOAD FACTOR (INCREMENTFROM 1 g TRIM) t_,_

FArA.,_x.¥,lz-J i-, iL_TAT IONNote: The variablesare complexin form but

have been written as _, q, _i, 6k, andwgas a space-savingnotation.

[( ) ()R + i ( ) _()g]representsreal andimaginaryparts of unsteady aerodynamicscoefficientform.

These equationsrepresentperturbationsfrom a 1 g trimmedflight condition.

• F_a

N TABLE III. - EQUATIONS OF MOTION FOR LATERAL-DIRECTIONAL, RIGID-BODY, AND

ANTISY_4_ETRICSTRUCRJRAL .MODES, FREQUF_NCYDOMAIN

[TotalVehicle,Body Axes, Units: ft, Ib, rad, sec]

SIDE-FORCE EQUATION, RIGID-BODYMODE

!

TABLE III. - Continued

ROLLING-MOMENT EQUATION, RIGID- BODY ~IODE

ANTISYMMETRIC STRUCTIJRAL MODE EQUATION, i = 1 to m

TABLE III. - Concluded

LATERAL LOAD FACTOR

STATATANy

_oN x,y, :_] i:l

Note: The variablesare complex in form but

have been writtenas _, r, p, qi, 8k,and Vg as a space-savingnotation.

+

[( ) ()R i ( ) ()i]representsreal

and imaginaryparts of unsteady aero-dynamicsin coefficientfom. Theseequationsrepresentperturbationsfrom1 g trimmedflightconditions.

+Y

+q

+vz,a

+X+r

+P

'+Z

_ Bqdy axes

Figure2. - Sign conventionfor rotationaland linearrates and accelerations.

+Cn +_I_

+N .6r+Y

+C N

+N .._ +C_- +L

' = Differental angle6H

+T between left and\ right panels

+Cz+Z

+X +6cv, Symmetric deflection +Z

+_cv, Differential deflection

Deflection magnitude and sense

referenced to right panel; right +Cm

panel trailing edge down, left +M

panel trailing edge up.

Body axes

Figure 3. - Sign conventionfor coefficientsand controlsurfacedeflections.

Symmetric Modes

Side view

+_fx

+ X Body axis_.d d,

ing point [ +Crli

AntisymmetricModes

+_/i_lizing point +C7/.,

<+x ' --_'x cG _+4 'xTop view

Looking forward

Figure 4. Structuralmode deflections,slopes,and generalizedforcessign convention.

17

modes. There are a number of advantagesto using the modal approach incontrastto the direct-influencecoefficientapproachfor such studiesasreferred to herein.

(i) It can describe the static as well as dynamic characteristicsofthe flexiblevehicle in a consistentmannerwithin the same format.

(2) Both inertiaand aerodynamicloadingsare accountedfor in themodal generalizedforce data for both the static and dynamiccases.

(3) The modal data are amenableto a number of simplificationschemes.

(4) The approachintegratesbest with control systemdesign requirementsin the handlingand ride qualitiesdesign areas.

Considerfirst, the most extensivelydetailed equationsof motion in thetime domain as describedin table I. These equationsneed be used in this

detailedform shown only when large-scalemaneuveringis studiedas in theterrain-followingproblem. More simplifiedequationswill serve other analysespurposes.

These equationswere developedassumingthat the angles of attack,_, andsideslipangles, 8, (figure5), would be small (lessthan i0 degrees),but thevehicleorientationin space as definedby the Euler angles9,@ , and • (fig-ure 6) would be unrestricted(exceptfor @ = 90 degrees).

Aerodynamicdata indicatedare preliminaryestimatesof those required.These data are shown in derivativeform exdeptwhere it is anticipatednon-linearcharacteristicswith _, _, or controldeflectionoccur. As an example,the normal force curve was expectedto be nonlinearwith _ so the normalforcecoefficientis expressedas CN(_) insteadof the more familiarlinear form

CN_-

The controlsurfacesexplicitlyshown are those anticipatedbeing requiredby either the Stabilityand ControlAugmentationSystem (SCAS)or SMCS. Other

control-surfaceinputsare shown in generalform as functionsof _k for thekth surface.

Becauseof the requirementto control structuralmotion at the frequenciesof the lower free-vibrationmodes, it is necessaryto considerunsteadyaero-

dynamic effectsof the controlmotion as well as the inertiareactionforcesof these surfaces. In the normal force-equationformatof _.ableI, the

unsteadyaerodynamicsare shown by the notion CN_k_k + CN. & for the kth6k k

18

E_

Body axes

Figure 5. - Angle-of-attack(_) and sideslip (8) definitions.

19

Vertical plane

Airplane plane

of symmetry

l

+ X-BodyI'II II Horizontal plane

m

I+ Y.Bod \ I

l \ I

\\\ + X-Earth

+ Z-Body+ YrEarth

I

3

+ Z-Earth

Rotation sequence indicated by numbers at arrow heads

Figure 6. Euler angle definitionsand rotationsequence.

2O

control surface. The control-surfaceinertiareactionforce in this equation

has the form (mk _k) 6"k- Similaraerodynamicand inertiareaction-forceterms may be recognizedin the moment and structuralmode equations.

Terms in table I involving4, 4, r, and p are generallyof small conse-quencebut have been includedto be consistentwith the frequency-domainequationsof tables II and III where unsteadyaerodynamicsmust be considered.The desirabilityof this consistencywill become apparentas these equations

" are discussedin subsequentparagraphs.

The B-I enginegyroscopicmoments, IR_Rr and IRO_Rq, shown in the equationshave not proved to be significantin handlingqualitiesor terrain-followingstudies. It had been anticipatedthat they might have been significantinlarge-scalemaneuvering. They have, however,been left in the equationsoftable I.

The equationsof table I includethe abilityto change speed. It isassumedthat if significantchangesin Mach numbersare to be realized,thesechangeswill be reflectedin the appropriateuse of Mach carpet data for theaerodynamicsrequired. Velocity changeswill show up directly in Vo, while

combinedaltitude-velocitychangeswill appear indirectlyin qo, (1/2 pV2).

The gust representationshown in table I is in the aerodynamictransfer-

functionform. Tilegust excitation,Wg and Vg, would come from random signalgenerationsources,shaped and scaledto reflectthe desiredgust powerspectraldensity and intensity.

The longitudinaland lateral-directionalrigid-bodymotionsare coupledduring large-scalemaneuversthroughthe inertiaterms and enginegyroscopiceffects. The symmetricand antisymnetricstructuralmode motionsare coupledby terms representingthe dihedraleffectdue to symmetricwing bending. Forsmall motions about a trim condition,the rigid-bodymode equationscan bedecoupledthrougheliminationof the inertiaand gyroscopiccoupling. Thestructuralmode equationscan be decoupledby using trimmedairplane static

-: symmetricstructuralresponseparametersat fixed values to determinetheeffectivedihedraldue to symmetricwing bending.

The load factor and rotationrates of the large-scalemaneuveringflexibleair vehicle (as read by accelerometersand gyros mounted on the fuselagestructure)are presentedin table I.

Euler angle equationsand earth axis velocitiesare given in table Iand can be used in terrain-followingstudiesto determinethe vehicle attitudeand locationwith respectto the earth'ssurface.

21

The frequency-domainequationsof motion for the flexibleairplanearegiven in table II for the longitudinal-sy_netriccase and table III for thelateral-directional-antisymmetriccase. These equationsare uncoupledandrepresentmotion perturbationsfromwing-level1 g trimmed flight. Compari-son of these frequency-domainequationswith the uncoupledtime-domainequationswill help in identifyingequivalentterms.

As shown, the vehicle responseaerodyn.amics(that is derivativesassociatedwith responseparameterssuch as a, _, q, q) are quasisteady,while the controlsurfacesand the gust are sho_cnas functionsof the forcingfrequencyin theform [( )R + i ( )I]- This formathas been convenientand sufficientlyaccurate for.preliminaryride qualitiesand structural-modecontrolanalyses.Digital programsare availableat Rochcell;however,that will also acceptvehicle responseaerodynamicsfrom unsteadyaerodynamictheoriesas a functionof frequency,permittingmore refinedride qualityand S_ICSstabilityanalyses.

The equationdescribingthe normal load-factorresponseat any locationin the flexibleaircraft is presentedin table II. The similarequationforthe lateralload factor is given in table III. These responsesare used inride qualityanalyses.

FLEXIBLE AIRCRAFT ANALYSES bDDEL

This sectiondescribeshow the data were obtainedto implementthe flexibleaircraftequationsof motion used for ride quality analyses. It will not bethe purposeof this presentationto providea complete set of data used inall analysesdiscussed;but it will be the intent to provideunderstandingofthe data used.

DYNAMICANALYSISSYST_

As a basis of understandingthe contentsof this sectionbetter,as wellas topics of other sections,the chart of figure 7 is presented. Shown is --the completedynamicanalysis system supportingthe developmentof flexiblevehicle dynamicanalysismodels for controlsystem developmentand ride qualityanalysesat RocMcell International'sNorth AmericanAircraftDivision,E1 Segundo. The path throughthis system as employedin developingthe SMCS isas follows. (No supersonicanalyseswere conductedduring SMCS developmentsothe Mach-box programcapabilitywas not used.)

Starting at the top of the chart, it is seen that the processbeginswith a definitionof the vehicle geometry,basic wind tunnel correlatedaerodynamics,structuralstiffness,and mass characteristicsbeing providedto

22

I I I il

Vehicle geometry I- Basic aerodynamics

Structural definition

I-- J T IMaSs characteristics,_

'' _ ' ,

STAR 6 stiffness'

vibration analysis El, GJ'SIC

Modes, frequencies

J Modal Iprocessing

_ Normalized modes,

Doubletlattice II lJ M>IMaChbox l_lizedm<] l I

Dynamics,J _ ' I_ , Generalized

Technolog_ J Generalized j aero forces

I I I

Control _ FH251/FH255 H STAR 4

system Dynamic Flutter

analysis response analysis

Control system requirements Flutterdynamic response, stability characteristics

• _ I al

-, Figure 7. - Dynamicanalysis system.

23

the DynamicsTechnology(,roup.As indicated,the stiffnesscan he used inthe EI-iU formator in the formof structuralinfluencecoefficients(SIC).The ,lassand stiffness&ata enter the STAR 6 program,and vibrationanalysesof the whole vehicleare accomplished. The output of the program,then, iswhole vehiclevibrationmode shapesand frequencies.

These mode-shapedata are next manipulatedto producenormalizedmodes.For tileB-I ride qualityanalyses,both the sy_netricand antisymmetricmodeswere normalizedto a I_int on the nose of the vehicle. Generalizedmass dataconsistentwith the normalizedmodes are produced. Finally_modal deflectiondata are developedthrough interpolationprogramsalong selectedstreamwisestrips for input into the aerodynamicprograms.

For tile ride quality related studies reported herein, the Doublet Lattice.Program has been used to provide theoretical aerodynamic generalized forcesas required. These generaI ized aerodynamic forces are in dimensional form.Programs have been developed which process the data from dimensional form tothe coefficient form required by the dynamic analyses programs. These dynamicanalyses programs employ the frequency-domain equations of motion discussedearl ier.

The 1:tt-251 program provides dynamic response results for the longitudinal-symmetric case, while the FH-255 program provides dynamic response results forthe lateral-directional-antisymmetric case. Both of these programs can accepteither frcxtuency-dependent or quasisteady data. Active controls can beincluded. Frequency responses due to gust or control forces may be obtained.When gust inputs are employed, ride quality parameters are output and control-system deflections and rate responses are obtained in power spectral densityform. Stability analyses are performed using the characteristic determinantfrequency evaluation technique of reference 4.

FREE-FREE VIBRATION MODAL DATA

The flexible aspects of the aircraft have been treated in the modal _format as opposed to the direct-influence-coefficient approach. Ninety per-cent of the ride quality analyses performed have been accomplished usingfree-free vibration modes which were obtained using an EI-GJ description ofthe vehicle stiffness; more recent modal data have been obtained usingstructural influence coefficients. The details of the EI-GJ approach arediscussed here.

Figure 8 shows the typicaldistributionof mass pointson the elasticaxes assumed. This is an earlymodel; figure 9 shows a refinementof thefuselage elasticaxis made at a later date. Each analysisincludeda flexible

24

. _FS 2461 (969) FS 4018 (1582)

Nace IIe

, points

Z

Typical control point motions calculated from EI-GJ approach

WL 72.4

IUk -81 (-32) FS 3942 (1552)

Figure 8. - Typicalelasticaxes and mass point locations.

25

I-o

Elastic axis., 'D f EA _ FRP

I -- L._ _1 (w,o)

Elastic axis coordinates*

Point Fus Sta WL

A O 38.6 (15.2)

B 406 (160) 34.0 (13.4)

C I003 (395) 143.8 (56.6)

D 1283 (505) 123.2 (48.5)

E 2477 (975) 80.0 (31.5)

F 3937 (1550) 154.9 (61.0)

G 4013 (1580) 50.8 (20.0)

H 4430 (1744) 57.7 (22.7)

*Straight lines between these points

Figure 9. - Fuselageelasticaxis refinement.

wing, fuseIage, horizontal tail, and vertical tail. Also included wereflexibly mounted engines/nacelles. The resuItant mode shapes consisted ofelastic axis deflections and rotations illustrated in figure 8.

Because the analyses were primarily oriented toward ride quality at thecrew station, the free-free vibration modes were normalized at the most forwardmass point at the nose of the aircraft. While most modes show a high degreeof coupling among vehicle components, table IV identifies the main component

; (wherethis is possible)and lists the associatedfreque.ncies.In the analysesdiscussedherein, I0 symmetricmodes and 12 antisy_netricmodes were used.Discussionsto follow later in this sectiondescribethe rationaleused in

selectingthe modes shown.

When groundvibrationtest (GVT)data became available,they were usodto upgrade the modal data. Symmetricorthogonalmodes were successfullyex-tractedfrom such tests. First, the effectsof the soft supportsystemwereremovedfrom the measuredorthogonalset of modes. These data were next used,with proper fuel weights included,to analyticallyobtainedorthogonalfree-free modes at the desiredweight condition. Itwas not possible,however,to directlyextractan orthogonalset of antisymmetricmodes from GVF data.To obtain usable, consistent, antisymnetric modal data reflecting test results,a technique was used of adjusting local stiffness data until a successfulapproximation oF the measured data were obtained analytically. Then, as inthe s_nmetric case, the effects of the soft suspension were deleted and thedesired fuel weights adeed analytically to obtain orthogonal antisymmetricmodes. The data of table IV are typical of those based on the GVT data ob-tained as described. It is to be noted also that structural damping wasextracted; these data were obtained using the oscillation decay method.Figuresi0 and II present typicalsymmetricand antisymmetricmode vector plots.

The computerprogramsused to determinethe aerodynamicdata requirevibrationmode shapes to be definedon a grid systemwhose chords are parallelto the free stream. Thus the basic mode shapeswere interpolatedto find

point deflectionsalong strean_isechords for all liftingsurfaces. (See fig-: ure 12.) The mode-deflection data on the grid system shown were used as input

to the Doublet Lattice Program where interpolations to the Doublet Lattice gridsystem were made and required slope data determined within the program.

AERODYNAqICDATA

The DoubletLatticeaerodynamictheorywas used to obtainmost of the

aerodynamicsdue to the flexiblestructure. In addition,extensivewind

27

TABI,Ii IV. - TYPICAl,ANALYTICALSTRUCTURAl,MODEI)ATA

Wt = 119,296.8 kg _(263 000 lb)A = 65°

Structural ablode Mode Frequencyno. description Hz damping

gs

Symmetric7

1 Wing, first bending 2.22 0.0622 Fuselage,first bending 2.84 .0943 Fi'rstnacelle 3.26 .024

4 IIorizontal tail, f_rst bending 4.19 .0285 Wing, fore and aft bending 4.23 .0526 Fuselage, second bending 6.28 .0167 Wing, second bending 7.57 .0228 llorizontal tail, fore and aft bending 8.31 .0649 Fuselage, third bending 11.15 .055

10 llorizontal tail, first torsion 27.35 .042

Antisymmetric

1 First nacelle I.73 .145

2 Wing, first bending 2.41 .0543 Horiz tail, first bending 3.51 .0434 llorizontaltail, first bending 3.96 .0255 tlorizontal tail, fore and aft bending 4.14 .0496 Wing, fore and aft bending 4.20 .0317 Fuselage, first bending 5.58 .0328 Wing, second bending 6.96 .0319 Vertical tail, first bending 7.21 .019

10 Fuselage, second torsion 9.72 .07811 Fuselage, second bending 10.30 .02212 Vertical tail, first torsion , 35.34 .020

aDeterminedfrom groundvibrationtests

28

Fore and aft motion

Lateral motion

Vertical motion

Mode l

_ Frequency = 2.23Hz

Figure I0. Typicalsymmetricmode vector plot.

29

Fore and aft motion

Lateral motion

Vertical motion •

\

\

\

_/ Mode l

Frequency = 1.73 Hz

t :-

Figure II. - Typicalantisymmetricmode vector plot.

3O

..

9

17

;>..1

' ..'..

t. 3Sb

03 ••

Figure 12. - Typical structural mode deflection grid points.

Horizontal tail location,

WL 320 (126)

WL 63.5 (25)

I FS 0

Horizontal tail pivot,

FS 4018 (1582)

Figure 13. - Panellingand box grid for Doublet Latticeaerodynamicsprogram.

t_mel data were also generatedto obtain static,rigid-bodyforce and moment-coefficientdata along with pressure-distributiondata. These two data sourceswere correlated (to be described)and used as input to the flexible aircraftanalyses.

The panelling,box grid, and control-surfacessetup for the DoubletLatticeProgramare shown in figure 13. The wing and fuselageforebodyhadfivemain panels with a total of 151 boxes. The horizontaltail had twopanels and a total of 64 boxes; the horizontaltail was _ut in, also, as anall-movablecontrolsurface. The vertical tail had two panels and 73 boxes.In addition,the vertical tailhad an end plate at its base consistingofone panel with 30 boxes. As shown in figure 13, the lower rudder controlsurfacewas also modeled.

For longitudinal-symmetricdata, includingcontrol effectivenessand gustdata, the wing/forebodyand horizontaltail were run as shown in figure13.This same wing/forebodyarrangementwas used along with the full empennage(horizontaltail, verticaltail, and end plate) to obtain lateral-directional-antisymmetric,whole-vehicledata. Vertical tail gust data, lower ruddercontroleffectiveness,and differentialhorizontaltail effectivenessdatawere obtainedby runningthe empennageplus end plate as an entity. Forfuselageside gust and generalizedaerodynamicforces,a modified slenderbody theorywas used which made use of wind tunneldevelopedside-forcedistributiondata.

It is to be noted that the SMCS vanes were not modeled for the Doublet

LatticeProgram. The aerodynamicsfor the vane were obtained fromwind tunneltests as describedin referencei. None of the vane-relatedaerodynamicswerefrequencydependent. The reducedfrequency(k) determinedon the basis of anassumedfrequencyof i0 Hz (62.8 rad/sec),the highestfrequencyrange thatvane is expectedto be effectiveat M = 0.85, is

k = _--_= (02.8)(2.46) = 0.0812: ZVo (2](9Sl)

Compared to k for thewing under the same circumstances

k- (62.8)(15.23) _- 0.503(2)(9Sl)

This is a relativelylow reducedfrequency,and the aerodynamicunsteadinesseffectsare judged to be acceptablysmall. In retrospect,this appearstohave been a valid judgement. In generatingthe structural-mddegeneralizedforce coefficients,the vane force was assuned to act at a point; i.e.,

33

FS 581.66 (229). Thus C_. = ¢iFS 581.66for symmetricmodes is anexample. ' 1_cv CN_cv

In developingthe DoubletLatticepanelingand box patterns,lifting-surfacespanloaddistributionswere comparedto the wind tunnel-relateddata.The gap betweenthe vehiclecenterlineand the first row of chordwiseboxesof the horizontaltail is one of the devicesused to obtain matcheddata sets

for the horizontaltail. It was reasonedthat with good matches of thelifting-surfacespanloadings,the computermodel developedwould give validanswers for both rigid body and structuralmodes.

The point of vi_¢ was adoptedthat the wind tunnel-relatedairloaddis-tributionfor the rigid vehiclewere the most accuratedata available. Thus,

the rigid vehicleaerodynamicscouplingsinto the structuralmodes (Cni_ and

Cni6 data are examples)were computedusing these distributionsratherthanDoublet Latticetheory,for the zero-frequencycase. These data, in turn,

were used to scale the frequency-dependentdata produced by the Doublet

Lattice Program. The Cnin.jand Cni_j data were used directlyas generatedbythe programwithout scalingfor both the synraetricand antisy_netricmodes.

As an exampleof how the wind tunnel data and the DoubletLatticefrequency-dependentgust aerodynamicdata were brought intoagreement,considerthetypical e_ampleof figure14. Shown is the pitching-momentcoefficientdatadue to a unit vertical gust velocity. The magnitudetrendsversus the reducedfrequency,k, were assumedbasicallycorrectand all adjustmentsto match windtunnel based data were made at low frequenciesas illustrated. The data magni-t_le of the real componentat zero frequencywas adjustedto match the windtunnel based data and then faired into the basic real curve at low reduced

frequencies. For the case illustrated,the gust coefficientat zero frequency

was determinedfrom angle-of-attackdata, CmWg (Cm_)/Vo. This matchingofthe angle-of-attackand gust velocitydata is essentialto obtainingvalidpower spectraldensityresponsesdue to gust data at low frequencies;a mis-match will producea load factorresponseat zero frequencywhich is not therein the real world. In the realworld, a stable aircraftwill weathervaneintothe resultantvelocitydue to the combinedmotion and gust velocity and haveno load factorat zero frequency.

As mentionedwhile discussingthe analysesflow, the data producedbythe DoubletLattice Programmust be reducedto the coefficientform of theequationsof motion discussedin the previous section. In order to do this,the data are processedin the followingmanner.

34

0 Real l Symbols denote

[] Image[ calculated data pointO.OO4

0.002

Crr_g

1 ,

ft/sec 0 __ .6 k .8 1.0 "_ I._I 6-0.002 ,--_r]F -- ..

-- ,ow-freqoencyadjustmenttofit wind-tunnel-test based data

-0.004

Gust reference point is FS 2649 (I043)

Figure 14. - Typical low-frequencyadjustmentmade to analyticalfrequency-dependentaerod_lamicdata.

The programoutputsdata normalizedto air density (2p)and to frequency

(o_2) as illustratedhere using the dimensionalliftforce due to plungingLh(forthe whole vehicle), [Lh/2P_]. The divisionby two is because the data

generatedin the program are for half of a vehicle, p is assumedto be unityby the program. Velocity is determinedby the data of Machnumber and velocityof sound at the altitudeselected. Frequency,o_,is input at a number ofselectedvalues.

Having this information,table V illustrateshow the frequency-dependentcoefficientdata are developedas a functionof frequency(mr reducedfrequency,k). Longitudinal-synmetricalcoefficientsare obtainedusing the informationof table V; lateral-directional-antisymmetriccoefficientdata are obtained ina similarmanner.

Reflectedin table V is the sign conventionbuilt into the Doublet LatticeProgramat Rockwelland the sign conventionassumed for the equationsofmotion of tables I, II, and III. The or_l°ydifferenceof importanceto theunderstandingof the derivationsof table V is the fact that lift force,L,of the DoubletLatticeProgram is of opposite sense to the equationof motionnormal force, N, and its associatedcoefficient,CN. Otherwise,the pitching

moment, M, and structuralmode generalizedforces,Qi' are identicalindefinition.

As an aid in understandinghow table V was assembled,the following

exampleis given for the derivationof CN.

For no pitching (@ =@= 0)

Vo Votherefore:

qoSw CN _= - [Lh] h force unitsIMAG

IMAG

v°2T CN= -o!

IMAG

INAG

36

TABLE V. LONGITUDINAL-SYMVIETRICAERO COEFFICIENTS FROMDOUBLET LATTICE PROGRAM

37

The bracketedquantity [ ] here and in table V comes directly from the DoubletLattice Programas a functionof frequency.

One of the questions to be resolved in using a modal approach to aircraftflexibility modeling is the one of how many modes to use. The technique usedto help make this decision for the described analyses is discussed here.First, the following criteria were developed as guides in the decision process.The modes contributing to main a_roelastic impacts on all short-period andDutch-roll characteristics were to be included. The modes contributing tomain aeroelastic impacts on control effectiveness were to be included. Themodes contributing most to flexible fuselage motion at the pilot station andactive control sensors were to be included.

Figure15 is typicalof the data generatedto assist this mode selectionprocess. For the longitudinal-syn_netriccase used as an illustration,some

25 whole-aircraftnormalizedmodes were input to the Doublet LatticeProgramand run at a frequencyof n/2 (a frequencyjudged to be in the regionof short

period and Dutch roll frequency). These data then were reducedto aerodynamiccoefficientform and enteredinto a programwhich calculatesquasisteadyflexible-to-rigid(F/R) ratios (referto 'TruncatedAnalyticalModels") for

all of the key aerodynamicderivatives(CNaandCn_ are used as illustrationshere). These F/R ratios are calculatedas one mode after the other is elimi-

nated. As a result,it is possibleto identifythe individualmodes contribut-ing most to a given derivative'saeroelasticimpact.

After the preceding procedure is accomplished for all important deriva-tives, plots similar to figure 15 are assembled and inspected as a whole. Thus,those modes making important contributions to all derivatives are selected forr et ent ion.

As far as fuselagemotion was concerned,as many modes as possible,havingfuselagemotion as a main component,were selected. Usuallymodes reflectingup to the third fuselage-bendingand the second fuselage-torsionmode couldbe selected.

• 2-

CONTROL-SURFACES INERTIA REACTION FORCES

The inertiareactionforces of the controlsurfacesare importantinputsto the stabilityanalysesof active controlsystems. The detailedfinal form

of these inputs for the B-I controlsurfacesare indicatedin the equationsofmotion of the first part of this report. The basic approachthat was used indevelopingthese expressionswill be discussedhere, but each control-surfaceinput will not be developedin detail.

38

For constant Mach number and altitude

l.O

0.8

CNol_ _-- ....F/R %,

%..-_

0.6 \\llll

0.4 |l -. /iI /"L_....----/ CM_x

0.2

0 I * I | | I I a s i I I I I I I I I I I I I I a

0 4 8 12 16 20 24

Number of modes

"L

Figure 15. Typicalaeroelasticflexible-to-rigidratiodata for aerodynamic

coefficientsas a functionof participatingstructuralmodes.

39

Figure 16a shows that when a typicalcontrol surfaceis acceleratedinthe positivesense of the deflection,a mass reaction force and moment aredevelopedat the centerof the mass of the controlsurface. This force andmoment are reactedinto the basic aircraftstructureat the control surface

hingelineas a force and moment as shown. The force acting on the aircraft isthe force shown,andthe moment acting about the aircraftcenter of gravityis (_m_+ I )6" where _ is the distancebetweenthe control-surfacehingeline_Eand the aircraftcenter of gravity.

The generalizedforce acting on a typicalstructuralmo'dedue to control"surfaceaccelerationis illustratedin figure16b where the specific exampleof a s_etric mode is used. Before getting i1_tothe specificexample,con-sider the conceptof a generalizedforce. A generalizedforce has units ofwork, that is m - N (foot-pounds). In this case, it is eitherforce multipliedby mode deflectionat the force applicationpoint,moment multipliedby modeslope at the moment applicationpoint, or both. A positivegeneralizedforcewould act to increasethe deflection (or virtualwork) of the structuralmode.Lookingat the example in the figure,it can be seen, for the example shown,that the reactionforce causesa generalizedmode force incrementof -_i_Lm_"

and the reactionmoment a generalizedmode force incrementof +¢iILI_L6".

The precedingdevelopmenthas proved an adequaterepresentationfor smallcontrol surfacessuch as the B-I SMCS controlvane and lower rudder control

surface. It is not accurate enough,however,for large control surfacessuchas the all-movablehorizontaltail. In this latter case, it was necessarytobreak up the mass characteristicsof the surfaceonto a distributedgrid system.Using the distributed:m_ssdata and the previouslydevelopedlogic, generalizedforceswere developedfor rigid body and structuralmodes. These data, devel-oped using distributedmasses,were input to the digitalprogramwhich imple-ments the equationsof motion of table II using an equivalentpoint mas.srepresentationof the data.

During the B-I development,checksof the pitch SCAS, with the aircrafton the ground resting on its landinggear, revealed a structuralmode-couplinginstabilitywhen excitedwith sharp horizontaltail Controlinputs. This • _-instabilitycould be analyticallyduplicatedby using generalizedcontrol-surfaceinertiareactionforces developedemployingthe distributedmassapproach but could not be duplicatedusing the single-pointmass representation.

The inertiareaction forcesare particularlyimportantto the stabilityof SMCS with the aircrafton the ground. The inertiareactionforces have anopposite s_se to theaerodynamic forces of the SMCS vanes. The stabilityofthe system is establishedby these aerodynamicforces. Thus, if the aerodynamicforces disappear,the feedbacksense is effectivelyr_versed,producingan

t

4O

Reaction force and moment actingon primary structure at H_

Reaction force and moment

about surface CG

° _ Primary

structu

I ' +(_

'IAircraft H_

, [ _l_._.._l CG of control

l l surface mass,m

Force at aircraft CG = +m{_"

Moment about aircraft CG = -(_m{ + IFL )i_"(pitch axis assumed

for illustration)

(a) Rigid-body generalized forces

Symmetric mode i

\ Reaction View from

X 4///_/_'(t s'CG Reaction"force _iFL left sidemoment \_

7L

Z

Note definitions of +(_liand +_iI'L.

q, 'm //5. Structural generalized forces =-( { _i H_

(b)Structural generalized forces

Figure 16..- Typical controlsurfaceinertiareactiongeneralizedforces.

41

instability. On the B-I, a switchon the landinggear preventsoperatingtheSMCS while on the ground, precludingany inadvertantdamage due to thispotential instability.

ACTIVE GONTROLSYSTt_/IS ..

Two types of active control systemswere includedin the analysesper-tainingto this study. One type, SCAS, is associatedwith controlof whole-

vehicle (shortperiodand Dutch roll) modes of motion. The second type,SMCS, has the functionto controlfuselagestructuralmotion to improveridequality.

The block diagramsand analyticalmodelingof the SCAS are given infigures17 through19 and S_ES in figures20 and 21. Flightcondition-dependentgains are shown for M = 0.85 at SL. These figures indicatethe type ofsensors,compensations,gains, and actuatormodeling assumedfor each of theindicatedsystems. The control-surfacedeflectionequationsare cast in aform directlyusable by the Rockwellresponseanalysesprograms. That is tosay, the overallgain is indicated,system dynamicsare representedby numeratorand denominatorroots of polynomials,and vehicle motionsare definedasmeasuredby the appropriatesensors.

COMPARISONSOF ANALYSESAND FLIGHT-TESTRESULTS

The ride quality indices,_z and Hy (referto reference5 for detaildefinitions),for the vertical and lateralaxes, respectively,cannotbemeasureddirectly in flight. One difficultyis the fact that the parametersare obtained from weighted power spectraldensitycurves of crew-stationaccelerations,and the weightingcan presentlyonly be done as a postflightoperation. Secondly,it is very difficultto preciselyfly a specificationvehicleweight at the specificationMach number and altitude. For the B-l,the approach has been to demonstratethat the analyticalmodel can duplicateflight-testresultsand then proceedto use the verifiedmodel in the required - _ride qualityanalyses. It is the intent of this task to presentthe datamatches which providedthe verificationand give an evaluationof the factorsaffectingthe matches.

The flight-testdata obtained for matchingpurposeswere frequencyresponsesof load factorsat FS 571.5(225)due to SMCS vane inputs. Bothvertical and lateralload-factordatawere obtained. The flight conditionflown was M = 0.85 at 762 meters (2500 feet). The vehicleweight was approxi-mately 119 296 kilograms (263000 pounds). Wing sweep was 65 degrees. Datawere taken with all controlsystems inoperative;then, with only the SCAS

42

•8 p.Ao

R,Ki"E aG,_o _<p (s_.o4S+9co_ .a(, 5'z.14. 1o

k_%, I'2.1.,,,L"_ 4,9 _01 AT _+ 57,14- S_|OF.S. 2;<.49 5.k.

:q Hi NOTCH MPEMSATIO SEP-.VO ACTUA_W|LTFJ_

RAD

I NOR.HAL L__ GAIN / //3--__--'2-('Y_;%'7"°_6_r'i'13"28__.S.2'¢49

mAD 1199<3.;_3

"L_ ;=1 ..J

For flight condition of M = 0.85 at SL

Figure 17. - Pitch axis SCAS analyticalmodel.

~~ Fl XEO F'OR ALL

g J=LIGHT CONDITIONS~ .....~ .44- 6.0

YA\N~TE COMPEN~TIOH K~...G'fRO 57.14 20

~.s. :l.(.!l19 S'+ 57.14- 9'+20( (043)

I<h NOTCH SERVO30 P ~ILTER

OR.Sr30 1.0

ATRAn

LATERAl COMPfNSATION l<MI"l S.L..Acr:.EL. "

For flight condition of M = 0.85 at 5L

Figure 18. - Yaw axis SeAS analytical model.

_S+ 1.05±J20.97) R_oRAo/_SC 57.14 I0(,5+ %.4 + J 19.35) .5? S -i-57.14- _ + IO

F_ATE - G_|N SERVO TUATOI_

GYI_,o r F ILT E I_

E s._649 FIXED gO_. RAD(IO4.3) ALL FLIGHT

CONOITIONS

325.7 m

' (S+l.os+_2o.O_I ! r57 5714 I0 - •-... ,. <a": -( )( s+_.4-JLg.3_S_7._X_-i_/_,Co)_+(,.olP :__AD J= I

For flight condition of M = 0.85 at SL

Figure 19. Roll axis SCAS analytical model.

4_OX

_0o) los" _o-7.31 c_+to .5"¢.5' .159 _+.lo _.SO

- ,0d@FH I_ F i+AT COFt P. _ ILTF_.R ASHOUT Tt._TOW_HF_

F.S. 57|, 5 CONSTANT C.OC.KPIT,, S.L. HA_ sE'r'r_,_- DF,J3

IO (C_'H57_'1"i57) H=.8S" o1=,O=1 VOLT-cj+jo

,_+!! /NGI_III_I_ LAG I-"

c-e-I I 1I:'.s._

(_o,_3) _o.2.7

IS.7±iIS(..Z) S' I

I--'I

IOZ.7, I"I

, .(£ __I.('.Is3X_XI31XI_Io)OoXIoXso)_I V, \(_+l£7+-JISt.._)/______/I IXI_. _,_<.__,_,/'_IO43 _. "L

RAD i=l

For flight condition of H = 0.85 at SL

Figure 20. - VerticalSMCS analyticalmodel.

_8 los 5o-7.31 S'+ 1.0 .5"(.,,5" .153 _ + |0 $+50

AT _,_ V COMP. _,L'rE_. uA'_oII I I'_"_1I !1 I_

F..s.S'/!,s" ¢OHs'rAHT _.iS.7:!:_ 1_(..2) AT ¢<:x:KPrrIo (_41sTX_,+_s'f)M=.ss o_,o=IVo-T

_.+IO

I_IOHINALI-_ IL/_4_ _-- -- Note: Apparent extra -_ due to

c.G.I I I simulatio,ofbankanglein lateral load factorF._.Z_ terms by

Oo4s) s8.0_

_:_v/_- (S_.sX3z.z) .... S,I_(.S+IST_+ISZ)_+IO_+SoAS/ ° "_' & --'_ V.dV, i

eo._

• _I

RADFor flight condition of M = 0.85 at SL

Figure 21. - LateralSMCS analyticalmodel.

operating;and finally,with both SCAS and SMCS (cockpitgains 1.5) operating.The forcingamplitudeand frequencysettingswere set manually in the cockpit.The measureddata were processedto obtain both magnitudeand phase charac-teristics.

The analyticalstructural-modecharacteristicsfor the data-match

analyseswere obtainedfrom test and analysis sources. The symmetricstructuralmodes were obtained directlyfrom GVT. It was not possibleto obtain a set of

orthogonalmodes for the antisymmetriccase directly from the GVT. Instead,the analyticalstructuralmodel was adjusted in stiffnessuntil the frequencyand mode-shapecharacteristicswere as close to the observed characteristicsas

possible. In supportof these tests, the basic vehicleweight characteristics(no fuel) wePe identified. For the specificdata-matchinganalyses,fuelloadingswere determinedfrom measurementsmade duringthe flight-testperiodwhen the frequencyresponseswere being executed.

The pitch SCAS characteristicsused in the analyseswere as describedinfigure 17; the yaw and roll SCAS were as describedin figures18 and 19,respectively. The verticalSMCS descriptionis given in figure 20; and thelateral SMCS descriptionis shown in figure 21.

Comparisonof the analyticaland flight-testresultsof the normal loadfactor of FS 571.5 (225) _'requencyresponsedue to synmetricSMCS vane deflec-tions (SMCSused as an excitationsystem)for the conditionof no activecon-

trols (basicaircraft)are sho_cnin figure 22. The three responsepeaks are(startingwith the lowest frequency)first fuselagebending, second fuselagebending, and third fuselagebending, respectively;there is some influenceonthe second peak from the wing secondbending. The qualityof the match isconsideredexcellent. To obtain this match, however, three adjustmentsweremade to the originalmodal characteristics: (I) the wing first bending-modefrequencywas reducedfrom 7.57 to 7.00 Hz, (2) the fuselagethird-bendingfrequencywas reducedfrom 11.15 Hz to 8.60 Hz (thislatter change is substan-tial, and no reasonhas been found to explainwhy the originalmode was off),and (3) the structuraldamping (gs)for the fuselagethird bendingwas changedfrom 0.055 to 0.025. Havingmade these adjustmentsfor the basic aircraft .responses,no additionaladjustmentswere made to the analyseswith controlsystemsoperating.

Figure 23 displays the data matches for the frequencyresponsewith theSCAS operating. Comparingthe first peak-responsemagnitudeof this figurewith the previous figure,it can be seen that the SCAS excitesthis peak some.Again, the analytical-to-testdata match is excellent.

Figure 24 shows the data matches for the frequencyresponsewith boththe SCAS and SMCS operating. The effectivenessof the SMCS in damping thefirst fuselage-bendingmode is demonstratedby these data. The data match is

48

Note: 8cv is control surface deflection. The flight-test data measurementsof the forcing command were analytically processed to remove effects

of actuator dynamics,which were measured,in order to permit com-

parisons with analytical results on this and similar subsequentfigures.

150

" !oo /_/ \ I

// \Phase

ang]e 150 I \ , _\deg /j/ _/'

Sl I I I I I I _// _

O_ / 2 4 6 8 _'_0% !%I Frequency (f) - Hz

-50:-"

Flight-test data

Analytical data

8

grad

L

00 2 4 6 8 I0

Frequency (f) - Hz

M = 0.85, air = 762 m (2500 ft)Wt = 119 296kg (263 000 lb), .A= 65 °

Figure 22. - Comparisonof flighttest and analyticaldata, frequencyresponseof normal load factor at FS 571.5 (225)due

to SMCS vane deflection,SCAS off, SMCS off.

49

olOO !Phase I

angle ,50 /_/ ,__l_ 2//deg I0 / " a

_,/ 2 4 6 8 1

Frequency(f) Hz-50

Flight-test data

Analytical data

12 -

g

4

,d

0 I l I I I n n

0 2 4 6 8 lO _Frequency (f) - Hz

M = 0.85, alt = 762 m (2500 ft)

Wt = 119 296kg (263 000 lb), A= 65°

Figure 23. - Comparisonof flight test and analyticaldata, frequencyresponseof normal load factorat FS 571.5 (225)due

to SMCS vane deflection,SCAS on, SMCS off.

5O

15o

Flight-testdata )SMCS vert gain 1.5

Analytical data 912

n z

8

g

4 / \\

7 "*

0 '_'I" I I I I I , , a i

0 2 4 6 8 I0

Frequency (f) - Hz

M = 0.85, alt = 762 m (2500ft)Wt = ll9 296kg (263 000 Ib), .A= 65°

Figure 24. - Comparisonof flight test and analyticaldata, frequencyresponseof normal load factor at FS 571.5 (225) due

to SMCS vane deflection,SCAS on, SMCS on.

• 51

excellentfor the first two responsepeaks and fair for the third peak. Thelack of better fit for this third peak was initiallyattributedto SMCS act-uator modeling in thishigher frequencyrange; however,using the flight-test-derived actuatormodel of figure 25 did not substantiallyimprovethe matchto the flight-testdata in the 8 to I0 Hz frequencyrange as shown in figure 26.

The comparisonof the frequencyresponselateralload factorat FS 571.5(225)due to differentialSMCS vane deflectionsfor the basic vehicle is shownin figure 27. The comparisonis fair; the frequenciesof the three peakresponsesare duplicatedby the analyses,but the amplitudeof the low-frequencypeak is off by a considerableamount. The phase angle is matchedreasonablywell in the midfrequencyrange only; specificreasons for why the data do notmatch better are not known. The peaks are identified(startingwith the low-frequencypeak) as the wing fore and aft mode with a large fuselageside-bendingcomponent,fuselagefirst side bending,and second fuselagetorsion. This lastpeak in the analysiswas obtainedby droppingthe frequencyfor this modefrom 9.72 to 7 Hz; although this mode was the only logicalone to adjust,there is no reason known for the noted discrepancy. Having made this logicaladjustmentfor the basic vehicle description,no other adjustmentswere madewhen the controlsystemswere operated.

The data comparisonof figure 28 are for the case of SCAS operating.The data are similarto the basic aircraftresponseof figure 27. Again,the agre_nentof analysisto test data is only fair.

In figure 29, the frequencyresponsecomparisonsare made for the caseof SCAS and SMCS operating. The trend of the analysesand the flight-testdata are similar;i.e., the first and secondpeaks are attenuatedbut thepeak around 7 Hz is increasedby the SMCS operation.

The implicationsof these analytical/flight-testdata comparisonsareimportantto the B-I ride qualityverification. The data imply that thevehicleanalyticalstiffnessand mass characteristics,whole-vehiclecontrol-surfaceaerodynamics,SMCS vane aerodynamicsand inertiareactionforces,andSCAS and SMCS modeling are fairlyaccurate. Thus, the ride qualitycharac-teristicscan be calculatedwith considerableaccuracyat specificationorany other set of flight conditions. The Iongitudinal-sy_netricaircraftcharacteristicshave been more accuratelydescribedthen the lateral-directional-antisymmetricset.

The aircraftride quality characteristicshave been calculatedusing thedescribeddata set and have been presented in referenceI, pages 56 and 57.

52

Flight control simulatorFlight-test-derived frequency response .........

Analytical model actuator, K + 50

3.0- K = 1.93 deg/volt

-154.4 (s-205.5)Fitted transfer function (S + 82.8 ± j95.6)

0 I I I I I I I I ! I

0 2 4 6 8 lO

Frequency (f) - Hz

0 2 4 6 8 lO

0 _ J w I I I i i J ,'

Phase -20, _'__'w"A A

: angle ____A

deg -40

-60

Figure 25. - Comparisonof analyticalSMCS actuatormodels to flightand simulatortest data.

53

150

Phase 1O0 langle k

50

deg :

0_" 2 4 6 8 lO

-SO Frequency (f) - Hz I "'

12

-- Flight test data

, ( 50------Analytical data _-_--_) actuator

I n z (S-205.5)_cv 8 ......Analytical data, -154.4 (S +82.8 .+-J95.6)actuator,

Test derived

grad

I _*,

4 X't_..

!

00 2 4 6 8 I0

Frequency (f) - Hz

M = 0.85, alt = 762 m (2500 ft)Wt = 119 296kg (263 000 lbs),A= 65 °

Figure 26. - Effect of test-derivedSMCS actuatormodel, frequencyresponseof normal load factor at FS 571.5 (225)due to SMCS

vane deflection,SCAS on, SMCS on.

54

400

F\300 /

I l

Phase 200 I x

"" angle I \/_._z,_.I i \

100 u I \. deg / _ \" 0 I . I * I I

2 4 6_ 8 JO

-100 - Frequency (f) - Hz

-200 -_" ....

Flight test data

Analytical data10

m

In]g I

rad ,_1 _ I4 I \ ;I _JIII

2 I l, \/I _

0 _* a_ , a , * * a ,0 2. 4 6 8 0

Frequency (f) - Hz

M = 0.85, alt = 762 m (2500 ft)

Wt = I19 296kg (263 000 lbs),A=65 °

Figure 27. - Comparisonof flight test and analyticaldata, frequencyresponseof lateralload factor at FS 571.5 (225)due to SMCS vane

deflection,SCAS off, SMCS off.

55

Figure 28. - Comparisonof flight test and analyticaldata, frequencyresponseof lateralload factor at FS 571.5 (225)due to SMCS vane

deflection,SCAS on, SMCS off.56

200 '

loo _ r\-. Phase 0 /"'_-_- _ _.,/ , _'._

angle 2 4 6_ L ' /-lO

de_ -100 _re_uen_If) Hz "7" _ J-200 "" ""

-300

10

Flight test data)

_SMCS 1at gain 1.5Analytical data

8

_Y I 6 [_cv I I

Ig

rad 4 !tI

t I |I i \2 I iI I \

00 2 4 6 8 0

Frequency (f) - Hz

M = 0.85, alt = 762 m (2500 ft)

Wt = 119 296kg (263 000 ]b),A= 65°

Figure 29. Comparisonof flight test and analyticaldata, frequencyresponseof lateralload factor at FS 571.5 (225) due to SMCS vane

deflection,SCAS on, SMCS on.

57

FORWARD SMCS SENSOR PACKAGE RELOCATION

On the whole, the SMCS has worked well in improvingthe ride qualityofthe low-altitude,high-speedflightregime operatingat cockpitsettinggainsof i.5 in the verticaland 1.5 in the lateral. The lateralaxis performance,however,has been below that of the vertical.

As part of the ongoing investigationto determinehow to improvethelateralSMCS performance,the data shown in figure 30 were obtained. Thedata shownare power spectraldensity (PSD)plots of the pilot stationlateralaccelerationresponsedue to turbulencemeasuredduring flight 1-20 whileflying at M = 0.80 at about 305 meters (i000 feet) altitude. The data showthat the SMCS, with the lateralgains set at i.5, significantlyreducesthekey peak response at 4.5 Hz, slightlymodifiesa secondpeak at 6 Hz, andincreasesthe responsesignificantlyat 7 Hz. The net effect is an improvementin ride qualitybut not a large one. When the pilot increasedthe lateralgainto 2.2, therewas a dramatic increaseof the approximately7 Hz responsetoa level which was feltby the crew to be not acceptableeven thoughthe 4.5and 6 Hz responseswere furtherreduced.

The comparisonof the flight-testand analyticallateralaccelerationfrequencyresponsesdue to vane excitationshown in figure 29 d_nonstratedthat the analyticalmodel could reproducethe essenceof 7 Hz increasedresponse.A study of the analysisresults indicatedthat the two peak responsesat 4.5and 6 Hz are due to fuselagesidebending componentsin these modes while theresponseat 7 Hz is the fuselagesecond torsionalmode. Figure 31 will assistin showinghow the responsephenomenonobservedoccurs. Key in the analysisis the locationof the SMCS sensor packageat fuselagestationFS 571.5 (225),WL 142.24 (56)and BP 60.96 (24). Both the vertical and lateralaccelerometersare located in this package. As the vanes are differentiallydeflected,a sideforce and a torqueare created. The lateralaccelerometersees lateralaccelerationdue to both the side force and torque. When the lateralaccelera-tion signalsare sent throughthe SMCS, the side bending-ir_lucedsignalsareproperly phasedbut the torsion-inducedsignalsare adverselyphased,resultingin a reducedgain margin of the 7 Hz mode. It is also importantto note thatthe verticalaccelerometeralso sees the torque-inducedmotion,and undesirableSMCS symmetricalvane motionsare causedby lateralSMCS operation. Data,however,have shown this not to be a large influence.

Analyses shown in figure 32 indicatethat observedadversetorsioncoupling,as well as the coupling into the verticalaxis, could be eliminatedor attenuatedby relocatingthe SMCS sensorpackageclose to the fuselagecenterlineand near the elasticaxis. Since there would be no lateralmoment

arm, the couplingto the vertical axis would actuallyresult in the torsionalsignal phasingbeing favorable.

58

SMCS off

SMCS on, vert gain 1.5,1at gain 1.5

• o..... SMCS on, vert gain 1.5,1at gain 2.2

Flight 1-20

M = 0.80, air = 305 m (lO00 ft)

A= 65°

Figure 30. - Effect of SMCS lateralgain on power spectraldensityof lateralload factor at pilot stationFS 746.8(294).

59

SMCS nz

sensor package torsion I

An iIonat F$ 571.5(2251 ' Ytorsold location

' 60.96 c£ -_(24. in.),

i142.24 cm__Estimated

(56 in.) //'elastic

axis

_-_ 16.05 cm(6.32 in.)

I 9..62 in.)

ISMCS

ISMCS vane

:sensor package Resultant 1

at FS 515,6(203) Force vectors" " due to vanenew location

deflection

Looking toward rearfrom nose

Figure 31. - SMCS sensorpackagelocationsand couplingcharacteristics.

SMCSoff

8 _k _ SMCS on -" ny 1 old sensor

locationI lat gain 3.0

g 6rad SMCS on -

new sensorlocation

4 FI lat gain 3.0'I

"q

O

o

2

%eO eS

•If •

0O 2 4 6 8 lO

Frequency (f) - Hz

I. • I I I .... I I ,I

0 I0 20 30 40 50 G0

Frequency (w) - rad/sec

M-- 0.85, alt = 762 m (2500 ft)

Wt = ll9 387kg (263 000 Ib), A = 65°

Figure 32. - Effect of SMCS sensorpackage locationon lateralload factorat FS 571.5 (225)due to differentialvane deflection,analyticaldata.

61

The advantagesof moving the SMCS packagemay be sLm_narizedas follows-

(i) Higher lateralgains could be used to improvelateralride quality.

(2) Higher effectiveverticalgains (dueto a farther forwardsensorlocation)would improvevertical ride quality.

(3) The couplingof lateralvane inputsto the verticalaxis would beeliminated.

As substantiatedby both pilot co_ent and flight-testdata, the originalSMCS considerablyimprovedthe ride quality. However, the gains in the systemwere limitedto values below those originallyintendedby the couplingdescribedherein and not by systemmaximum capability. Based on the describedadvances,action was taken to move the forwardSMCS sensor packageto the new locationat FS 515.6 (203),WL 19.36 (7.62)and BP 16.05 (6.32). This relocationwasfirst accomplishedon A/C-I and then on A/C'2; this modificationwas neveraccomplishedon A/C-3. A/C-4 has the sensorpackage at the new location.

SMCS stabilitytests were initiallyperformedat high altitudeto evaluatethe forwardSMCS sensorpackage relocationeffects. The flight condition/configurationwas M = 0.85, altitude 6096meters (20 000 feet) and weight119 297 kilograms(263 000 pounds).

The First fuselageverticalbending-modedamping obtainedfrom the pitch-pulse transientdata is shown in figure 33 and is seen to be a linear functionof the S_4CSgain. All other modes were stable at the indicatedgain conditions.The first fuselagebending-modedampingwith the SMCS forwardaccelerometerrelocated (flight1-41) is comparedwith the resultswith the accelerometerin its previouslocation (flight1-7). The mode dampingappears to be signifi-

cantly larger at the higher gains with the new sensor location.

The lateralbendingmodes were not stimulatedsignificantlyby the rudderpulses, so that similardampingcharacteristicscould not be obtained. However,the lateralSMCS was stable for all values of gain tested (maxim_ cockpitknob settingof 6). Resultsobtainedwith the previous sensor location(flight1-7) showedthe SMCS to be unstableat a settingof 6.

• Followingthe high-altitudetest, stabilitytests were conductedat lowaltitude. The initialflight condition/configurationwas M = 0.85, altitude914.4 meters (3,000feet), A = 65 degrees,and weight 119 297 kilograms(263 000 pounds). The excitationswere horizontaltail and rudder pulses.First, each axis was tested to a maximum gain for that axis (V-gain= 3.0 andL-gain= 3.0, respectively). Followingthis, tests were conductedwith the

62

Legend: "--O-----O-- _

Flight 1-41 Flight I-7Mach = 0.85 Mach= 0.85

A = 65 degrees A = 65 degreesAir - 6096 m Alt = 6096 m

(20000ft) (200ooft)Wt = I19 297 kg Wt = 131 544 kg

(263oooIb) (290ooo Ib)Relocated SMCS Original SMCSsensor sensor

0.3

JI

J.E JO_

E f0.2--

o jI=

"O

_o.1- /7 ._ "

12 (

L !: 0 ' !, ,o 2 4 6

Vertical SMCS gain (cockpit knob set)

Figure 33. - First fuselageverticalbendingmode dampingversus SMCSgains for original and relocatedforwardsensorpackage.

63

SMCS operatingin each axis simultaneously. The maximum combinedsettingwas V-gain and L-gain = 3.0. The qualityof the response data was such thatthe d_qlpingfactorcould not be numericallydetermined;however,these qualita-tive data indicatedthat the SMCS was stable.

Followingthe horizontaltail and rudder pulse-stabilitychecks, SMCSresponseto SMCS vane excitationevaluationswere made. Figure 34 containsthe vertical-axis-responsedata. Shown are-thenormal load factoratFS 515.6 (203) (sensorlocation)frequencyresponsedue to SMCS sy_netricvane deflectionsfor SMCS off and SMCS on at vertical gains of 1.5 and 1.8.

These data show good performancefor the verticalSMCS and are not signifi-cantly differentfrom the similardata of figure 24 for the SMCS forwardsensor packageat the original location.

Figure 35 shows the lateralload factor at FS 515.6 (203) (sensorlocation)frequencyresponsesdue to differentialSMCS vane deflectionsfor SMCS off andon at lateralgains of i.5 and 2.2. As shown, the large 5 Hz (approximate)first fuselagelateralbending mode was significantlyattenuatedat the highergain without the previouslynoted large adversetorsionalcouplingresponseat approximately7 Hz.

Figure 36 shows the lateralload factor at FS 746.8 (294) (pilotstation)frequencyresponsesdue to differentialSMCS vane deflectionsfor the SMCSoff and on at lateralgains of 1.5 and 2.2. As on the previousplot, the

large 5 Hz first fuselagelateralbendingmode was significantlyattenuatedat higher gains. The 7 Hz response,however, shows a slight increaseinmagnitudeover the zero gain response.

Additionaldata similarto that just described,but at a higherweight

condition,are shown in figures 37 and 38. These data indicateless attenua-tion of the 5 Hz mode and more excitationof the next higher frequencymode

peak.

Rememberingthat the lateralnominal gain setting is i.5, all of thesedata indicateda substantialnet improvementin lateralload-factorresponsedue to the SMCS.

To furthercheck out the relocatedforward sensorpackage,tests were

conductedat the off-designconditionat M = 0.55, altitude762 meters(2500 feet),A : 55 degrees. Figure 39 presentsthe normal load factoratFS 515.6 (203)frequencyresponsedue to sy_aetricSMCS vane deflectionforSMCS off and on at vertical gains 2.5 and 3.0. These data show good perform-ance for the vertical SMCS at this off-designcondition.

64

SCAS only

SCAS + SMCS, vert gain 1.5...............SCAS + SMCS, vert gain 1.8

.I I I I I , I I

: 0 I0 20 30 40 50 , 60

Frequency (w) - rad/sec

Fit 2-20, runs 9.1,9.2,9.3M = 0.85, alt = 914m (3000 ft)

Wt = 129 502kg (284 500 lbs), A= 65°

Figure 34. - SMCS vertical axis performancewith relocatedforwardsensorpackage, frequencyresponseof normal load factor at

FS 515.6 (203)due to S_CS vane deflection,case i.

65

SCAS only

SCAS + SMCS, lat gain 1.5............SCAS + SMCS, lat gain 2.2

16

12

8g

r-_

/ \4 / ._ .

/ m

00 2 4 6 8 10

Frequency (f) - Hz

I, I I I I I I

0 I0 20 30 40 50 60

Frequency (w) - rad/secf

M = 0.85, alt = 914m (3000 ft)

Wt = 122 018kg (269 000 Ibs), A= 65°

Figure 35. - SMCS lateralaxis performancewith relocatedforwardsensorpackage, frequencyresponseof lateralload factor

at FS 515.6 (203)due to SMCS differentialvane deflection,case I.

66

r

SCAS 0nlySCAS + SMCS, lat gain 1.5

............SCAS + SMCS, lat gain 2.2

-- I0

8

cv I

0 , . .--- '. • _"Ii ''" t. I I ! IO 2 4 6 8 I0

Frequency (f) - Hz

I I I I I ' I

0 I0 20 30 40 50 60

"_ Frequency iW) - rad/sec

M = 0.85, alt = 914m (3000 ft)

Wt = 122 018kg (269 000 Ib), A= 65°

l:igure36. S_S lateral axis performance with relocated forward sensorpackage, frequency response of lateral load factor at FS 746.8 (294)

due to SMCS differential vane deflection, case i.

67

SCAS o_ly' SCAS + SMCS, lat gain .75

SCAS + SMCS, lat gain 1.5+ + + + + +SCAS + SMCS, lat gain 3.0

16

12

CV I[

8g

rad |'I

.

4

• ..I.4

0 s i i I I I I I I

0 2 4 6 8 I0

Frequency (f) - Hz

I I I I I I I0 I0 20 30 40 50 60

Frequency (o_)- rad/sec

M = 0.85, alt = 487.m (1600 ft) '

Wt = 139 709,kg (308.000 Ib), A = 65°

Figure 37. - SMCS lateralaxis performancewith relocatedforwardsensorpackage, frequencyresponseof lateralload factor at FS 515.6 (203)

due to SMCS differentialvane deflection,case 2.

68

SCAS only

SCAS + SMCS, fat,gain .75

SCAS + SMCS, lat gain 1.5

+ + + + + SCAS + SMCS, fat gain 3.0

12

lO

8

ny .

.6cv . .

6 . .. .

g 44

rad .

4 _"4

4-2 /,

@

0 ' I I I I I I I I

0 2 4 6 8 lO

Frequency (f) - Hz

I I I I I , I i

"- 0 l0 20 30 40 50! 60

Frequency (0_)-rad/sec

M = 0.85, alt =487m (1600 ft)Wt = 139 709 kg (308000 Ib), A = 65°

Figure 38. SMCS lateral axis performance with relocated forward sensor

package, frequency response of lateral load factor at FS 746.8 _294)due to SMCS differential vane deflection, case 2.

69

SCAS onlySCAS + SMCS, vert gain 2.5

....... SCAS + SMCS, vert gain 3.0

12

8

4

g

rad

00 2 4 6 8 0

Frequency (f) - Hz

I I I I I I I0 I0 20 30 40 50 60

Frequency(_) - rad/sec

M : 0.55, alt : 762 m (2500 ft)

Wt = 132 451 kg (292 000 Ibs),A: 55°

Figure 39. - SMCS verticalaxis performancewith relocatedforwardsensorpackage, frequencyresponseof normal load factorat FS 515.6 (203)

due to SMCS vane deflection,case 2.

70

Figure40 shows the lateral load factor at F$ 515.6 (203) frequencyresponsedue to differentialdeflectionof the SMCS vanes for the SMCS offand on at lateralgains of 2.5 and 3.7. As shown, the large 5 Hz firstlateralside-bendingmode was significantlyattenuatedat the higher gainswithoutthe previouslynoted adversetorsionalresponse at approximately7 Hz.

"" In additionto the precedingfrequencyresponse data, SMCS performancedata in turbulencewere obtained to evaluatethe relocatedforwardSMCS

sensor package. Figure 41 is a power spectraldensity plot of the pilot station(FS 746.8 (294))vertical load-factorresponsedue to ttmbulencewith the SMCSoff and on at vertical gains of 1.5 and 1.9. As shown, the SMCS was veryeffectivein attenuatingthe first fuselageverticalbending response.

The power spectraldensityof the load-factorresponsewas normalized

to unit root mean square (RMS) gust intensityCOw_) as derivedfrom the angleof attack (_ vane) data. This is not an accurate_techniquebut is the best

availablein absenceof gust boom data. A similarnormalization(OVg) has beenattemptedfor the lateralaxis data using sideslip (B vane) data.

Figure 42 shows the lateralload-factorresponseat the pilot stationwith the lateralSMCS gain zero but the vertical gains at 0, l.S, and 1.9.Since there is r_ mechanismfor the vertical axis SMCS motion to couple intothe lateralaxis, these data indicatethe level of repeatabilityof thelateraldata.

Figure 43 containsvertical load-factorresponse data at the pilot stationwith the SMCS off and on at vertical gains of 1.5 and 1.9, togetherwithlateralgains of i.5 and 2.2, respectively. Comparisonsof these data withthe data of figure 41 show that littleor no coupling of the lateralaxisactivity is evident in the vertical axis response.

Figure 44 shows the S_S performancein the lateralaxis at FS 515.6(203)with the SMCS off and lateralgains of i.5 and 2.2. These data showthat the 5 Hz (approximately30 radiansper second)first fuselagelateral

: bendingmode is significantlyattenuatedat both gain settings. However, atgain settingI.5, the secondfuselagelateralbending-modepeak response near6 Hz r_aainsabout the same in magnitudebut shifts slightlyupward in fre-quency. At lateralgain of 2.2, a significantincrease in magnitudedevelopswith this frequencyshift.

Similar data to figure 44 for the pilot stationare shown in figure 45.These data show the S Hz mode reductionbut indicatean increasedcouplingwith gain increaseof the highermodes.

Attention is directedto the power spectraldensity scales for figures144and 45. The data of figure 45 are a factorof I00 smallerthan figure 44.

71

SCAS only

SCAS + SMCS, fat gain 2.5

........ SCAS + SMCS, lat gain 3.7

2

g

rad

00 2 4 6 8 10

Frequency (f) - Hz

I, I ! a I m./___ e I

0 l0 20 30 40 50 60Frequency (OJ)- rad/sec

M = 0.55, alt = 762 m (2500 ft)

Wt = 129 276 kg (285 000 Ibs),A= 55 °

Figure 40. - SMCS lateralaxis performancewith relocatedforwardsensor

package,frequencyresponseof lateralload factor at FS 515.6 (203)due to SMCS differentialvane deflection,case 3.

72

Run ll.l vert gain = 0.0, lat gain = O.0------ Run II.2 vert gain = 1.5, lat gain = 0.0

........ Run ll.3 vert gain = 1.9, lat gain = 0.0SCAS on

200

xlO_ 6 "200

xlO-5

160

160 _nz

120

12o g2g2 (ft/sec) 2

(m/sec)2 "rad/secrad/sec 80

8O

|40 40

0 00 10 20 30 40 50 60

Frequency (_) - rad/sec

I I "1 I I • I :_ | , I I I I

0 2 4 6 8 I0

Frequency (f) - Hz

M = 0.85, alto152 m(500 ft)AGL

Wts 128 369kg(283 000 Ib),A = 65°

Figure 41. - SMCS performancewith relocatedsensor,PSD of verticalload factor at FS 746.8(294)- pilot station,case I.

73

Run II.I vert gain = 0.0, lat gain = 0.0

.... Run 11.2 vert gain = 1.5, lat gain = 0.0

..........Run ll.3 vert gain = 1.9, lat gain = 0.0i SCAS on

0.28xlO-6 .-

o.28x]O-5

0.24

0.24

0.20

O.20 _ny

o.160.16 2

2 gg (f t-TTsec) 2

(m/--_ec)2 "rad/sec

rad/sec O. 12

0.12 !o. 08 o. 08 I .

0.04 0.04 \_ _1_/ _\

V 'IQQ

e -foe

0 0 I I I I l .-0 ]0 20 30 40 50 60

Frequency (u_)- rad/secI I _L I I I I I • I IO 2 4 6 8 lO

Frequency (f) - Hz

M = 0.85, altu]52 m(500 ft)AGL

Wtm 128 369kg(283 000 Ib),A= 65 °

Figure 42. SMCS performance with relocated sensor, PSD of lateral

load factor at FS 746.8(294) pilot station,case i.

74

-- Run 11.1, vert gain = 0.0, 1at gain = 0.0.... Run 11.6, vert gain = 1.5, lat gain = 1.5............ Run 11.7, vert gain = 1.9, lat gain = 2.2

SCAS on

200

xl2oo

xlO-5

160 ....

160 _nz

120

120 g2

g2 (ft/sec) 2(m/sec)2 .rad/sec

rad/sec 808o

'140 i

40

0 00 Io 20 30 40 50 60

= Frequency (_) - rad/sec' I I I I I I I I'" I I I

0 2 4 6 8 10

Frequency (f) - Hz

M = 0.85, altB152m(500 ft)AGLWts 128 369kg(283 000 Ib),A= 65°

Figure 43. SMCS performancewith relocatedsensor PSD of verticalload factorat FS 746.8(294)- pilot station,case 2.

75

Run If.l, vert gain = 0.0, lat gain = 0.0

Run ll.4, vert gain = 0.0, lat gain = 1.5

.......... Run ll.5, vert gain = 0.0, lat gain = 2.2

SCAS on --28 ....

28 x

x10-5

244%

24" "eQeeee

..

20 "'

20 _m. : '__LIZ

O_v :;!2 16

2 16 g(ft/sec) _

(m/see) 2 rad/sec !ill i T,,,i

rad/sec 1212

P,I

/\ I:._

': "'" "'",; "...-_ - .-

0 0 I',-. I i .....0 I0 20. 30 40 50 60

Frequency ((_) - rad/secI I I, i_ I ,,, I , _ m.... I0 2 ...4 _ 6_. 8 I0

Frequency (f) - Hz

M = 0.85, alt--152m(500 ft)AGL

Wt =128 369 kg(283 000 Ib),A= 65 °_

Figure 44. - SMCS performancewith relocatedsensor,PSD of lateralload factor at FS 515.6(203)- sensor location,case i.

76

Run 11.I vert gain = O.0, lat gain O.O

------Run 11.4 vert gain = 0.0, lat gain 1.5

.......... Run II.5 vert gain = 0.0, fat gain 2.2SCAS on

0.28

xlO -6O.28xlO-5

_0.24

0.24

eege

._ e • 60.20 : .':.

2 (_2V • "._--g 2 - "'".Q ..(m/sec)rad/sec 2 -!I :

0.16 _ 2 _I| "',ft/sec) ,_.i_;I "

rad/sec :| [ •

o.12 :J I :O. 12 ;I I •

l %..:\ e

O.08 \_- '

o_::o.04 .._\0.04

: Y

•e

O O I0 10 20 30 40 50 60

Frequency (_) -rad/sec -L_ I, I . , .A " • -_ .' i I !0 2 4 6 8 lO

Frequency (f) - hz

M = 0.85, alt B152 m(500 ft) AGL

Wt =128 369kg(283 000 Ib),!%-- 65°

Figure 45. - SMCS performance with relocated sensor, PSD of lateral

load factor at FS 746.8(294) pilot station,case 2.

77

This large responsedifferencebetweentwo fuselagestationswhich arerelativelyclose is not reasonable. Figure 45 data are believed to be thedata in error. Subsequentto flight 1-56, the accelerometerat the pilotstationwas found to be unable to hold a calibrationand was replaced.Becausethe pilot stationresponseis of prime importancein ride qualityevaluations,it was requestedthat the data of flight 1-55 be rerun. Thiscould not be accomplishedin the schedule;however,the data of this flightdo have limitedvalue and are presentedhere for completeness.

Figures46 and 47 are similarto the data of figures44 and 45, respec-tively,but with the vertical SMCS on togetherwith the lateral SMCS.

All of the data in figures41 through47 were obtainedat an aircraftweight of approximately128 369 kilograms(283 000 pounds). Data similartofigures44 and 45 are shown in figures48 and 49, respectively,for a weightof approximately120 204 kilograms(265 000 pounds). Comparisonsof thesedata indicatemore adversehigh-frequencymode couplingwith increasinglateralgains for the lighterweight configuration.

It is concludedthat the frequencyresponsedata show a significantimprovementin pilot-stationresponsedue to relocationof the lateralaccelerometer. The PSD data of the lateralresponseatthe pilot station,however, still show considerablehigh-frequencymode excitation. Taking allevidenceinto account, the lateralSMCS still appearsto providea netlateralresponseimprovementfor the nominallateralgain of I.5.

TRUNCATED ANALYTICAL MODELS

As previouslymentioned,the flexibleanalyticalmodel of the B-I usedin SMCS designanalyses has been describedusing nomalized vibrationmodes

of the structure. The model used I0 symmetricand 12antiso_ymetricmodes inmost analyses. In de_ng a_]xrti.cg_l_modglstO supportre?ringbase simulatorstudies,however,it was found that this number of modes caused computerequT_m-e_iYYequirementsto become excessive. The challengewas to retain theaccurateaeroelasticimpacton short-periodand Dutch-rollcharacteristicsand the main essenceof the structuraldynamicmotion as seen at the pilotstationand SCAS and SMCS sensors.

78

Run II.1, vert gain = O.0, fat gain = 0.OIII Run II.6, vert gain = 1.5, lat gain = 1.5

........Run ll.7, vert gain = 1.9, fat gain = 2.2-- SCAS on

_6 '"

._ 28- x I0-5

x 1024

24

2O

20 _ny

g2 16

(m/sec)'2 ft/sec)2

8

,\ r.4 4 .i

• ° I e

0 0 l0 20 30 40 50 60

' .,. Frequency (0Ji - rad/sec

I I I I " I I , I I I ' -'0 2 4 6 8 |0

J

Frequency (f) - Hz

M = 0.85, alt _152 m (500 ft) AGL

Wt _ 128 369 kg (283 000 Ib), A = 65°.


load factor at FS 515.6 (203) - sensor location, case 2.

79

Run II.I, vert gain = O.O, lat gain = O.O

..m_ Run 11.6, vert gain = 1.5, lat gain = 1.5_..........Run ll.7, vert gain = 1.9, fat gain = 2.2::

• . SCAS on ;' .:0.28 , , ::

0.28 x 10 6 ." : :: .-5 :i!! "

xlO : :,.I i •

0.24 : ::,0.24 '.

0.20 jI :0 20

_ j •o.16 I I:2

2°'16 g _. I:g 22 (ft/sec) _,: li

(mlsec) radlsec I: ! i

rad/sec O. 12 I; _.0.12\-\'.

0.08 0.08 1_ij_il:I: I-

0.0/4 0.04 I.

I! ,,I:

%,G" Z ""• e

o o0 i-O 2-_ 30 ;40 ,50 60

.Frequenc, ) (oJ) - rad2._ecI I I I ,I I :I" _" I I ' - •

0 2 4 6 8 lO

Frequency (f) - Hz

M = 0.85, alt _ 152 m (500 ft) AGL

Wt = 128 369 kg (283 000 Ib), A = 65°

Figure 47. SMCS performance with relocated sensor, PSD of lateralload factor at FS 746.8 (294) - pilot station, case 3.

8O

Run 11.8, vert gain = 0.0, 1at gain = 0.0Run 11.11, vert gain = 0.0, 1at gain = 1.5

......... Run 11.12, vert gain = 0.0, 1at gain = 2.2SCAS on

28 x .........-6

"- 28 x lO ,.

-5 ":I0 "

24 !_

24 _ '.

y • I i!

2 16

g (ft/sec)(m/sec)2 rad/sec

rad/sec 12

12t

|.

8 \8 :\

00 10 20 30 40 50 60

Frequency (m) - rad/sec

I I I I I I I I I I !0 2 4 6 8 lO

Frequency (f) - Hz

M = 0.85, alt _ 152 m (500 ft) AGLWt _ 120 204 kg (265 000 Ib), A : 65°


load factor at FS 515.6 (203) - sensor location, case 3.

81

Run 11.8, vert gain = O.0, lat gain = O.00.72 Run 11.11, vert gain = 0.O, lat gain = 1.5

x 10-5 0.64 _. .......... Run 11.12, vert gain = O.O, 1at gain = 2.2-

x 10-6 SCAS on ..oe

0.64 ""eo

e •

56 "'0 . m • o

0.56 - ••

O.48-:'.: •

0.48 - • :• •: •

0.40 - ,. :

0.40 - _ny

2 0.32-g

(m/sec) 2 IIo

rad/sec 2 0.24 - 1: - !_0.24- g 1: ::

(ft/sec) 2 I!rad/seCo.16 'li :_ :o.]6- I: I

j;

I: k0.08 - 0.08 \ :

0 _ 0 _ I I I0 10 20 30 40 50 60

Frequency (_) - rad/secI I I i I i I l I I Io .2 4 6 8 lO

Frequency (f) - Hz

M = 0.85, alt _ i52 m (5OO ft) AGL

jWt = 120 204 kg (265 000 lb),./L= 65°

Figure 49. - S_S performancewith relocatedsensor,PSD of lateralload factor at FS 746.8 (294) - pilot station, case 4.

82

l)ynamic analyses Were conducted to identify the key modes contribut.[ngto dynamic motion at the pilot station (the SMCSsensor was nearby) andSCAS sensors located at the nominal center-of-gravity (CG) location. Thetechnique will be illustrated using the longitudinal-symnetric case. rnthis instance, the normal acceleration frequency responses at the pilQtstation and SCASsensors, due to excitation by the horizontal tail, were

- employed as criteria to judge the degree of accuracy achieved with truncatedmodels.

Thus, given a flexibleaircraftdescribedby rigid-bodymodes plus anumber of structuraldynamicmodes, the structuraldynamicmode set istruncatedwhile none of the quasistaticaeroelasticeffectsof the eliminatedmodes are lost.

The data used to modify the aerodynamic derivatives are generated by themethod described herein and are identified as F_/Roratios, or [ IF/[ IN"The approach to generating the F/R ratios is as follows:

(1) Select sufficient modes to represent accurately the dynamiccharacteristics of the real system.

(2) Assume those modes not selectedfor the dynamicsimulationto bequasistatic.

(3) Excite the quasistaticmodes with the aerodynamicloadingsassociatedwith the dynamicmode,S control-surfacedisplacements,and rigid-bodydisplace-ments. The loads picked up in each mode are"determinedby solvingthesimultaneousmodal equations.

(4) The solutionsto the precedingequationsprovidethe informationnecessaryto calculatethe F/R ratios used to correctthe aerodynamicderiva-tives of the rigid-bodyand structuraldynamicmodes selectedfor thesimulation.

The equationsto be used in the exampleare as follows:

Rigid-BodyPlungeand Pitch ModesT

83

Symmetric Structural Mbdes, i = 1 to n

Before approaching the details of defining the specifics of the truncatedsimulation, a brief description: of the basic P/Rtechnique is discussed. Ifall of the n structural equations were eliminated but the aeroelastic impact onthe short period retained, the equations would appear as follows:

Rigid-Body Plunge and Pitch Mbdes

84

The F/R ratios st_wn, where all 10 structure modes are ilwolved, areobta ined as follows. The set of structural mode equations are set up asshown below and _1 through _10 are solved for each of the indicated unit

loadings of _, (_w/2Vo) &, (Vw/2Vo)q, and _, respectively.

] il i I i! I II I I. I , , ' i

As an exampleof how the F/R ratios are developed,considerthe derivative

CN_. From the normal force equationwhere _ = 1.0 and all other rigid-bodyvariablesare zero, the followingrelationshipis obtained.

nl through nlO were obtained from the simultaneoussolutionof the previous

e_uationset-_ora loadingCni_ for a unit value of _.

The expressionis reformedto obtain the F/R ratio.

85

The F/R ratios

are obtained in a similarmanner.

The extensionto the pitching-momentcoefficientsfollowsthe same lineof logic.

Considernow the longitudinal-symmetricequationswhere three structuralmodes (I, 3, and 5) of the i0 are dynamic,but the quasistaticflexibilityeffectsof the eliminatedmodes are retained. The bar over the coefficients

indicatesa modificationdue to the quasistaticeffectsof structuralmodes

2, 4, and 6 throughI0. TypicallyCN = ([CN]/[CN])CN, where CN is for a rigidvehicle.

Rigid-BodyPlungeand Pitch Modes

Symmetricstructuralmodes, i = i, 3, 5

86

For mode 1 (typical of modes 3 and 5 also)

• The F/R ratio correctionsfor these equationsare obtained in the follow-

_ng manner. Ass_ning that structuralmodes I, 3, .and5 will be dynamic,thesimultaneousequationsto be solved for unit loadingsappear as follows:

87

The F/P, ratio data for CN and Cm coefficients,given the solutionof theseequationsfor unit loadingsof _, (_w/2Vo)&, (_w/2Vo)q, and 6, are

obtainedas previouslyexplained. The F/R ratio corrections,to CNni, CN_i,

Cmni, and Cm_. are obtained as follows:l

where 72, 74 and 76 through710 are obtained from the simultaneoussolutionof the previous equationset for unit 71 loading.

c_,zJ c_,<,and

c._3 c_ s

are obtained for unit loadingsof 73 and 75, respectively.

m m

Ccnq' i Cm_s , and -----.Cm_$

Cm_Zi CrowI Cm_/,_

,r"

obtainedusing the pitchingmoment___elationshipsand the 7i solutionsforare

the unit loadingof 71, 73, and _5' respectively.

The

data are obtained in a similarmanner for unit loadingsof (_i/Vo),(_3/Vo),and(_5/Vo).

88

'llm ratios for modifyingthe coefficientsof the dynamicstructuralmodes

are obtainedusing the n2, n4, and n6 throughql0 responsesof the simultaneous

equationsfor unit loadingsof _I' _3' 775'(_i/VO)'(_3/V°)'and (@5/Vo).Typ i ca1] y :

Cn,._.c_,n24+C_,n.7?.Cn,n,o_,O-- %,_,= I 4-__I

cn,n,

Where n2, n4, and n6 through nl0 are from unit nl loadings.

Where n2, n4, and n6 through nl0 are from unit (_l/VO) loadings.

It can be shown that all of the quasistatic aeroelastic information isin the truncated modal equations by using the following logic for a typical.rig id-body aerodynamic co efficient.

A B

f

(F/R)3-mode modified X (F/R)7_mode quasistatic = (F/R)10-mode quasi-system system static original

system

" The followingset of numbersfor a test case of the techniqueillustratesthe accuracyretained in the coefficients.

89

S

F/R F/RCoefficient from A from B

CN_ 0.752514 0.752519

CN& I.015199 1.015201

CNq 0.661091 0.661086

CN_ 0.620383 0.620382

CM_ 0.602100 0.602126

CM& 0.948420 0.948413

Q_ 0.732947 0.732943

CM6 0.588629 0.588754

Figure 50 illustratesthe degree of quasistaticand dynamicaeroelasticinformationretainedin the three structuralmodes plus correctionsset ofequations. Shown is the frequencyresponseplot of the normal accelerationatpilot stationand CG. Note that at zero frequency,the truncatedmodel responseoverlays exactlythe original 10-modemodel response, indicatingthat all ofthe quasistaticinformationof the 10-modemodel has been retained in thetruncatedmodel. Furthermore,the truncatedmodel dynamicresponse is a goodrepresentationof the 10-modemodel dynamicresponse.

Truncateddynamic equationsmay be developedfor the lateral-directional-antisymmetriccase in a manner similarto that shown here for the longitudinal-sy_netriccase.

ANALYSISOF ,SMCSVANEAERODYNAMIC INTERFERENCEEFFECT

During the developmentof the SMCS vane configuration,wind tunnel testswere conductedto determinethe aerodynamiccharacteristicsof the vane.Fairlyextensiveaerodynamicinterferenceeffectswere observed in the forceand moment data for both the longitudinaland lateralcases during componentbuilduptests. Reference1 containssome of these wind tunnel data andanalysisof their sources;refer to this referenceas backgroundfor thematerialto be presentedhere. It is the purposeof this sectionto reportthe resultsof an analyticalstudy made to assess the importanceof the vaneaerodynamicinterferenceeffectson the dynamicsof the aircraftresponse.

9O

0

0 10 20 30 40 50Frequency (_) - rad/sec

lO structural modes,

120 1 _H unsteady aero

1_-_I 1 3 structural modes + corrections,

pilot / _H quasistatic aero80

rad

4O

0

0 10 20 30 40 50

Frequency (_) - rad/sec

M = 0.85 S.L.

Medium weight, A= 65°

Figure 50. - Normal load factordue to horizontaltail frequencyresponsecon_arisonsof full and truncateddynamicanalyticalmodel, SCAS on.

91

The analysesconductedwere made in thefrequencydomainwherenormaland lateral load-factor responses at the vane station were calculated dueto SMCSvane oscillatory deflections at various frequencies. The main reasonfor doing this was that flight-test data existed for these frequency responsesagainst which to check the reality of the interference effects modeled.Another reason was that the basic analytical model was available to conductthis study; only the vane aerodynamic characteristics had to be reworked toinciude the aerodynamic interference effects estimated. The following para-graphs describe how this was done and discusses the resLtlts of the analysesmade. Finally, an evaluation is made of these interference factors relativeto the importance to future similar designs.

Discussions of reference 1 identified general areas where the forcescaused by vane deflections were acting. As shown in figure 51, they were on(1) vane itself, (2) forebody, (3) wing-aft body, and (4) empennage. As asimpi:ification, it was assumed that these forces acted at point locations ineach of the general areas identified. These point locations were determinedfrom the wind tunnel test data of forces and moments for various stages ofconfigurat ion buildup.

The SMCSvane alone force and moment coefficient data were determinedanalytically (reference 1) while the associated interference force and momentcoefficients (CN, Cm, Cy, Cg, Cn) were determined directly from the wind-tunneldata. The structural mode generalized forces were calculated for each modeknowing the forces acting at the points defined previously and the structuralmode deflections at these points (F¢i). However, before any of these datacould be added, the transport time lag effect from the vane to the point ofload impact had to be considered. In the frequency domain, this lag effectwas included for each interference load by mutfiplying by

-1vo

e

whereX is the distance(+aft) fromthevane stationto the pointin question,is the forcing frequency, and Vo the velocity.

TableVI showsthe rigid-bodyand structuralgeneralizedforcecoefficientfonmlationforthe SMCSvane,includingintereferenceeffectsforthelongitudinal-symmetriccase. The SMCSvanelateral-directional-antisymmetricdata,includinginterferenceeffects,were assembledin an analogousmanner.

The frequency response data of figures 22 through 24 and 27 through 29 are

repeatedhere in figures52through57. The normalloadfactorresponsedataof figures52 through54 showthatthe interferenceeffectsmodeledproducea

92

-- Horizontal tail

Vane

Wing-aftbody Forbody+ a L , + a R

(a) Normal forces

Vertical tail

Wing-aftbody

J

Forebody

(b) Lateral forces vane + _R' - _L

Figure 51. - Typicalvane-inducedinterferenceforces.

93

TABLE r1. - S\ICS \":\.\E EFFECTI\E\ESS 1\CLUD1\G 1\TERfERE\CE EFFECTS,LO\G1TUD1\.-li.-SY~I.'IETR1C C\SE, FRECUE\CY DO~l-U\

150 f_,_\ -f %

,oo:.-7,"".,\degPhase i/ / "..!_ /'_/// ":_ _.

angle 5G

/ i I I I ° I I i•.._"-i • --

o L ? 2 : 4! 6 8 . "lo• /_, ' : ". I

-50 _'I_l" "..°,i_' Frequency (f) - Hz "I

Flight test data

Analytical data, no vane aero interference

.......... Analytical data, with vane aero interference12 P,

I " "nz : :"

. a% •o

8 " :

g ": : •

:'" -'::_\ _

/// -,_. /•"JI "- _- -" -" -"",, ..'/" .. ._,:,€---/

• • °" /,' ° .- _.J:; - ,_ _.. _# -. ..0 ""_"_"I_'_i_'"I I I "" I I I I I '

0 2, 4 6 8 10

Frequency (f) - Hz

M = 0.85, alt = 762 m (2500 ft)

Wt = I19 296 kg (263 000 Ib), A= 65 °

Figure 52. - Effect of S_S vane aerodynamicinterferences,frequencyresponseof normal load factor at FS 571.5 (225)due to

SMCS vane deflection,SCAS off, SMCS off.

95

15°

Flight test data

Analytical data, no vane aero interference

..............Analytical data, with vane aero interference

12 . .."

"i :'

nz .. ::e • o •

_CV ' " ' '

• , e

8 " "

rad

4 •

oe° • •

", ..'"0 I I"" I I I I I i0 2 4 6 8 lO

Frequency (f) - Hz

M = 0.85, alt = 762 m (2500 ft)

Wt = ll9 296 kg (263 000 Ib), A= 65°

Figure 53. Effect of SMCS vane aerodynamicinterferences,frequencyresponseof normal load factor at FS 571.5 (225)due to

SMCS vane deflection,SCAS on, SMCS off.

96

•.C,.",xPhase 100 n _ ,, .-'" I °" -.l.._.."| ."" _,q_, :".

eoo • • _ • •

-. angle • ",,'i ; -" °°'_X " t _. ,* "*

deg 50 : I _' _'.\. : J/ _ .'IV'"

•e 2 ; 4: 6 8 "',. 0

"'" Frequency (f) - Hz "'_-50

Flight test dataAnalytical data, no vane aero interference

..............Analytical data, with vane aero interference :".:..

e el •

• ".12 ' '

: .# e

nz , ,

8 .... : ": "_

grad

o°°

4

00 2 4 6 8 O

Frequency (f) - Hz

M = 0.85, Air = 762 m (2500 ft)

Wt = I19 296 kg (263 000 Ib), A_= 65°

Figure 54. Effect of SMCS vane aerodynamic interferences, frequencyresponse of normal load factor at FS 571.5 (225) due to

SMCS vane deflection, SCAS on, SMCS on.

97

Flight test data

-- Analytical data, no vane aero interference................Analytical data, with vane aero interference

4OO

300 !"_ "/', :\ .

200 l _"" - _',r..--. ,,,,,_.

anglePhaseI00 / "J/"_ I_.

_; Frequency (f) - Hzdeg 0 "- ' ! I I • _la,... I I I 1

....... _vL-W_ 02 ; 4 6

-200-l O0 " %%_._ _ ..........j

owI i il I

lO

ji o

8

I6

rad

4k/

Ii:

I: _

I: t i

;.....;.-• "_, a-.-_" • n I I I00- "-- ..2 4 6 8 0

Frequency (f) - HzM = 0.85, alt = 762 m (2500 ft)Wt = 119 296 kg (263 000 lb), A= 65°

Figure 55. - Effect of SMCS vane aerodynamic interferences, frequencyresponse of lateral load factor at FS 571.5 (225) due to

SMCS vane deflection, SCAS off, SMCS off.

98

Flight test data

Analytical data, no vane interference

...............Analytical data, with vane aero interference

8

6grad

2

• ooS ••_' i_._ _

00 2 4 6 8 0

Frequency (f) - Hz

M = 0.85, alt = 762 m (2500 ft)

Wt = I19 296 kg (263 000 Ib), A = 65°

Figure 56. - Effect of SMCS vane aerodynamic interferences, frequency

response of lateral load factor at FS 571.5 (225) due toSMCS vane deflection, SCAS on, SMCS off.

99

Flight test data

.... Analytical data, no vane interference

...... Analytical data, with vane aero interference

200

""____-,_lOO "_--_-_ _

Phase _ /-I lb,_ _.

angle- 0 _'______._ ,/ . .__'_,., 0

deg .... 2""_" 4 6_ _ ' _!Frequency (f) - Hz_._

-100 .

-200 % "-

-300 ...." 'lO

8

o,

6

g

rad _I _.

4t

2 _ _

_o_

00 2 4 6 8 I0

Frequency (f) - Hz

M = 0.85, alt = 762 m (2500 ft)

Wt = If9 296 kg (263 000 Ib), _A= 65 °

Figure 57. Effect of SMCS vane aerodynamicinterferences,frequencyresponseof lateralload factor at FS 571.5 (225)due to

SMCS vane deflection,SCAS on, SMCS on.

I00

significantdegradationof the analyticalto flight-test-datamatches inamplitudeand phase. A similar,but less severe degradationof the lateralload factoranalyticalto flight-testdata matches due to inclusionof theinterferencesis indicatedin figures 55 through 57.

Why better agreementswere not obtainedof analyticalresults (with inter-"" ferenceeffects included)to flight-testdata is difficultto assess. First,

however, it is to be noted that the interferencedata were obtainedusingstaticmeasurementtechniques,whereasthe flight-testdata were dynamic. Thevane on the 0.036-scaleforce model from which most of the componentbuildup

data were obtainedproduced small force readingsof questionableaccuracyforthis type of study (momentdata were judged more satisfactory). The 0.1-scaleforebodymodel was judged to produceacceptableforce and moment data. This,then, impliesthe need for more expensiveand larger scale models tested,using dynamictechniquesin order to supportaccurateanalyticalvane aero-dynamicinterferencestudies. The analyticalto flight-test-datacomparisonswithout interferenceeffects displayedin the referencedfiguresare quite

good, particularlyfor the normal load factors,and suggestthat such anexpensiveapproach is not warrantedto obtain satisfactoryaerodynamicdesigndata. A key conclusionreachedas a result of this interferenceanalysis isthat a static 0.l-scaleforebodymodel is adequate for obtainingvane aero-

dynamics data that includethe significantlocal fuselageinterferenceeffects;this model producedthe basic data used in obtainingthe best analyticalto flight-testdata match displayedin this report.

IMPACTOF SMCS ON SELECTEDLOADS

BACKGROUNDOF USING SMCS IN THE B-I FATIGUEANALYSIS

Becausethe SMCS systemwas designedto be fail-safe,not fail-operational,

the originalB-I structuraldesignconceptwas that the aircraftwould havefull structuralintegritywith or without the SMCS. That is to say, thestructuralloads for both designand fatigueanalysesof the air£ramewere

: to be done with the SCAS operativeand the SMCS inoperative. The B-I SCASwas designed as triple-redundant,fail-operational,and (forthe gust loadsanalysis)fully effectiveat all times.

During the B-I design development,studiesof the expected airframeloadswere accomplishedusing the availableweight, stiffness,and controlsystemsdata as it evolved. Severalyears into the program, it was found that filterchanges in the symmetricSCAS were having a considerableimpacton the forebodyfatiguespectrum. Furtheranalyses indicatedthat operatingthe SMCS mini-mized the effectsof the SCAS changesand gave forebodyloads of a more consis-

tant magnitude. Also, the lower statisticalload levels computedwith the

I01

SM(_ ;letirewere thoughtto be more representativeof those that would besee, i_ serviceusage. In view of these facts and with Air Force concurrence,it wCJsdecidedthat the B-I synmetricfatigueanalyses for the low-altitude

imnetrationmission segmentswould be accomplishedwith the SMCS operating.l:atigueanalysesfor all other flight conditionsand all design limit loadanalysescontinuedto be performedwith only the SCAS activated.

GUST LOADS ANALYSIS DESCRIFFION

As with the ride quality analysisdescribedearlier, the dynamicgustloads analysisperformedwas a generalizedmodal analysiswhere the equationsof motionwere solved in the frequencydomain. This analysisused 14 modes of

motion: plunge and pitch rigid-bodymodes and 12 symmetricfree-freenormalstructuralmodes. Also includedwere two active controlsurfaces: the hori-zontalstabilizerand the structuralmode controlvane.

For the B-I aircraftfatigueanalysis,the missions for the expectedserviceusage were each divided into segments. Flight conditionsrepresentingthe mission segmentswere selectedfor analysis. For each flight condition,staticaeroelastictrim loads (to be used as mean load levels)and gust

dynamicresponse incranentalloads were computed. Both the trim loads and gustincrementalloads were issued for fatigueand fracturemechanicsanalysesas distributedgrid loads. Along with the gust loads were the estimatedtimeto be spent at the flight conditionand a graph of the expectedload exceedancesper hour flight.

Aeroelasticloads for the steady-flighttrim conditionswere computed

using wind tunnel-derivednonlinearaerodynamicdata for the rigid andtheoreticallycomputedaeroelasticincrementloads. These static aeroelasticcalculationswere performedwith structuralinfluencecoefficientsfixed at apoint near the aircraftCG.

The gust response equationsof motion that were used are presentedasfollowsin matrix notation (ref. 6). See the appendixfor symbol definitions.

102

Frequencyresponsefunctionsof the generalizedcoordinates,q, wereobtainedby solvingthe equationsfor a unit sinusoidalgust inputat each ofi00 equallyspaced frequencies.

GUST L(I_DSSTRUCTURALMDDEL

A completeaircraft structuralmodel was derivedfrom structuralflexibil-ity influencecoefficientsgeneratedby the B-1 InternalLoads Group usingfinite-elementmethods. These SIC's,which were used for both staticand

dynamicanalyses,were for partiallybuckled skins so as to be representativeof the aircraftstiffnessat the design-limitload level. On the lifting

surfaces,the SIC points were arranged in streamwiserows to make it easy tocomputethe required slopesand deflectionsfor the aeroelasticanalyses.Figure 58 shows the arrangementof SIC poin_ts,and table VII presents thestructuraldegreesof freedomused in this model.

TABLE VII.- GUST LOADS MODEL STRUCTURALDEGREESOF FREEDOM

Motion Type

Component X Y Z @x @y @z

Fuselage,fixed wing and fairings 57Nacelle, includingengines 4 6 17 2 2 2Moveablewing 4 4 58llorizontalstabilizer 2 45

Total = 203 Z i0 I0 177 2 2 2

Normal elasticmodes, free from rigid-bodyplunge and pitch constraints,were computedby the method of reference 7, as shown in the followingequation.

where:-i T

103

'. GUST LOADS AERODYNAMICS

Aerodynamicgeneralizedforces were developedusing the DoubletLatticemethod with the wing-fuselage-horizontal-tailcombinationrun at one time.l:igure58 presentsa diagram of the aerodynamicpanel arrangement. Geometryadjustmentswere made to the theoreticalmodel to improvethe correlationof the zero-frequencystabilityderivativesand pressuredistributionswithwind-tunneltest results. One particularadjustmentwas to leave a small gapat the root of the horizontalstabilizer. At eachMach number,the width of

the gap was adjustedto get the best match of stabilizerCL_ and spanwiseloading. For the SMCS vane, the aerodynamicswere run separatelyusing thevane modeledwith 30 boxes (5 chordwiseand 6 spanwise)in the form of asymmetricinvertedV. These data were correlatedagainst the averagewind-tunnel test data. Generalizedvane forcesdue to modal motion and gust werethen added to those from the wing-fuselage-horizontalcombination. No attemptwas made to simulateany vortex wake producedby the vane.

Deflectionsused to define the structuralshape for aerodynamiccalcula-tions were selectedfrom the SIC points availablein the gust analysismodel. •It was found to be necessaryto review the shape of each mode to insurethatthe deflectionschosen defineda smooth shape with no sudden changesor rever-ses in slope. Points eliminatedrepresentedconcentratedmass it_ns andfairing-nacelledeflectionsthat were not representativeof the true wing-bodystreamwiseshape.

Generalizedforcesdue to modal, vane, and horizontal-tailmotions and aunit sinusoidalvertical gust were computedat zero frequencyand eightfrequenciesbetween zero and approximatelyII Hz. In the processof computing•the aerodynamics,the downwash inductionmatrices generatedwere retainedonmagnetic files so that they could be used again with differentmode shapes,thus savingconsiderablecomputingcost.

At each of the i00 solutionfrequenciesrequired,generalizedaerodynamicforceswere obtainedby spline-curvefits of real and imaginaryparts of thegeneralizedforces computedat the nine frequenciesas previouslydescribed.To improvethe qualityof the spline interpolationsof the gust forces,the °gust referencepoint was transferredto a point just forwardof the aircraftaerodynamiccenter. This transfertends to flatten the curves;i.e., reducesthe rate of oscillationof the functionswith frequency. After fitting,thegust referencepoint was transferredback to the nose of the aircraft. Forconvenience,the gust referencepoint is normalllyplaced at the nose of theaircraftso that when the solutionfrequencyresponse functionsare used tocompute time historiesdue to a discretegust input,the penetrationof thegust starts at time equalszero.

1.04

Fuselage station centimeters

500 lO00 1500 2000 2500 3000 3500 4000 4500I I I I I I I I I I

Inches

200 400 600 800 1000 1200 1400 1600 1800

Component Panel No. of boxes _____'._'_..'_"_*I*/

Forebody 1 6 _I_

Body + wing stub 2 40

Outer wing 3 6

Outer wing 4 30 *SIC points used to define aero shape

Wingtip 5 12 +SIC points not used to define aero shapeHorizontal 6 6

Horizontal 7 36

136

Y motion SIC points --

FRL . t t , T. . +

oFigure 58. Loads analysis SIC point locationsand DoubletLattice geometry.

l:inaladjustments were made by direct ratioing to bring the generalized

forces in the rigid-body modes, due to rigid-body motions and gust, into

agreement with the available wind tunnel measured values of CL_ and CM_.I:actorsdeveloped in the region of zero frequency were applied to the fullrange of solution frequencies.

GUST LOADS CONTROLSYSTI_V_

Descriptionsof the active control systems(ref. 8) used in this gustanalys[sare presentedin figure 59. Also shown is the method used to formthe transformationmatrix, IT],for control-systemsfeedback,relatingcontrol-surfacedeflectionsto the generalizedcoordinates.

LOAD METHOD DISOJSSION

Gust responseloads were computedat selectedstructuralstations. (Seefigure 60.) The mode displacementmethod was used at each frequencyof solu-tion to computedistributedloads at the SIC points. Shears,moments,andtorqueswere then computedby finite sL_mnationsof the loads or the loads

tilnesthe appropriatemoment arms. Figure 61 presents the matrix equationsused to performthe load computations. The mode displacementmethod finds aset of net externalloads that are equivalentto the loads requiredto holdthe structureinthe deflectedshape attained (ref. 3). To obtain accurateresults,a linear superpositionof the elasticmodes that are used in thesolutionmust give a good descriptionof the structuraldeflections. This

requiresthat at least severalmodes of vibrationsthat are primaryto eachstructuralcomponentbe included. Modes selected shouldnot be restricted infrequencyto the maximumFourier frequencyof the solution,but shouldbechosen to obtain all significantcontributionsto the structuraldeflections.

Althoughthe final loads for structuralanalysiswere issued as grid ormass point loads, it was necessaryto compute shears and moments in order to

keep the ntmber of items used in the load calculationsto a manageablesize.On each structuralcomponent,a referenceaxis and load stationswere estab-lished for computingthe shears,moments, and torques. The LOAD GEOM matrixcontainedone row for each load item. Basicallythe LOAD GEOM row elements_Jre: for a shear, ones in the columnsrequiredto select loads outboardof theload stationaxis, and for a moment or torque,the arms from the load stationor referenceaxis to each required load point.

To improvethe accuracyof the computedshears,moments, and torques,each load acting on a SIC point was consideredto he a pressure evenlydistributedover a load box around the point. For any load box cut by a loadstationaxis, only the box area outboardof the axis was consideredin

106

Symmetric SCAS at SL, Khp -- .36

Compensation Notch Low pass Servo Actuator

Degrees

Symmetric SMCS for M = .85 at SL, K_o= .565

Notch Shaping Washout Actuators

Degrees

The matrix representing the control systems in the solution equationswas formed as:

Gains x polynomials Mode defl.ectionsand slopesevaluated at S = j0_ at sen_or locations

7-n_c_ ec_ _,l 41

-nzc_ 0 .n_.c cuz /

Where: CG = FS 2649 cm (1043 in), CV = FS 516 cm (203 in), _'=

o

t:igure 59. Control systems for gust loads..

FS 2367 FS 2896

oc (932 (l 40)

FS 922 FS 1377 FS 1872 FS 2520 FS 3366 FS 4018 (1582)

(363) (542) (737) (992) (1325) Horizontal tail pivot

i I' I I' Pivot

Dimensions: cm (in.)

+S Forebody Aftbody

ZFB

, Yt FB I I _ iI II I I

Figure 60. - Structuralstations for gust loads analysis.

I

c

Point load at each mass point, by mode displacement method

Calculation of loads; shears, moments & torques

NLI_I NLIj N_I

Acceleration at point i

NMjI

Pitch rate at point i

=

i

" Control surface deflections Where: N = Number of mass pointsNM = Number of modes, rigid

_u = T(_J) ta) NEM = Number of elastic modes_V(_){ NLI = Number of S, M, & T load

2.aI Z_KIM NM_I iterns

Figure 61. - Load calculationequations.

109

computing t:he load. The ._ment and torsion arms for this-load were taken totile centroid o1 the outboard area.

The partitioningof the load boxes was computermechanizedby subdividinga box cut by a load stationaxis into I00 small boxes (i0 equal division on aside). Area ratio and centroid coordinatesof each subboxwere computedusingas data the coordinatesof the load box corner points. The area ratio wasthe ratio of the subboxarea to the load box area. Coordinatesof each subbox

centro[dwere then checked, and all those outboardof the load stationwereretained. The retained subboxarea ratios and area ratio momentswere accunu-lated to obtain the total elementvalues for the LOAD GEOM matrix.

Other load items computedwere load factorsat selectedstations,pitchFate and acceleration,and control-surfacedeflections.

Frequencyresponse functionsresultingfrom these load item calculationswere computerplotted for visual review and also saved on magnetic files foruse in computingthe load exceedancecurves and in-phasecomponentloadcond i t ions.

GUST STATISTICALLOAD CALCULATIONS

For each load item, the turbulenceresponse power spectrtm,_o(_),responsequantitiesA and No, and a frequencyof exceedanceof load curve werecomputed. The methods used were as presentedin reference9 and as follows:

2

o

3600 VN -

0

Ii0

The Von Karman continuousturbulencespectrumwas used with the gustscale L equal to 152.4 meters (500feet). Also used were the sea,level vertical

gust parametersfor low-levelcontourflying- Pl = 1.0, bI = 0.823 m/sec(2.70 ft/sec),P2 = 1 × 10-5, and b2 = 3.246 m/sec (10.65Tt/sec). In the cal-

_.. culationsto computeA and No, the requiredintegrationswere performedfromzero throughthe highestfrequencyof the analysis,approximatelyii Hz.

• The load level that could be expectedto be exceededone time per airplanelife in each mission segmentwas determinedfor each load item. This was done

to define a set of loads that could be used to developdistributedloadingconditionsfor the fatigueanalysis. These load values were read from the

exceedancecurve of each item at one over the total number of flighthours forthe mission segment. By nature,all of these loads were positivenumbers,and

althoughthey were consideredto representa load-cyclecondition,the phasingrelationshipsbetween the loads were unknown.

LOAD PHASING

The problemof load phasingwas solvedby applicationof the correlationcoefficientmethod to producewhat are referredto as in-phasecomponentloadconditions. Referencei0 developedand used the correlationcoefficientto

expressthe statisticalcorrelationbetweentwo gust responseparameters.

o

For this gust load analysis,correlationwas developedbetweenall ofthe load items. Correlationcoefficientswere not actuallycomputed;but a

matrix [B],the elementsof which were proportionalto Pij_i_j,was developed.- Here i and j indicateload item numbersand, thus, the corresponding[B] matrix

row and column numbers.

[B]= Real [H(_)] I_w(_)l[H_(_)I_]

iii

l'_chcolumn of [B]was then normalizedon the diagonal elementand

weightedw_th the correspondingload item expectedvalue.

where:

i=j

Columnsof the resulting [S]matrix representload conditionswhere thediagonal elementsare the load item expectedvalues and the off-diagonalelementsare the statisticallyin-phasecomponentsof the other load items.

CONDITIONMATCHING

To developdistributedgrid loads that matched the in-phasecomponentload conditionsthe mode displacementmethod was again used. As shown infigure 61, this method can be used to computedistributedloads for a givenset of generalizedcoordinates. For this matching problem,the shears,moments, and torquesat each load stationwere known, and the solutionhad tobe made for the generalizedcoordinates. The load generationmatrix wasdefined as:

Then the coordinatesand the shear,moment, and torque loads from the [S]matrix are related as:

LOAD] hSLOADS = GEN

112

_is equation is usually overdetermined,and the coordinatesare solvedfor in a least-squaressense. Due to the large differencesin magnitudeoftileshears and moments involved,the solutionresultsgenerallyprovidea poor

load match. To improvethe solutionquality,each load item was weightedso that it had equal magnitudeand, thus, equal significancein the solution.

. By multiplyingthroughby a diagonalmatrix of one over the loads, the followingequationwas obtained.

['°Ih1 = LOADS GEN

Taking one columnfrom the IS]matrix at a time, a generalleast-squaressolutionprogramwas used to solve for the generalizedcoordinates. Distributedi_int loads for the in-phasematchingconditionswere then formedby the modedtsplacementmethod.

EXAMPLE CONDITION LOAD RESULTS

Load resultsfor a B-I aircraftlow-altitudepenetrationconditionof

M '--0.85, altitude = SL, using sweep --67.5 degrees,and weight = 140 614 kilo-

grams (310 000 pounds)are presentedin table VIII. Also, presented infigures62 throughi00 are plots of the load-itemfrequency-responsefunctions,power spectrums,and load exceedancecurves for the items listed:

(i) Wing, WS 985 (387.6), bending moment

(2) Forebody,FS 1377 (542),bending moment

(3) Forebody, FS 2367 (932), bending moment

- (4) Normal load factor, nz, at CG

- (5) Normal load factor, nz, at pilot

(6) Delta (deflection) of horizontal stabilizer

(7) Delta (deflection)of SMCS vane (SMCSon)

113

'I'AI_I,I!VIII. I,OADCOMPARISON,SMCSOFF VERSUSSMCSON

lixpected loads, increments from trim, one o_ccurrence per 1000 hours• M = 0.85 at SL, wing sweep = 67.5°, wt = 140 616 kg (310 000 ib)

SCAS on

Loads

l,oad :items SMCSoff SMCSon

Shear WS 1341 (528) 38 330 (8 617) 39 149 (8 801)Bend. morn. 67 232 (49 588) 69 567 (51 310)Torsion ,' 3 243 (2 392) 3 246 (2 394)

Shear WS 985 (387.6) 80 406 (18 076) 78 622 (17 675)Bend. morn. I 278 977 (205 764) 280 389 (206 805)Torsion } 11 977 (8 834) 11 413 (8 418)Shear WS566 (223) 123 958 (27 867) 117 842 (26 492)Bend. morn. 698 319 (515 056) 681 015 (502 293)

Torsion _, 20 984 (15 477) 19.975 (14 733)Shear at wing pivot 171 710 (38 602) 155 567 (34 973)Roll morn. 649 463 (479 022 612 953 (452 093)

Pitch morn. 'i 1 355 627 (999 864) 1 280 991 (944 815)Shear at IrFroot 58 681 (13 192) 60 189 (13 531)

Roll morn. 237 236 (174 977) 238 159 (175 658)l'itchmorn. ,r 126 113 (93 017) 127 587 (94 104)

I:B morn. FS 922 (363) 381 315 (281 245) 188 692 (139 173)FB mom. FS 1377 (542) 1 060 369 (782 092) 349 942 (258 105)I;I_morn. FS 1872 (737) 1 848 893 (i 363 680) 364 677 (268 973)

I.B lllonl, tzS 2367 (932) 2 261 236 (1 667 810) 627 525 (462 841)AB morn. I-S2520 (992) 4 106 793 (3 029 030) 2 366 297 (I 745 300)

AB mom. FS 2896 (1140) 2 438 061 (i 798 230) 1 I00 992 (812 054)AB mom. FS 3366 (1325) 983 160 (725 145) 603 152 (444 864)

~

Nacelle Sz 85 953 (19 323) 91 366 (20 540) -

Anz at (:(; 0.956 0.949

An z at pilot 2.007 1.041 -horizontal, degrees 0.699 0.728

mode wine, degrees - 19.955

Stations, cm (in.)Shears, N (lb)bloments and torques, N-m (lb-ft)

1.14

-. N-m Ib-ft

m/sec ft/sec

14000. , ! ,

60 000 13ooo. I

,L,ooo. !! ' ''II I I

llO00. IIII ° I ,

'0000.40 000 M 9o00.8000.

'700o.6000, . ,

_ooo. ,/\_

20 000 4ooo.3ooo./f'_-j _' - \ I'2000. I :

1000. _ ._.

0 o0 20 40 ' 60

180.00

,oooo140.00 1120.00

!00.00

80.00 K

_.oo \ \P 40.0o\ _. \

H 20.00_ \ \

As o ',, \E -20.O0 \

"-40.00

-60.O0

-80.O0 X-loo.oo \

:: -120.00 _ !

-140.00

-160.00 t- " -180.00

20 q0 60

FREQUENCY, RAD IANSISECOND

Figure 62. SMCS off, wing bendingmoment frequencyresponse -WS 985 cm (387.6in.).

115

(:_(,b_ft)2m/sec/ ift/secrad/m rad/ft

I .OXlO+G9 , \

,o_- ,.ox,o._ \ Ill .................. _ III ....... -

\ T r'- 1 I I\ /I I zzz

r__z III ...........

P

V

,o,- , !lfl ",I HII

I]H.....'I.OXIO+05

I II I

0 .02 .0_ .06

SPAT|_ FREQUENCYRAD/FT

' ' ' ' ' 2'00 .04 .08 .12 .16 .

rad/m

Figure 63. SMCS off, wing bending moment response power spectrum -WS 985 cm (387.6 in.).

116

-. l.OXlO+03

l.OXlO+02

l.OXlO.01

EX I.OXIO +00CEEDANCES I.OXIO -01

/

H0UR

l.OXlO -02

l.OXlO-o3

I.OXIO-04

z 40000 80000 120000 160000 200000 240000 280000

13rr_NDING14C)IHEN'TAT NING STATION 387.6 lb-ft

• I ! i I

" 1O0000 .200000• 300000 .400000N-m

Figure 64. - SMCS off, wing bendingmoment exceedancesWS 985 cm (387.6inl).

117

N-m Ib-ft

m/sec ft/sec

150000

140000

6 x 105 13oooo120000

I10000

I00000

90000

4 x I05 i 8ooooTI VO000

oU 60000E

500001"

2 × lO_ 4oooo

ooo20000 ,-,

,oooo L/\0 o_ "----------__ _ _--------

0 20 40 60180.00

l_.oo14o.oo "\ 1120.00I00.00

_.0o

60.00 i

40.00 \

H 20.00 IA

_ o

-20. O0 t

-_o.oo I

-60.00-8o.ooi-I00.00

-120.00

-lqO.O0 \ I-)oo.oo r_ k_ I-180.00

20 40 60

FREQUENCY, RADIANSISECOND

Figure 65. SMCS off, forebodybendingmoment frequencyresponse -FS 1377 cm _542 in.).

II_

Figure 66. - SMCS off, forebodybendingmoment responsepower spectrum -FS 1377 cm (542 in.).

119

l.OXlO+02

l.OXlO.01

EX 1.0XI0 .00¢EEOAN

' CES 1.0Xl0 -01

/

H0UR

1.0Xl0 -02

1.0Xl0 -03

I.OXIO-04

4.0XIO+I 8.OXIO"

F'0REBDOYBO_[:)ING_ AT FUSELAGESTATIOXl_-t2 I b-ft

' 06 ' 06 ' 60.4 x 1 0.8 x 1 1.2 x lO N-m ""

Figure 67. SMCS off, forebodybendingmoment exceedances-FS 1377 cm (542in.).

120

N-m Ib-ft

-. m/see ft/sec_0000

300000

_0000

1.2 x lO6 eeoooo_0000

220000

M 2000O0

06 AG 180000

0.8 x 1 _ ,6oooo140000

120000I00000

20000

O- o_--_" __" _-___O 20 40 60

180.00

160.00 _

140.00

120.00

I00.00

80.00

6o.oo\

H 20.00AS 0E -20.00

-40.00 _ \_0.00 __

-80.00 _

-I00.00

-120.00

.. -140.00-160.00

-180.00 \20 40 60

FREQUENCY, RADIANS/SECOND

Figure 68. - SMCS off, forebodybendingmoment frequencyresponse -FS 2367 cm (932in.).

121

Figure 69. - SMCS of, forebody bending moment response power spectrum -FS 2367 cm (932 in.).

122

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1111111111111!111111111111111111 III11

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X .OXlO+00

C )z() ( .............. II ...... II t, ' 'E lll:lkllllllll ;]l;:::l .... i

< IllD IIIIIIIIII_IIIIIIIIIIIIIIIIIII IIIIIA IIIIIIIt111_1111111!1111111111 IIIII

Iliililiiliii i_'i illiliIii 'limeD' i'''i_ .o×,o-O,IIIlilflllllt_!ill II Illllll lit:::::::i:i:::lill_llllll iiiiiiiii'iliii_/ 'IIIIIIII,IIIIII[I_ ...... ,,,,,, ....

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........................... iiiii]lill]l;lllll]llllllll]l]illlli_]l]'

,.ox,o-°,I!11111111111I I! IIIIIIIIIII ilqB.OXlO 1.6XlO' 2.4X1_ 06

F_Y B__J',(3INC_I'IOHENTAT FUS_AC_.,ESTATIONg_ 1b-ft

. ° # !

1.2 x 106 2.4 x 106 N-m

Figure 70. - SMCS off, forebodybendingmoment exceedances-FS 2367 cm (932in.).

123

g g

m/see ft/sec

•07000

.06_00

0.20 .06o0o.05500

•05000

M "04500

o. 15 _ .moooNI .0_::)00 |T

O. 10 u 03O00Q

E ._00

ooo/. \O.05 .ol_oo _.01000 W

.00500

0 .00 20 40 60

1130.00

160.00

140.00

120.00I00.00

80.00

60. O0

P 40.00 _

H 20.00AS 0E -20.00

-40.00

-60.O0 X,_-80. O0

-ioo.oo .!-120.00 \

, ooo-160.00

-180.00 1"_"0 20 40 60 -

FREQUENCY, RAD I ANS/SECOND

Figure 71. - SMCS off, normal load factor frequency response - aircraft CG.

[24

-3lO

! .OXlO-04

-4I0

0 .02 .0_ .06

SPATIAL FREQUENCY RAD/FT

f I I I I I

0 0.04 0.08 O. 12 O.16 0.20

rad/m

Figure 72. - SMCS off, normal load factor responsepower spectrum - aircraft C G.

125

l.OXlO +03iii

I

l.OXlO +02

\\

,ox,o.°, I\\

\

\E

X I.OXIO+00 \CE -\Eo \A \Nc \E \S 1.0XlO -01

/ \\

H0 \u \R

l.OXlO-02

\\

\\ ,

1.0XlO -03 \

I,OXlO -04 _0.2 0.4 0.6 0.8 1.0 1.2 1.4

LOADFACTORAT AIRCRAFT CO

Figure 73. SMCS off, normal load factor exceedances-aircraft CG.

1.20

g g

--. m/sec ft/sec

l.2 .35ooo

1.0 .3oooo

.25OOO0.8

MAG .20000N

0.6 IDU .15000E

0.4•I0000

O. 2 .05000 J L0 .0 0 20 40 60

180.00

160.00

140.00

120.00

I00.00

80.00

60.00

P 4o.oo!_ _.H 20.00

As o "X

E -20.00 _ _

_0.00

-60. O0

-80.O0

-mo.oo-120.00

-140.00 \-160.00

-180.00" " 0 20 40 60

FREQUENCY, RAD IANS/SECOND

Figure 74. - SMCS off, normal load factor frequencyresponse-pilot station.

127

rad_m rad/ft -_I.OXlO+00

i'ii ........!111I ........t1111I

,o° '_l"ll .....,.o,,o-o,i/lll

II i 11 I

i| Ill i[I Ill l

___ I Ill tII]]l

I ,! III I-1 I\ ./: tll ....

1o !1! /i t11

o, !1/![ Ill : ..... "_-

P

V _.... t-H_'_ ---2 " -

10 ,_ I11 .... T--,.............° !!1 /! _____-....-

=_ II I / ...........ill _ -

'3 III /

........ t-tl -- --------'----'--2 _---_._'-r_£t I I

t ] I __z: ::__rz_:TI=_-,= -o .02 .0,.t .o6

SPATIAL FREQUEN(;Y RAD/FT

I I I I I I

0 0.04 0.08 0.12 0.16 0.20

rad/m

Figure 75. SMCS off, normal load factorresponsepower spectrum -pilot station.

128

I.OXIO +03

l.OXlO+02

I.OXIO+01

EX l.OXlO +00CEEDANCES l.OXlO -Oj

/

H0UR

l.OXlO -02

l.OXlO -03

. _

I.OXlO-040.4 O.B i .2 1.6 2.0 2.4 2.8

LOAD FACTOR AT PILOT STATION

Figure 76. - SMCS off, normal load factor exceedancespilot station.

129

i

deg degm/sec ft/sec

.06500

O. 20 .06000.05500

.05000

O.15 .msoo#4 .04000A

O.10 I .030o0U

_ .0,2500

O.05 .0,.5o0.01000

•00500

0 .o _---..__ __0 20 40 60180.00

160.O0

140.00

!20.00 _ >

I00.o0

BO.O0

60.00

P 40.00 "_H 20.00AS 0E -20.O0

-40.00

-80.00

-80.00

,oooo / \ /-120.00-140.00

-I_::_).00 _,-IBO.O0

0 20 40 60

FREQUENCY. RAD IANS/SECOND

Figure 77. S_S off, delta horizontalstabilizerfrequencyresponse.

1.30

Figure 78. - SMCS off, delta horizontalstabilizerresponsepower spectrum.

131

I.OXIO+03

I.OXlO +02

I.OXIO+01

EX I.OXlO +00CEE

.DANCES l.OXlO -01

/

H0UR

l.OXlO -02

l.OXlO-03

F£gure 79. - SMCS off, delta horizontal stabilizer exceedances.

1.32

-. N-m Ib-ft

m/sec ft/sec

10000.

40 000 9ooo. I 1

8000.

30 000 M 7000. i

GA 6000.

lIN 5t700. _ /_

20 000 U 4000"://'_-'/3000. !_I0 000 2ooo. _..

1000.

0 o0 20 40 60

180.00

120.00

I00.00 " X

80.00

_°'°° \..... ",H _.oo \ \A ,,,

-2o. oo _-40.00

-60.00

-80.00 __-I0O.00

\ 1-140.00

-160.00

. . -180.00, 0 20 40

FREQUENCY, RADIANS/SECOND

Figure 80. - SMCS on, wing bendingmoment frequencyresponseWS 985 cm _387.6in.).

133

( N-m )2m/.secrad/m

(lb-ft )2ft/secrad/ft

J •OX 10+09

I.OXIO+08

oUTPUT

I OXIO+07P .5o

J .OXIO+06

\--I- -

\\

1/-f-

II -,

I' II-

--\ -

-1\

- -

-

-\

---.-

-\.,

- ~_.._- -- -

1\

-

\o .02 .04 .05

SPATIM. FRElU:t-.CY R/lD/FT

o 0.04 0.08 0.12

rad/m

0.16 0.20

134

rib~rc 81. - SMCS on, wing bending moment response power spectrum WS 985 cm (387.6 in.).

I. OXl0-°3

i

- - I.OXlO-°4100000 21_0000 300000

eENOmO.OMe_TATUmO STATION_37.6 Ib-ft

I I I I

lO0 000 200 000 30"0 000 400 000

N-m

Figure 82. - SMCS on,.wing bendingmoment exceedancesFS 985 cm _387.6in.).

135

N-m Ib-ft

m/sec ft/sec

40000.

35000.

1.5 x I0530000.

M _00.A

1.0 x 105 G

_ 20000.U

,,.o.Jr,t0.5 x 105 IOOOO. j_

_ooo. / "h

o0 0 20 40 60

180.00

i_o.oo _ ',.,

140.00 _,

120.00

lOO.O0

130.00 _,

\40.00

H 20.00A

°-20. O0

-40.00

-6°.00

-BO.O0

-I00.00-leo.oo \

J

-140.00 i _ k-160.00 _,-180.00

0 20 40 60

FREQUENCY, RAD I ANS/SECONO

Figure 83. - SMCSon, forebody bending moment frequency response -FS 1377 cm (542 in.).

136

(

2N-m

m/sec)rad/m

(lb-ft)2

ft/secrad/ft

I.OXIO+09

I.OXIO+08

oUTPUTp5o

1.0XIO+07

I.OXIO+06

l- - - .

1- -

IIIV\

- l- -I----1- - I-

,I-I-

I\y.-- - --

II II

1\1/

- -1-- - ,J..., -l--1 -

- -1\ -

- -

\

1\

\.-

-

1\1- -

- \1\

1\- - - ..

l-

I-- ~\I-- ''1,.0

o .02 .04

SPATlM... F"RECUN:V fW)/FT

.06

oI

0.04I

0.08. I

0.12

rad/m I

I

0.16 0.20

Figure 84. - SMCS on, forebody bending moment response power spectrum FS 1377 em (542 in.).

137

I. OXI0+03

~

,°

I. 0X10+02

1.0Xl0 +01

EX 1.0XI0 +00CEEDANcE \

• 5 1.0XlO -01

/

H0UR

1.0Xl0 -02

!

1.0Xl0 -03

i

I.ox10-°4 i100000 200000 ,.300000

FOF_EY3OYBO',OINGMOMENTAT _ STATION 542lb-ft

II I I I I

100 000 200 000 300 000 400 000 500 000

N-m

Figure 85. - SMCS on, forebodybendingmoment exceedances-FS 1377 cm (542 in.).

138

N-m Ib-ft

m/sec ft/sec"

2.0 x 105

I

45000.!

40000.

,.s× '°s . _oooo._°°"1 _NGk:_O0.

1.0 x 105 u_ 2oooo.D

0.5 x 105 loooo.

0 o _0 20 40 60

IBO.O0

160.00

140.00 L

120.00

I00.00

BO.O0

60.00

40.00

H 20.00AS 0

E -20.O0

_0.00

.oo 1

-_. O0

-I00.00

, oo ,\ /_.\-140.00

" -,_o.oo _., "\-,_.oo \

0 2O 40 60

-" FREQUENCY, RADIANS/SECOND

Figure 86. SMCS on, forebodybendingmoment frequencyresponse -FS 2367 cm (932in.).

139

,-

f-- - -f-

1\

J 1\ \\ II /~

-- f-

- - ---

1\ \.. - - . -

- - f-0- . -

"K- ._.-

\

-1- - - --

-f-

\-- -. f-

f- - - ~--

- \\-.

~r- -- .

- -

1.0XIO+06

I.OXIO+09

I.OXIO+oa

(Ib-ft)2

ft/secrad/ft

I.OXIO+ 10

ou~ I.OXIO+07

UT

P5o

610

( N-m )2m/secrad/m

0 .02 .O't .06

5PAT1~ FR£Cll.EN::Y RAD/FT -

'- , I I I I0 0.04 0.08 0.12 0.16 0.20

rad/m

Figure 87. - SMCS on, forebody bending moment response power spectrum FS 2367 cm (932 in.).

lila

I.OXIO+03

l.OXlO+02

1.0XI0+01

EX 1.0XI0+00CEEOANCES 1.0Xl0-01

I

N0UR

1.0Xl0-02

I.OXIO -03

I .OXlO-04100000 200000 300000 400000 ..500000 600000

Fg__.E_OYBO'-DING_ AT _ STATION 932 I b-ft

• i _- I ..... I !" 200 000 400 000 600 000 800 000 N-m

~n

Figure 88. - SMCS on, forebodybendingmoment exceedances-FS 2367 cm _932 in.).

141

, g g

m/sec ft/sec

.08000 _.

o.25 .o_o /!•07000

.06500

O. 20 .o6ooo. [Y5500

HA"05000

o.15 No ._oo [%/

I .04000Tu .03500

O. 10 _ .o_oo.02500

O.05 .o15oo \ _ X.oloooI "'xJ

.00500__ _ _-

0 .00 20 40 GO

180.00

160.00 1

140.00

120.00

I00.00

80.00

60.00

40.00

H 20.00A

-20. O0

-40.00

• "-60.00

-80.00 _N-I00.00

-120.00 -N._. ,,.

-1_o.oo _ %-160.00-18o.oo \ N

0 _ 40 GO

FREQUENCY. RADIANS/SECOND

Figure 89. - SMCS on, normal load factor frequencyresponse-aircraft CG.

1.42

.06.04.02o

1\-l- f-

- ~ - - - I- f-l- f-

1\

1\-- --~

2 1\-

- i- f-

-.-

1\- r-

II.i' I\U !\ IA

.-

-- ~

1\

II1\1\

i-

- 1\ i-iI,

~ \

\. - -- i-

- - -~ 1-. -1\ .. -- - - -

--1\ 1;< - -~. -~ ~

1\ i-

.. -t-.~ i-

" .... 1...-

1.0XIO-04

I.OXIO-O

oUT~ I.OXIO-03

T

P5o

(~2rad/ft

-210

-110

rad/m

SPATlf.L FREa..EN:Y RAD/FT

oI

0.04I

0.08 0.12 0.16 0.20

rad/m

Figure 90. - SMCS on, normal load factor response power spectrum aircraft CG.

i43

I.OXlO.03

I.OXIO+02

l.OXlO+01

EX l.OXlO+00CEEDANCES I.OXIO-01

I

H0UR

l.OXlO-02

l.OXlO-03

1.0XlO -°q

0.2 0.4 0.6 O.B 1.0 1.2 1.4- .

LOA_)FACT_ AT AIF_I:_-T CG

Figure 91. - SMCS on, normal load factor exceedances- aircraft CG.

144

g g

m/sec ft/sec

O. __00

.08000

.07000

O. 10 .o3ooo

ooo/ %.01000

0 .0 0 20 40 6OI_.00

160.00 I

I

140. O0 1 _ '120.00

I00.00J t i

80.00i i

60.OO

40.O0p t l , iH _ .00A t is o \/-,,,E __.00 ___.

-40.O0i

' \--60.00-80. O0

-100.00-120.00

-140. O0

-I_.00-180.00

" O 20 40 6O

FREOUENCY, RAD I ANS/SECONO

Figure 92. SMCS on, normal load factorfrequencyresponse - pilot station.

145

g 2 g 2m/sec ft/-sec

F-

10-1_ ' _

l.OXlO -02 /

-2 _ __i0

o

y i,OXlO_O3P ~_UT

P

O I

-3- /lo

i.OXlO -04

m

-410

o .02 .oq .06

SPATIAL FREQUENCY RADIFT

| I I I . .. I I

0 0.04 0.08 0.12 0.16 0.20

radim

Figure 93. - SMCS on, normal load factor responsepower spect_Jm - pilot station.

140

I.OXlO+03

\

I.OXIO+02

\

l.OXlO+01 \

\

i

\ JEX l.OXlO+00

[ \ 'Eo \A \Nc kE \S 1.0Xl0 -01

/ \\

H

o \uR

|.OXlO-02

l.OXlO-03

i

I .OXlO-°h0.2 0.4 0.6 0.8 1.0 1.2 1.4

LOAD FACTORAT PILOT STATION

Figure 94. - SMCS on, normal load factor exceedances-pilot station.

147

deg degm/sec ft/sec

0.1 0 - .o3ooo

o_ooA J i

// ".02600

O.08 .o24oo .•OP_2OO

.02000J

0.06 _ .o16oo,NO .olr_o : .

i .01400

0.04 u .oI_o ,E

.01000

.OOSO0 _ {

0.02 .oo6oo , I •.oo4oo I.0_0

0 .o "'_ _ , T_ ..-.L20 40 60180.00

160.00 _

140.00

120.00I00.00

130.00

60.00 1

o.oo !p -

20.O0 /H

A 0

[-_1.00

-40.00

-60.00-_.OO

-I00.00

_,_.oo \_-140.00-I_.00 \

-IBO.000 20 40

FREQUENCY, RADIANSISECOND

Figure 95. SMCS on, delta horizontalstabilizerfrequencyresponse.

1.48

/ deg _2 f deg ,_2m/sec _ft/sec/\ra--_m/ -r_--_T-

lo-1 _L_'X -

, x_

!

:-___\,OXlO_ I

o \ ---_v ......_--"........ _,.o×,o-O_ / Ip ..... / \

1o-5 s -- \ \ / \ I,.ox,o-°_ / - i "T

............. _.___-_-¥,=___--__-,-,_,_.-.-

,ox,oO, ! r.j.k-- ' -_--_-t:

10-7- _ .....I-

f.OXlO-°8

_=1==!:_--- -0 .o2 .1_._ .o6

,_PATI._L FREQUENCY RAD/F'T

L _l i I t

0 0.04 0.'08 0.12 0.16' 0.20rad/m

Figure 96. - SMCS on, delta horizontalstabilizerresponsepower spectrum.

149

150

EI.OXlO+00X

CEE0ANCE

J .OXIO-015

I

H0UR

I.OX1O-02 1\

oo...TA ~'ZCNT~ ST.6BILIZER. CE~S

Figure 97. - SMCS on, delta horizontal stabilizer exceedances.

m/sec ft/sec

1.7000

" - 1.6000

1.50001.4000 A

I.3000 /_

/I

1.20001'4 I. I000

1.ooo0

3 _ o.9ooo0.8000

g o.7ooo

0._00 I0.4000

I 0.3o000.2000

O. I000 _ __

0 o 40 60180.00

160.00 _,_

140.00 _

IL='O.O0 _ "-.,1oo.oo _ \

80.00

40.00PH c_.obA

o-;20. O0 N_-4O. 00

-60.00 "_-80.00 _

-I00.00

-120.00 _

-140.00-IBO.O0

-180.00'_ 0 2O 40 6O

FREQUENCY, RAOIANSISECOND

Figure 98. - SMCS on, delta mode controlvane frequencyresponse.

151

( deg _2 (deg _2\mlsec/ \ftlsec/rad/m rad/ft _.

I .OXlO.01 "

lO1

I.OXIO+Do

0I0

I .OXlO-010UTP

-] ulO T

PsO

I .OXlO-02

-210

l.OXlO -03

-3lO

0 .02 .0'-I• .06

SPATIAL FEQUEI',_Y' R,_/FT

I I I I I f

0 0.04 0.08 O. 12 O. 16 0.20

rad/m

Figure 99. - SMCS on, delta mode controlvaneresponsepower spectrum.

152

I.OXIO+03I_ Ill;....... [:::::11 I]I: !':_l[: ill]

[llj Ijjlljjj ijll ' _ill tljl

IIII ....................IIII]111 1tll I111 lilt1111 IIIIIIII IIII Itll I!ll

_llil lillilli Illi iili lili,.ox,o+o__t! t! tl IIli ltl i ii

_ ikii: _ iiiiiiii _; iiii iiii__!_iiii...................... IIIII I _1 11111111 I I I I I I I IItlM II111111 IIII ]111 Ill]til_ lltlllll_ ll,l Iljl t111illl' liltilil ltli Ilil IIII

,.=o+O,Illi' Ililli II Il! II::iiii::,.iiiiii!! .................llll TIlL z 1111

.................. III]IIII IIIIIIII Illl IIII

IIII \11111111 Ill! Ill! IIII

TITT i_lllill Illl litl IIIt

!!!! .... ,,!!]:lle IIII IIIi_III :-iiii .... :III ....I111 Ill I1"1 IIII IIII ]Illo IIII IIIll\lll I1tl Illl III1A IIII Illlt_ll lilt till IIitn Illi Illill_l fill IIII liltC

_,-°×,°-°',..! Illi1, I il! .it# ll:i iiiiiii[},i"iiii .... !!ii _:' iiii

IIII I1111111 IIII Ill[

IIII IIII111! IIII IIII illlH llll I lllll II It It Illl _. IIII0 IIII Iit11111 \ till Itll IIIIu IIII iliiilli ,ill! lill liliR

,.o×,o-°_Jill I1!1 N I J!! IJ......................... lil :

tr,,I III, _ i i ' ' i , ,\, I I II 1111 "ilil ........

'"' IIlitl ItlltlJt ll!i \ Itll IIIlli iillllti Itli _illl li

,.ox,o-°_II Ililti il "N. !!!till [Itzl[tl Tltl ._ tlll1111 IllllIlt IIll lIlt _ ill/

I111 IIIII!ll IIII II[I _ IIIIIIII IIIIIIII IIII t111 nb,.]llJill Illillll Iltl Ilil t f't.,.]

'. IIII IIIIIIII IIII IIII III1",.ox,o-_Ilil IIIIIIII III Jill III 7-_

4 8 12 18 20 _

" " DO-TA _ _ V,N_E, CE_:_E_S

Figure i00. SMCS on, delta mode controlvane exceedances.

153

The tabularload summaryresultsand the plots are presentedfor both_ICS off (gains= 0), and SMCS on. In eithercase, the SCAS was consideredto be operatingnormally.

As can be seen in the summarytable,the effectof the active SMCS was

to substantiallylower the fuselageforebodyexpectedbendingmoments. Alsoloweredwere the aft fuselagebendingmoments. The most obvious effectofthe &ICS, as can be seen in the forebody frequencyresponseand PSD plots, wasto considerablysuppressthe responseof the fuselage first bendingmode. Atthe same time, the operationof the SMCS caused increasedresponsesin someof the higher frequencystructuralmodes. The effectof this increasedmodal

activity,as seen in the fuselageand wing frequencyresponseplots, was tocause slight increasesin the wing tip, nacelle,and horizontaltail expectedloads.

Althoughthe effectof :theSMCS activitydramaticallyreducedthe magnitudeof the forebodyfatigueloads for low-altitudepenetration,the total forebodyfatiguespectrumswere not so drasticallyreduced. The SMCS was not usedduring the high-altitudecruise mission segments. The large numberof flighthours spent cruisingand the higher wing lift curve slope at a 25-degreewingsweep, cause the expectedloads to be relativelyhigh. Thus, with thereductionin the low-altitudepenetrationloads, the cruise conditionsbecamedominant in the forebodyfatigueload spectrum.

SMCSVANE EFFECT ON INLET/ENGINECHARACTERISTICS

]he objectiveof this section is to brieflydescribe and summarizeB-I

flight-testresultsthat identifyeffectsof ingestingvortices generatedby the SMCS vanes into the inlet. Although vorticesgeneratedby the SMCSvanes were ingestedfrequently,no engine incidentsrelatingto operationof the SMCS were identifiedduring the approximately1200 flighthours accu-mulated to date. This program includedmore than 200 flightswith threeaircraftand 29 engines. Wind-tunnelresultswith sub- and full-scalemodelswere summarizedin referenceI.

Portionsof specificflightswere dedicatedto demonstratingoperationalsuitabilityduring aircraftmaneuverswith the SMCS vanes deflected. Emphasiswas placed on exploringcombinationsof SMCS vanes deflectionangles and air-craft maneuversduringoperation at Mach 0.85. Effectson inlet total-pressurerecoveryand engine-facedistortionwere measuredby a 40-probe instrumentationmatrix at the inlet/engineaerodynamicinterfaceplane (AIP). Instrumentation,includingan automaticin-flightcalibrationsystem,and flight-testproceduresare described. Resultsare summarizedto doct:nentthis B-I experienceas anaid to futureprograms employingsimilarsystems. Nomenclatureassociatedwith this sectionmay be found in the appendix.

15,1

TEST DESCRIFFION

The B-I propulsionsystem is arrangedin two nacellesunder the fixedportionof the wing as shown in figure I01. Each nacellecontainstwo inde-pendenttwo-dimensional(2-D) externalcompressioninletsand two GeneralElectric F-101 afterburningturbofanengines. Relativelocationsbetween

- the SMCS vanes and inlet nacellesare also shown.

Most flight tests investigatinginlet characteristicswith the SMCS vanesdeflectedwere conductedwith the simulated,fixed-inletconfigurationshown

:infigure102. Ramp configurationfor both inletsconsistsof the initialtworamps set at 7 degrees. On the inboard inlet,the third ramp is set at5 degrees. In the outboardinlet, the third ramp is set at 9 degrees. Smalldifferencesbetween inboardand outboardramp configurationsreflect anattemptto maintain good performancecharacteristicsduringboth subsonicand supersonicoperation. The movable cowl lip is shown in its normal,takeoff,and landingpositions. Duct flaw area distributionsare shown infigure103. Maximumflow area is based on an averagethird ramp angle of7 degrees. Design flow area is shown for referenceand representssupersonicoperationwith a variablegeometry inlet.

Inlet boundary-layerair is removedthroughporous surfaceson the secondmovable ramp, throat panel, and small regionson the upper and lower end plates.The bleed air is collectedin two compartments. The air exits from the forwardcompartmentthroughfixed louversand from the aft compartmentthroughtwo-positiondoors. The doors are open above MachL_l.4and closedat lower speeds.Aft bleed doors were closed during all tests with the SMCS activated.

A bypass systan operatesat supersonicspeedsabove Mach 1.4 to match theinletsupply and enginedemand. The bypass doors open to compensateforreduced engineairflow such as occur on a hot day or during low-powersettings.The bypass doors remainedclosed during all tests with the SMCS activated.

The SMCS vane configurationis definedin referenceI. Flight-testpro-visions includeda black box located in the crew compartmentto drive the

2

SMCS vanes to a fixed deflectionangle. From this position,the vanes couldalso be driven at selectedfrequenciesto oscillate+i0 degrees. Using these

-- provisionsto generate full-scalevane deflections,flight tests were conductedto identifywake ingestioneffectsduring combinedaircraftmaneuversand enginethrottletransientsas measuredat the inlet/engineAIP.

155

Figure 101. - B-1 flight test aircraft, wings swept 65 degrees.

: Throat fixed link--

___'_- fimed

linkBypass

Ramp fixed link

Movable lip__." _:::::`_:_:_i_i_i_i_iiiii_ii!:iiii_!iiii!!ii_iii!iii_i;iiiiiiiiiii_iiiiiii_iiiiiiiiii!i_ii_i_iiiii!i_iiiii!ii_iiiii!i!iii)iiii_iii!i_ii%_iii;!iiiiii_iiiii_!i_

(takeoff/landing) __ _ . _____

"-- " _ _._ _ 2:2:i:i:fi22:2:i2)i:r:2:i:i:i:5:2!iii21211:i[:2:i:r:i:iiii:i21:2:2i:i:2:i::-_2:!:f:!:i:2:i:£_:2!i:ii1_:_:J:!:_:2:!:_ii!ii_:_:i:_:_:J:_:1!i_ii_i_:y:i:i:f:i:2:5:_::::::2:i:i:[:_:i:i:i!_:[£_i_2i_:_:[:J:_:i:_:J:[_i5!!_Ji_:_:i:_:2:J:J:!:2

Fixed throatY_ I°v:_::rI.o.,

Fixed b

_--BLC 1-[ two-position doorsAdjustable link(ground adjustable)

Figure 102. - B-I air inductionsystem.

157

18ooDuct flow area distribution

ll 000 ECI-8 Aerodynamic

interface _Plane_ l

Movable lip inlO000

I10258 cm 2

9000 (1590 in.2)((cm2 9258 cm2

(1435 in.2)8000i

4-1U

"0<[:,

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600C J500O

4ooc I i I a I _ I I I _ I I I l I l I i I80 lO0 12o 140 16o 180 200 220 240 260

in.

I I I I I |__ I I I I I I200 300 cm 400 500 600 700

Nacelle station (at duct _)

° Figure 103. - B-I inlet subsonic diffuserflow areadistribution.

FLIGHT-TEST INSTRUMENTATION

B-1 flLght-testaircraftwere instrumentedto measuremore than 1600

parametersof which approximately700 pertainedto the propulsionsystem.InstrL_entationat the inlet/engineAIP included40 dual-purposeprobes inboth engines in the left-handnacelleto measureboth the steady-stateanddynamic componentsof total pressure. Probeswere installedintegralwitheight engine-inletguide vanes, eachwith five probes locatedat the centerof equal areas. (seefigure 104.)

Computeddistortionparametersare sensitiveto errors in individual

total-pressuremeasur_aents. Several techniqueswere employedduring theflight-testprogramto minimizethese errors. High-response,differentialtransducersmanufacturedby Kulitewere referencedto a duct staticpressureupstreamof the inlet/engineAIP to maintain signalresolutionover a widerange of operatingconditions. The referencepressurewas measured by anaccurate,digital, absolutetransducerinstalledin a conditionedcompartmentin the aircraft. The referencesyst_iwas constructedwith sufficientvolume

and orificesto restrictrapid changesduring aircraftand/or inlet transients.Referencepressurewas sampledfour times per second.

An in-flightcalibrationsystemwas developedto update individualprobecalibrationsonce per minute throughouteach flight. A schematicof thissystem,using a three-waypneumaticvalve for each total-pressureprobe, isshown in figure 105. The valve alternatelysequencesthe back side of thetransducerfrom the normal referencepressure (operatemode) first, to acalibratepressureregulatedto a pressureapproximately5 pounds per squareinch above the referencepressure (calibratemode), and then to the samepressureseen by the front side of the transducer (zeromode) (hencethe nameZOC valves).

The calibrateand zero positionswere each held for 2 seconds,and theoperate positionwas held for the remaining56 seconds. In the zero position,the total pressurebeing sensed is routed to the back side of the transducer

: throughan infinitecoil, approximatedby a coiled line 20 feet in length.This provisionwas necessaryto preventreflectingwaves affectingthe frequencyresponseof the probes and was determinedempirically. Both coils and pneu-

-" matic valves, the latter groupedin gangs of five, are shown installedonthe engine in figure 104.

159

.....Q\o

lguide vanes

Total-pressure probe

Kul ite AlP probes359° aO

IView looking aft 359° aO

I

80°

3au

71°

440

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Engine

B-] j nlet/engine aerOd)l1amicEngine 2

tntcrfacc plane instrumentation.

Fuselage Engine compartment Inlet-- ----I i -II I l II I I l I

zocI' ' m '! Icalibrate Pneumatic ICal + 15 psi I ' ! I _ . 'I actuator

I I Ii l Zero "_ i[ I II

l uperate _ I

Cal I. II ZOC valve I

R2+ 5 psiI I I i • Ii I I

I I _.Infinite coil I Engine I

I R2 . I I I airflow I

Inletl'I I t__ "i I! I Ii

static I ! .AIP Drobej I , ,-

-- I _ ....... , ,_ ___j_-..... Si%n-al--co--ndT ti-one'-r............ I

I P°werII 7.5v supply IICrew station I _. [

i I Cal, zero _s I lTape

rec°rdeff ICBW Operate. _x2__._.j I lHP F _._j_ - III.__R2>10 ps!aTM _-- filter I

' ; R2_ O.4Hz I16 ips I L • Normalize

,.--I' ' iiI "_J ILP [- _rans_u_er 'A/D I LP fi1ter,

J 16 spsi_ filter,0.4Hz II 12HzI_ l' I II I L I

MV multiplexer I

_ Figure 105. - B-I flight-testAIP data acquisitionand signal conditioningsystem.

lhlet_naticvalves were actuatedby a pressure sourceregulatedto 15 poundsper square inchabove the calibratedpressure. Provisionswere includedtovent trapped pressures within the valve to either the engine compartment(ambient pressure) or an engine-face static pressure, depending on the operat-ing condition. Vent pressures were required to be the lowest pressure in thesystem at all times. Selection of vent pressure was accomplished by a crewcompartment switch. A pressurized nitrogen bottle was used as a pressuresource for the actuate and calibrate tanks.

Signal conditioningfor the transduceroutput is also shown in figure 105.l.hcbsignalis initiallypassed througha variable amplifierthat is used toestablisha full-scalerange for the steady-statecomponentat approximately24 pounds per square inch. Zero output levels can also be biased to favornormal operation. Subsequently,signalsare split to record low- and high-frequencycomponentsseparately.

l_w-frequencyor quasi steady-statecomponentsare generatedby low-passfilterstl_ateffectivelyeliminatefrequencycomponentsabove 0.4 Hz (3 decibellevel). Signalsare sampledby a sequentialanalog-to-digitalconverteronboard the aircraftand stored as digitalwords on a tape recorderlocatedin the crew compartment.

lligh-frequencyor dynamiccomponentsare generatedby high-passfilterstlmt effectivelyeliminatefrequencycomponentsbelow 0.4 Hz (3-decibellevel).Signalsare furtheramplifiedbefore passingthroughvoltage-controlledoscillators,multiplexed,and recordedon the same tape recorder (differenttracks)used to record the low-frequencycomponents. With 21 continuousbandwidth (CBI_)signalsmultiplexedper track, four trackswere requiredtorecord the 80 high-frequencysignalsfrom both instrumentedinlets.

I)uringthe zero and calibratesteps, all filters are bypassed,and thetotal signal (minusthe referencepressure)is recordedboth as pulse-codemodulation (PCM)and CBW parameters. Calibrationsequenceis operate-calibrate-zero-operate.Transitionbetween the zero and operatepositionsiml_sesa step change in pressure differentialacross the transducerdiaphragm.This has no essentialimpact on the low-frequencycomponentbecauseof thelow-passfilter. However, the high-passfilterrespondsto thisstep changeand res_altsin a damping characteristicwith a period of approximatelyi0 sec-onds. As a consequence,dynamic data are invalidduring this time, and datarecords for detaileddynamicanalysisare selectedto circumventany problems.Typicaloutput signals, includingcalibrationcycles, are illustratedinfigures106 and 107.

102

Parameter identification

(scale in counts)

PT = (CTSOP-CTSZ) CAL + 6002 CAL = (6002-6631)/(CTSOP-CTSCAL)

"' '" 7"'_oPE.CTSOP I , ,I CTSCAL RATE

_ _3_6 _, ....i_---_I/t-l, I 633CTSZF_ , -r I I ....

*.o _ .... I,..... ,,,[u, ..... !1.... ,I.... L_,..,u.JJ,,,I o.o

6327 ' I I 6332

...."......' _........_ L....I ....... ' I : : I'[ .... : _: ": ' r _ I i i1 L.... i J,., .......... , ..... ..;-.j............ ,., ' '7

, , , "'"_• I I , ' --76328 ...............,_ i_ __ I: CALIBRATE .-,.._I'_I I

6333I I''''......,--1 ] ............... P--i mr......,.o ,.................. L......... _........... ,., ,........................ _............

6329 , i t 6334 .

................_ r _ Zili,i III111111 iiiiiiiii iiii1111 iii iiii Iii i iii l.i I

-Ittl ....i [1t6330 ', t 6335 I I I

I I Ii,O __JIIh,LLIII _'_tlT nit Itl O,I IIIllUll Illllltnl itQmatllln iJll*_

• leO _o _o _oo _0 too loo _o _eo _€

6336

'_-' _ 1 -..:r,,.. ......,_.....r,......L......... I,L. i t .......O.O '' I' IIII ill ill , , , 11_ l, , ,, , , , , , 11 , , , , , , , , _.O ' ' I I ' I 'I '

-LI .....iSi, I u E I END OF DATA RECORD I

6337 ; , , 6342 .................,....._............ -4 i .....

O ............ _ ...... _1 _ ..... "111 , , 0.0 , .................. _ ..... _ I'-| L&

..... I "gt t'4L't,1''6330........... ,, ,, 63_3 ..... 'I.I lima InlllII I'_" hie O.l IIII n l.IJJ lIlIIlnIeIJII.U I1_#

..... 13ltt ,].....I _ I f I I

6339 - "............. I.... _.__.F.... 634-_ ............. [" , _._L/-..... .....

_" ,.,.....................I.....,li....I................ ,., ....................!.....' _* ,_,,,,,

°LAtt !J .... , , ,.... I l" .... I ,',"I I "- j-I,I I._. .ii iiiiiii ii_ ii I.I

• Ill Ill li,OI '_ll log IOl Ill loll i00 Ill

SAP.1PLES SAMPLES

Figure 106. In-flight calibration cycle, total pressuresat aerodynamic interface plane, flight 1-5, 64 sps.

163

_" F--- - -_ ................=--- - ' -:_-- _----- --- ---_ - - --=- ........ ............-------- _ " =O',: - -= : --- = ..... - - -- --- =IL_+4V FS disc

'- _ --- I _--- ......-_- --_- -: :-output (62.5 Hz)6 8 .... -- " -1 sec -- _:-- 32 --- :-_--_:----= ------° _= L=_--_ .--=_..... -_ _ ! - - :-

erate .-Zer Cal __- Ol_e -- " ---_-._- --- ---= -_=- ----=:- ___

V=i o fi, --igh-pass N (High-pass --- .... -i ..... =L_ L- -- -- ___--_--=:__=:-Z-__--- ! - - ---_- _ NO valid CBW data during filter __:

] .... _ .... . ........... =-.................. I _ . --=. _ stabilization time of 6-I0 seconds _---_ I --_=-' .... _= =_T_._-_- .......- .... -- 3 _ _-= ....._ i_----_...... _ -:--_- ---- ___.._-1-._ -=- -1:_ - _ _. - __..... _ _. ___.... _ . =--.:_

...... I_- _ -- _-- _=FT_-__-=- -...... ,.. I _.-z.._.-_-_-_--. -_ -_- .... _- _-- _ ............

I-_6329------- - .... - ...... - -- = =---_- - -_-==_ - ------ ---= ..... -= .... - ........- __. _=__- -_ ..... ___--.... _-_ -- ... __.___----_.: _--:=___

_- [ . .' _- _-- ._-'-=-_.-_-_=_ ==-_..... _ ...... I - _- _-, _ _ -_ -. _ - _.-

+- -- _ I _-- --- "...... - _=-_:_=-'--...... =---_ ....= " - _..................... - - i ...... L ....... -- =...........................

_ ............. _ .....

........_ _-_ ......... =T ..... _ i -- " .......... i r£-.... ____ . L ............ -- 4- ---- =--

- - - -'-_;- ] __Determined from PCM data-_--_ _ ........... --_--- ...... _-_--.... _.-9:- . CTSZ_H ....... _ --_ -....... - --_...... _ - ! - -

I

I m

6330 =_ -- Ap- _ CTSCAL ................ _ _ ....-- _' -=_==_-=-=-----_-_:-_--T_ _.L............................ _ CTSMEAN- __.....=....-_.=:_.... ...-.................................. _. _=-Lq ..... !

_- I , -_I ..... _ I _pJ- + PPCM.__-----=_--_...... - _..... _ _ ._-_. PDYN: (CTS-CTSMEAN) (CTSZ-CTSCAL)__

......... ..__Spur_ous ZOC valve operation ....... _ I_-_,=__- _: - , _ ._ _...... .j=-_ ---.. _ '..... i _ _'==.... - ..... -"....... -_---

__ ' __-. _ _.... , , _ _ ,i--6331 ..... .......". _ ..... _ _ l t_=_ ........ __=-__-=____==__ r...... ]_._.-:-d..-.---! ..... -..... _-

- "_ ........ 4- ...... :---= ....... ] F I _

_Ap.p=_rox t_ime l :27: 14, fl-t- -- I ,- ---_.. _ - _ .... ] __ - ..... --_ ......... _ -- _

_ -.-- =-_,,_ ____. ' " - =----2-.-- -

Figure 107. Repmesentative discriminator output, AIP total-p_essumeCBWdata, £11ght 1-10.

In stmm_ry,particularemphasisduring the flight-testprogramwasplacedon obtainingaccurate,total-pressuresurveysat the inlet/engineAIP.As described,these featuresincludedthe following:

(1) Differentialtransducerswith a floatingreferencepressure

(2) Accuratelymeasuredreferenceand calibratepressures

(3) Automaticin-flightcalibrationprocedures

(4) Signal conditioningto obtain good resolutionof both high- andIow-frequencycomponentsfrom the same transducer

F1ight-testexperiencehas shown that, as one might expect with all thepneumaticand electricalconnections,the systemrequired considerablemainte-nance. However, it is believedthat system accuracyapproachedlevelsavailablewith a well-constructedwind-tunnelprogram,and this seems to beconfirmedby comparisonsbetweenwind-tunneland flight-testresults.

TEST RESULTS

A total of 5 flight-testhours were dedicatedto exploringSMCS vaneeffectson inlet/enginecharacteristics. Flight conditionsand operatingvariablesare summarizedin table IX.

Several computedparametersfrom the 40 total-pressuremeasurementsat theinlet/engineAIP are used to present test results. Recovery (PTI/PT0)is theaverageof the 40 low-responsesignalsreferencedto free-stream,totalpressure.

Several distortionindexesare used and were computedby digitaltech-niques. High-responsesignalswere filteredto 62.5Hz and sampledat 360

samplesper second to representa one-per-revolutionengine-frequencyresponse• (signalsfully attenuatedat 125 Hz). Circumferential(IDC)and radial (IDR)

distortioncomponentswere computedfor each ring (eachgroup of eightAIPpressuresat the same radii) and combinedmathematicallyto form a fanstall-marginindex (IDL). This latter index is normalizedto stall-marginallocations,and thus, values of unity computedfrom the high-responsesignalsidentifydistortionlevels approachingdesign limits. Inlet distor-tion, (I_MAX- PT_,IIN)/PTAVG,was also computed. Record lengthswere in therange between 5 and 20 seconds. Scans resultingin maximumvalues of stall-margin index are generallyused to identifytrends.

165

'I'ABI,IiIX. - SMCSVANEEFFECTSON INLET/ENGINECHARACTERISTICS,FLIGHTTEST INVESTIGATION

Dedicated AlPflt timea instrumentation

Flt no. (min) Objectives/accomplishments status

2-19 65 Initialflight-testinves- Referencepressuretigationswith oscillating transducermal-vanes at 1 Hz, vane angle function,no con-±i0 degrees, u = 3-8 degrees, version to= 0-3 degrees, initial engineeringunits

throttle transients

2-33 45 Static vane deflectionangles ZOC valves mal-in 5 degree increments, functionedon± full scale, u = i, 3, no. 1 eng, data5 degrees, 8 = 0-4 degrees, reducedonly forno. 2 throttletransients no. 2 engine.

2-36 60 Repeat of flight 2-33 Leaks in refer-

- ence systempre-cludedAIP datareduction.

2-37 55 Repeat of flight 2-36 Satisfactory

2-38 60 Max rate throttletransients Satisfactorywith vane deflected20 degreescombinedwith nose left

sideslip -

2-42 20 Conductpushoversand pullups Satisfactorywith vane deflected20 degrees

aTotal _5 hours

166

Engine-face,total-pressurecontoursare used to illustratevariations in

distortionpatterns. High-pressureregions (pressureshigher than average)are shadedon the contourplots. Low-pressureregions (pressureslower thanaverage) are unshaded. The magnitudeof the differenceabove or below theaveragepressurelevel is definedby the number on the contour. The annularregion is formedby protrusionof the enginebullet nose at the AIP.

Envelopesencompassingall combinationsof circumferentialand radial

distortioncomponentsare used to identifythe magnitudeof dynamidactivityand to comparethem to design goals. Two types of time historieshave also

been found useful in documentinginlet turbulencecharacteristics.Analogstrip chartsof the high-responseAIP signalshelp to identifywake ingestionduring transientmaneuvers. Digitaltime historieshelp to define transientconditionsbased on computedparameters.

OscillatingVanes

Initialtests with the SMCS operativewere conductedwith the vanes

deflectedsymmetricallyat nominalangles of +I0 degreesand then oscillated+i0 degreesabout that mean at a frequencyof 1 Hz. Vane deflectionanglethus oscillatedbetween 0 and _+20degrees;the latterrepresentingfull-scaledeflection. Note that with these procedures,the vane is being used as anexcitorand inducessome discomfortto the crew members. With the vane

oscillating,aircraftmaneuversinvolvingcombinedangles of attack and side-slip were performed.

Time historiesof aircraftattitude and vane deflectionangles recordedduring a 30-minutesegmentof flight 2-19 are shown schematicallyin figure 108duringoperationat Mach 0.83. Angle of attackwas varied between 3 and 8

degrees; sideslipangle was varied between 0 and 3 degrees. Tests were con-ducted at positive sideslipangles only to ingestthe wake in the instrumentednacelle.

Flight times where increaseddynamicactivity (identifiedfrom stripcharts of the high-responseAIP instrumentation)could be definitelyattributed

- to SMCS vane deflectionare indicated. Increaseddynamicactivitywas alsonoted at other times; however,effectsof sideslipand vane deflectionanglescould not be definitelyseparated. Ingestionis generallyrestrictedtopositivevane deflectionangles (leadingedge up) greaterthan i0 degreesincombinationwith aircraftsideslipoperation.

One major advantageof oscillatingthe vanes is to producerecognizablewake ingestioneffectsin the data. The 1 Hz vane oscillationbetween 1.5 and

18.5 degreesduring a sideslipmaneuver is shown in figure 109. Time histories

167

Leading

edge up L_.Vaneosc'illation at l Hz20 .::.:;.:r.:: :.:.".-,..:.:: .... :.,_--_-....:::.r:-!::=========================l": ....:-: : . ).:! = ",.: :'".... :..... t'--":'::;' _:(.':-_-::_:"..:-T-7_T ":"T-.-T--F.- " ":-::---:._.-:T:..':_'-".'T,7:" "_'7 • • , ;. _ ":: -" ::"=.-:i-

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u mmO " : 4 : " 8 ....t:; : ]2: .... ... 16 .-i 20 . • 24 28 i.... :..-::-;:-::: - .-.-.:- :-. -:::i ...... " " : :";--:!_-!7: _..:::iE _ " : :"' : " " : " " " Flight time - minutes_I_ _ .... " ............... -- .................. ...... :-_ r': "T-- °e- _) : "" NO. 1 "_- I • : ! • " : .... :" : :':.":::";"::: :.: • ........>'_ _I_. = ....... ,

_ ..... inlet:2. : :___ . _i:,i ::--i.-::: ;:-: Increased A-IP dyr)activity_:-:.-----'_ • .i ] .:i.i due SMCSvane deflection"_ u " No.•- _ ' inlet ............................... "..... ;

__ _ l:igure 108. - AIS/SMCS investigation - £1ight 2-19, Math 0.83.

168

0 ,,, -

x3029RHSMC 0 _ ]Vane - _- l Hz-

I

- SMCS vane deflection angle, clegrees

--20LI I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I !

5 i . |

- i Sideslip angl- .. -.'T '.I . _,degrees.

MlOl2 0Beta '_

-5 I l I If I If I I II II I II I I I III I II I II Ill IIII I II IIII II

2400 , . _ ,- AIP PcMdata

Counts

00:53"40 Flight time- seconds2000 I I I I I I I I I I I I I I I I ! 11 I I I I I I I I I 11 I I I I I I I 11 I I I I I I

O 2 4 6 8 lO

I .... ._-_:..--.._._'_L -_ '__,l .... W-data l.."_:"-:- ......Representative AIP C-B I .......... ; .............. '.......o !L ...... ._..'_'--1 ....

= D6338

: _ High pressure> " F.,_4-J

m D6339

' • ' ......:i_ _ Low pressure: _ , -,, _ i ",._-

Figure 109. - AIS/S_S investigation flight 2-19, Mach 0.83/16 000 feet,

= 3 degrees,vortex ingestion in no. 2 inlet during sideslip operation,PLA = max.

169

of both low- and high-responseAIP signals are shown duringan increase insideslipfrom 2.2 to 2.7 degrees. Wake ingestionis evidentthroughoutthetransient,and turbulencelevels increasesignificantlyat the higher sideslipangle. Although the individualplots are not preciselyaligned in time, lowpressuresgenerallycoincidewith positivevane deflectionangles. Similarresultswere obtainedwith the no. 1 inlet. (Seefigure ii0.) Smalleramplitudesare partiallyattributedto reduced engine airflow indicatedbylower fan correctedspeeds.

I_gineairflowdemand has a significantinfluenceon peak-to-peakampli-tudes. RepresentativeAIP pressuresare shown in figure iii during a throttleburst from IDLE to INTERMEDIATE(maximumairflow)power settings. Peak-to-peak amplitudesat INTERMEDIATEpower are greaterby a factor of 1.7 than theamplitudesrecordedduring IDLE power. Again, the lower pressuresin thecycle appear to coincidewith maximum,positive,vane-deflectionangles.

Static Vane Deflections

Comparisonof resultsfrom tests describedpreviouslyindicatedthatresTJltssimilarto those obtainedwith oscillatingvanes could be obtainedwith the vanes positionedat a constantdeflectionangle. Crew members found

these proceduresless objectionable,and their work load was slightlyreduced.All subsequentflight tests exploringSMCS vane effectson inlet/enginecharac-teristicsduringmaneuversand enginethrottletransientswere conductedinthis manner.

Maneuvers.-Tests conditionsinvestigatedduring flight 2-33 (withSMCSvanes deflected+20 degrees)are summarizedin figure 112. Tests were con-ducted at Mach 0.85 and includedmaneuversat combinedangles of attack and

sideslipwhere the maximum angle of attack was 6 degrees and maximum sideslipangle was 4 degrees. RepresentativeAIP high-frequency,total-pressuresignalsare shown duringmaneuverswhere vane vortex/wakeingestionwas measured inthe no. 2 inlet. Faulty ZOC operationprecludedanalyses of data recorded in J

the no. 1 inlet. A strip-charttrace,recordedduringnormal cruise attitudes,is shown for reference. Aircraft attitudesresultingin ingestionduringflight test generallyagree with those recordedpreviouslyduring wind-tunneltests.

Steady-statetotal-pressurerecovery and distortionparametersare shownin figures113 and 114 as functionsof sideslipangle. Angle of attack is used

as the independentvariable in figure 113 with the vanes deflected 20 degrees.The largestdefect in recovery (approximately0.06) was measured at 6 degreesangle of attack and approximately1.75 degrees of sideslip. Effects of the

wake diminish as sideslipis further increased. At approximately3 degrees

170

100 ------..------r-----,.--------r------,Percent fan speed

Mll12NF

f-

-iSidesl ip angle, degrees

,i-

i="

i- ,

i- t·i- .'i-

,f-, I I I I I I r I r r r r I , , r , r , , , I I I I I I I I I I I I , , J I I I I I I I I

- 20 L1..LLLLLLLLJ-L.L.J....l...1...J...L...l...1......L.J......L.J....l..J...l..J...l..J...l..J...J...L..J...L..J...L..J...L..J...L-1-L....I....l.....I....l....J.......J.:o

5

OLLLLLLLLLLLL.L.L.L.L.L.L.L.L.L...L..L...L..L...L..L...L..J-l...J-l...J-l...J-l..~~~..............................~

20 -----.........-----.....,...----~-----r_---_,

-52800 ------,,------;,.------r-----,------;

Ml012Beta 0

X3029RHSMCVane 0I------~------+----..,---I-----_t_----___t

D6288Counts

AlP PCM signal. I

40 CTS at 0.01 0.4 psi

Figure 110. - AIS/SMCS investigation, flight 2-19, O! = 3 degrees,vortex ingestion in no. 1 inlet during sideslip operation.

171

100 ,..-----.----r---"l:::::===---;t;r-----'

INT

Mll13NF

Percent fan speed

tIDLE

.- Sides 1i pang 1e, degrees::::J

I--

I--

l- tl-

I--

tu-u.-LL.LlJ I I I'll I I I II I I I I I I I I I I 1 II I I 111111111

OLU-l..LLLLLl--L.LLL~__LL.L.LL.LL-L.L.L.L.JL.LL-L.L.L.L.J_.L.L~L..L.l_L.L.l.._l_LL..L..J

20,.------r-----,-----,------,-------,

- 20 LU-l...l....u....l..L.J-I...L.L.IL..J....l-l-L..J...J....L...I..~L_1_l_l...J......L...L...L_L_...............................-'-'-..........'__'_'_...I....I........

5

-52400 r------,-------.--,--..-----:T~--_;:.V""--r___,

X3029RHSMC 0 I-------+----:.---+------+-----+-------tVane

Ml012 0Beta

ps i

Flight time

2 4 6 8 10

• .• -- ... -'. - ·.A. jp -.c"sw- ;i~n.~.l·~. -- ~'. -:1.---... ~ f.-. ---.1.".;; . . . . . 'I --- i ... JJI _.•~ ... ~

[ 06338 -'\C~'(-t~~"'''li~ti"J~~r-~~\''r. ~I.· -- ft11;VWI.~I'~~~----: .Low pores sure -" ---,----1 . 7X _. . -- '_'_'_1 .'--;

!D6339)IR~~~A~~~,:Ji~I~~~ .~~~_. !~\~I_'III~...~.. ~--_.::. ...~.:.f.:" -.~_.__LH, gh pressure . .' I Y: "; .._l~\·

Figure 111. - AIS/SMCS invest;igation, night 2-19,a = 3 degrees,vortex ingestion in no. 2 inlet

Juring IDLE-INT throttle transient.

D6338Counts

172

Fit 2-33 test matrixEnvelope of SMC vane 10 Angle of attack,

wake ingestion from degrees • Normal turbulence Normal AIP dynamic activity

wind tunnel data 8 O Increased turbulence (No vortex ingestion)

....Throttle transients -i_i_-- _ _:i '7-J_'2-_ :_- "--: - -i-"--_:--_

• _ --- ....... :- :

Mo re " _6 _0.3 psi -" - -_ "

4 6347 -_-....:_:-..:_-:_---:-:--._--_-_-_4.-----

...... •..: _: ..... - . -

"!"i!i!iiii:_":.... 20 : 26 : 44

0 oz= 3° _= 0°

-8 -6 -4 -2 2 4 6 8

-2" Sideslip angle, degreest

AIP dynamic activity with vortex ingestion

_:-=--'--= 2.3 psi .... _ .... 2.4 psi ................ :----:'_........... 2.45 psi :_--_

A,P probe _--.......=-_=__.:_"_-__:__L"_-=_.=_':;_-.._;--:_-_--_--:.:::_:"_.............._: _-_,-_......_ _- _--i_ii_._____-_-i._-_.iL----_i i:_:_!--__._.

19:58:00 20:28:I0 20:43:00

e = 5"5 ° _ = 2° _ = 3° _= 4° OZ= 1° /_= 4°

Figure 112• Flight 2-33 test matrix and AlP signals during8ach 0.85 operationwith SMCS vanes deflected20 degrees,

no 2 inlet,RB = 7 degrees,RC = 5 degrees

INo.2 INLETIM= 0.85, alpha = variable, beta = variable

RB = 7 deg, RC = 5 degSMCV = 20'deg

l.O0 Recove ry

0. .......:!:

0.92 .....i i.1i.:......:i. /;!i_.:__..!:: ......i.i:.i. :..

•.. i.--i-::!.::i.-ii-:_!Sym Alpha WIR .i_i...]..i!ii-ii:i:i:.-::

•' : :--:: "i_":i O I deg 297 pps ::'._.::. " : ::: :..---"..............-:.-;.A 3 deg 323 pps .!" ::.-.!'.:::i:[.i:::

• : :. :.:':.:.-:: :; {'1 6 deg 348 pps .i. : .i ...i. iJ.i.:

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I DT :. : .:..'I:. "; ;I. :.'::::.: ::- :::::::::::::::::::::::::::::::::::::: F. : _ :T.:';_ ._:.::.'..

---A'-"'_.........:_. . TT-,-_......".....;'-'+.........._,.... -t---_._'--AL-..T.......*......................."-- -

:::_ ; : _- : : : : : _.;:. ; - ........ : A: :: :::::::-

.---:.+;+--:--.:---=_ i._ +:-::.:H-_ _-::_:_::__:::_::::.?'_::.__?:+:__:_:_ _::: _

- 0 ! 2 3 4 5Indicated angle of sides]ip, beta, degrees

Figure i13. No. 2 inlet, effect of sideslip angle on steady-state inlet

characteristics at SMCV = 20 degrees, flight 2-33.

[74

M= 0.85, alpha = 3 deg, beta = variable

RB = 7 deg, RC = 5 degl.O0 . Recover,'

L........._...-.%. _Z_.:.:Z_:£1ZZ. Z5£Z.!'.-Z.:'-_-

o.96: !• i. ' : : ".__i" "'.:[7:_'.':::':x:..:x:t'.:T_x::t_":. _ ':;.. ;. .! " ' : .:i ..... ,_......

PT1 . ...: .:r.:__l_ i..q.::_:::::b_:}::::!.::::Fr:_::::':.:t..::': 4 : 'i i:V :i:! " ::-

.:: .;.::L:__:L_:l.:j:_Ti:£::i_L?i£_£{. ,::.: ......... r:'::..._ . ::::..:. :::': :.:-:_ .:-:". ,- :::.:- :" :;t:- .-"-: .... :'--,:"%::: "- ::::::::::::::::::::::::::::::: "- :- :!:'-':"::'" - : _'::::::-

0.92 ..--_-.......... -.....: '!....."...............•i'..'.::. '. ....:.... ._. :... :. .....:: ..... SYM SMCV WIR .!::::i::...... _. • . .

;-N_t.; + .:i:> !:.:_;_.._:::i?_::>:.i :-: ,: • '-:-.::! ::.i_"_-_:! _ 0 316 pps-:---i_i"

•" : : i'. ":".'t:...!::.:_'::::'."::-.. '. i l d :: _.... : :"___-.._[. :.!_::{__.:k_:____.-.._._ _ 3.3 eg 316 pps ._-:_-.::.i!_:--:-: -" ::. :-__'_-_'_!::! g ,20 deg 323 Dps :'- :": ::.- :-. ::::::::::::::::::::::::::::::::::: . :_:l:.:_:::.!_k ":::: :.: ::.: :i ,. ..'._

............... I ..... _.... _ "" _....... _........................... _..... _.......• : " [ : ::'":::-:::: I"':"' i::'.:::|:-';" .:.t::-::::-:l -':". I" .': "; "" :: .:.'.. :.: : : : "

0.8 ." _ ............:..........r,-_-_...........:......._:_:._ ...........: . ....-.:;::_.--:- -t:::!::-it!N--!:iStall-margin index :::q.;::":i.:;::-i -.--.-

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0.4 :.:.......:....": ........".........' ...."="_' ..... i " _......i.........

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-".:. "._,':::::.:'i::xi:-::_'::::':::}::::::.:.,t':.::,:r_:*,-_n::::it_'::.:x'::: ;:. ::::::::::::::::::::::: _.::":_.:':". _:_._.A_;:_r%:_:_::_:'_:::-£._T:t.N_._[7_r:_.-_.:_:_:_::N..:.°" :.L.:_:," _.:" _.:::._t:t:: _. ::...... ,......... t .... -..........,.......!.......... _...... -+--:r=.-" I- ........................... • ................

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.................. --I .... I ......... _.... _ .... _............... _.., ....... _.... , ........... | ....... . .................... _........ ' ........ _........ I ........ _"....................... _........ l .............

0.2 :: ::.:.i}::}} :. .:.::_.:::ib:::::_:!.:::!:: Distortion .:-!:. !: :_: ;:::-!-::._::-!:.::.L-::-" ::::.Z :::::::::::::::::::::::::: _::ZZ---L--!-:Z[LL:::Z-_F£f.:-

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- 0 l 2 3 4 5Indicated angle of sideslip, beta, degrees

F]gure 114. No. 2 inlet, effect of sideslipangle on steady-stateinletcharacteristics at various SMC vane angles, flight 2-33.

175

sideslipwake effectsare no longerpresent in the inboard inlet. At lowerangles of attack,sideslipangles greaterthan 3 degrees are requiredbeforewake ingestionbecomes evident.

Effectof vane deflectionangle is shown in figure 114 during operationat3 degreesangle of attack. Pressurerecoveryvalues from this figure indicatea similarloss during sideslipoperationfor both the 13- and 20-degreevanedeflections. No losses in recoveryare evidentwith the vane held in a neutralpos ition.

No definitivetrends are evident in the steady-statedistortionindexes.

floweret,with the measured increase in turbulencepreviouslyshown in figure 112,this is sometimesindicativeof an in-phasecomponentpresent in the flow asshown in the time historiesof total-pressurerecovery (figure115) where SMCSwake ingestionwas most evident. Time historiesof stall-marginindex are alsoshown. Similartraces during operationat normal attitudeswith no vane deflec-tion are shown for comparison.

Resultingenvelopesof circumferentialand radial distortioncomponentsare s}_wn in figure 116. Total pressurecontoursrepresentingmaximumvaluesof stall-marginindex are shown in figure 117. Maximum dynamiceffectscorre-spond to operatingconditionsresultingin lowest total-pressurerecoveryat6 degreesangle-of-attackand 2 degreesof sideslip. Maximum stall-marginindexwas 0.74, and the associatedtotal-pressurecontour shows a well-developed,low-pressureregion in the hub.

An extensive series of tests with the SMCS vanes was also conducted duringflight 2-37. Performance and distortion characteristics are summarized for the

11o.2 inlet at normal aircraft attitudes (_ = 2.7 degrees, _ = 0 degrees) infigure I18 for the complete range of SMCS vane deflection angles from 20 to

-20 degrees. No indications of wake ingestion are evident.

A seriesof tests investigatingSMCS vane effectsduring sideslipopera-tLon were conductedover a range of angle of attacks. Resultsare generallyshown for both inlets,and in each case, resultsare shown with the vane at aneutral positionto help separatevane deflectioneffectsfrom sideslipeffects.Re_Its at a nominal1 degree angle-of-attackare shown in figures119 through123, 2.5 degreesangle-of-attackin figures124 through126, and 6.0 degrees _-angle-of-attackin figures127 through132.

Generally,these figures illustratethat the higher the angle-of-attack,the less sidesliprequiredfor the wake to enter the inlet and affect thedistortionpatterns. Data recorded during tests at 6 degreesangle of attackare also used to illustrateadditionaldata analysistechniquesbesides thedigitaltime historiesof quasi steady-statemeasured and computedparameters.

170

, . , .

a= 5.5° {3 = 2° SMC vane =+20°

PT1/PT~ ,0613:40.8 11 '111 ~ I

Fi Itered to 62.5 Hzsampled at 360 sps

a= 3° {3= 0° SMC vane = 0°

PTi IPT:,: l... 0 0 0 0 0[J"""", I, " , '" "I

20: I 9: 57 I .0 2.0Time - seconds'

1. 0 ...----,.-- ---r-----r----,

IDL ~J\'1',~t~':~:\V!f~, \'i/\>-'~tJ.l.I, \Wt,\I,r'lf\'0.0

19: 57: 49 I .0 2.0Time - seconds

a= 3° f3 = 4° SMC vane =+20°

PTl/PT:':ttt=u

2.01.0

1. 0 r----,---;r:- dIDL

o.0 ..........-W-L-Iu....L.J.-L~U-LLLLl..I....L :: L1-1..L1-1.

20:43:03

Time - seconds Time - seconds

Figure 115. - Time histories of total-pressure recovery and stall-margin index, Mach 0.85,with-~~S vane deflected 0 and 20 degrees; no. 2 inlet, RB = 7 degrees, RC = 5 degrees, flight 2-33.

o.og.oo 0.02 0.0" 0.06 O.OB 0.10 0.12 D....IDC

IDC IDC

Figure 116. - Dynamic circumferential and radial distortion components, Mach 0.85,with SMCS vane deflected 0 and 20 dep,rcc-~~no. 2 inlet, RB = 7 degrees, flight 2-33.

, ,

Pressure above PTI

Filtered to 62.5 Hz in shaded area

sampled at 360 sps Pressure below PTImax IDL scans in unshaded area

,number onPTI - PTI [I+ _ contour ) _ ]

Use + if contour in shaded area

Use - if contour in u_shaded area= 3°

SHC vane = 20°

Figure 117. - Dynamic total pressure contours during Mach 0.85 operation with SMCS

vane deflected deflected 0 and 20 de_rees, no. 2 inlet, RB = 7 degrees, RC = 5 degrees, flight 2-33.

Recovery Flight 2-37

1.00 ?._i_i::i::.::.-.: ::.:: _-i::.:!::!-ii ,!:: . " -:-: . : • ".................................. : • : T.. ::...: ,, • i ....

0.98 ?:];::?i:i-_"-i!i_i_i_iiii;:i!iT-; I NO. 2 INLET I i!71:!]!:. !i=:::7:_i:ii=:i;i:!ii_i: ;

:=LJ--.L__:-_ RB = 7 deg_ RC = 5 deg _-............' _ . "I. : : i_.;..;...PT1 _:-:.. -- -.PTO .... .-! _. M = 0.85, alpha = 2.7 deg, beta =Odeg. -.;: : '...

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.-:_'-% ;:i.=_4_-__;ii_P_;ii.%__:_ :__:_t_:_ i%_._L %:_4i_"

". ':::.'. =:'-:-.:'t:=. ::::::::::::::::::::::::: u::.::".;::'::.'::::: : ::i-:F:_.-T::::I.::2-.'_TT-:T:IDT 0.1 .......:.... :: .,._::::_::"'':":-"::;:::=:::::_:.:...... _..:: .......... _. : ===================-===========t=====,, .......

::::h :.:ii:.]i::::!i."_ ...... :-:;:::.='::'v ::::!!::':::::l':::]:7"h::'!:!:!:: .':i !": i' _ . :." :............................. 'LLI:-_:--2. LZI...L.L: "_-i2:=J%__.L." "' -- • " L =.= " __.-'_ .... ;_. "......... _: -F:J.............,.... _ .......T ......-.--_.._ ...... - -_4: ::: __::!_:=:f_ =::..;m"_..::.A.: !:,,_::::_!.:G).I_.:_::.::.-._:: "

...... =- _ ..... l ........ _v ........... ----_,,i. _ . _ -- " ....: " ':"_:"::-_ :::_:--::=i:i::::i.i:!!i--.:i:" i"i:t-::_'::=i::!-!.:? -:::i!-'

: i-::::f:!'iir-:;':-.-.!:::!-::T:_:_---, ....................... _....... t:-_--:::: .....:i ................... ! ................................. " ......

-20 -10 0 10 20

Structural mode control vane deflection angle, deg-rees

Figure 118. - Effects of SMCvane position on no. 2 inlet steady-state anddynamic characteristics, Mach= 0.85 and alpha = 2.7 degrees, flight 2-37.

[8O

Mo (PEG) I

0-

-" 0.5 JII LI[1111I_ 11111 L.LLLL -2 I_I_.L.L.L.J

ALTITUDE (DEG)(K FEET)

o IILAI''' ,,,_,_i,, ,,llJllll k_,,,,,,J ,_,,i,1 .... ,, ,, .5 IIItllll_JllJlllllJqlllJL_

'% I I t I I 25 'E -- I STRUC'_JRAL HOOE CONTROL VANE _SI_ ....

-_0 J.tl tl[ll ll_ 1_ U!l[llll tlllllltl !ILLL2LJ. -2_ 1 JJJ.J._LLLI lJ0 _.0 80 120 0 _0 80 120

_ 20: k9:32 SECONOS SECONDS

I No.l inletI INo.2inletIRB = 7 deg, RC = 9 deg RB = 7 deg, RC = 5 deg

_ i __ _ LO --_ ..... , STALL_RGIN INDEX "-- 70 ]1 f t 111[ I III 0 I III , . _.... _ ,

0 Itl[ ]1111 IIII Iiiiiiii I I II I 1

°" r-i ]FAN DISCHARGEHACHNO. SIGNAL GNAL

(AFT CiNTROL)

/

iitiil[i itll_l:_ ::ii_111 I.i]ll_ll II!ll;llI_[illlll L 0 F_l_'l_ll/t:_ll_ lllrillll: :1, _:.: t_--.--._._-L .... _L_I

0 gO 80 120 0 _.0 SO 120

SECONDS SECONDS

Figure I19. - Effects of sideslip angle on steady-state inlet characteristicswith SMCV= 0 degrees and alpha = 1 degree, flight 2-37.

181

MO (DEG)

0._ kLl I CI_: [_a ,EL _ -_ IIIIllllllllllllllll IltlllFtr 111111111 IIIIlllllll)[l[ll[ll

,,,, _, , 11 ,, .c IIIJ_ll_[_l_l_LLLlITIllll_*ll_lJllrl_JllllllllJl!

(_F) 0 - S,_'V O-C_G)

o k.0 80 !20 0 _.0 80 12o

i_ SECONDS SECONDS20:W]:02

INo.1io,e_I Imo._io,e_IRB = 7 deg, RC = 9 deg RB -- 7 deg, RC = 5 deg

1.0 ...... %T_LI._qGtN INDE.X_ -_ 1.0- STALLMARGINI,NDE:X

0 .... 0.......................................i.........1

_60. 360CORRECTEDMEIGHTFL.O_ GORRECT1EDV_EIG,HTFLO_

WIR .(PPS) " (PPS) -

o._., i i,_ o_ _ No.s_r._.. I I o.s t _A_O_s,:_,_r__c, No.s_r,_. " "

_AN FAN p_ -F -,--

I :0 JJJ ilJi il_Ji ILIJlJlLI Itllllllf ItJilJlit ILIIJlilf O_ {{{IJ[lll ili_lli:llli[lil L_[I_II[I IliJ_ll il?_li,_

0 WO 80 120 0 40 80 120SECONDS SECONDS

Figure 120. Effectsof sideslipangle on steady-stateinlet characteristicswith SMCV = 20 degreesand alpha = 1 degree, flight= 2-37.

1.82

" ..

Ramp

Wi ng

~Looking aft

Cowl

5 deg

INo. 2 i n1e t IM = 0.84, RB = 7 deg, RCSMCV = 20 deg,~]

Pressure above PTI \in shaded areaPressure below PTJin unshaded area

number onPTI - PT I [1 + ( can tau r )- ~

fUse + if contour in shaded areaUse.- if contour in unshaded area

Alpha =

W1RIDL =

lOT =

IBeta =Beta = 0 degJ

Alpha = 1.4 degW1R = 312 ppsIDL = 0.296lOT = 0.056

!Beta = 1.6 degl

1.3 deg313 pps0.3040.055

AlphaW1RIDLlOT

4.0 deg I1.4 deg311 pps

= 0.4200.076

Dynamic total-pressure contours(max IDL scans, filtered to 62.5 Hz, sampled at 360 sps)

Figure 121. - Effects of S~~l1 wake on no. 2 inlet dynamic total-pressure contours duringsideslip, SMCV = 20 degrees and alpha = 1.4 degrees, flight 2-37.

_0 (DEG)

0

i_ ALTIT_JDE INDICATEDAi_,LE OF SIDESLIP

aLr IT'J_I_ E_TA(v FEET) (DEC,) 0-

,_TT _MCV,.oF) 0 C_E[G)

L_ L _.--_J J-U J- _LLU,--U._U-I

0 20 _0 60 80 100 0 20 _,0 60 80 100

_ 20:5].:[15 SECC_NDS " SECONDS

i.o., 'n'o! ' I"o. 'nlo lRB = 7 deg, RC = 9 deg RB= 7 deg, RC = 5 deg

PT-"_ PTO

0.9-L,L,U..I _.LI L.L .L.L; L_.LL L _:L. L L.L-L . i i J J i i _ i i 1

IDI. I_.

0 IJ,_LL.L L_. _ & _ I 0 I I t I I I I I I I I I I I I I I ] I I _ I I I 11 I ] ] I J I ] I I 1

IDt _ _ J '_ IDT

WlR __ WIR

l_O -_J [_ . I I I I I I I _ I t I _ I 162 I I I I I I I 11 i I I I I I I I I I [ I I I I I I I

o._ ---- FAN Dl SCHARGEMACHNO. SIGtW_L J / 0._t FAN DIS_HI_qGIEI'_C_ NO. S l G/'_/..

1 _'T co_o_) P_" c,u,TCOm_OL)

0 20 '_0 €'0 80 I00 0 20 kO 60 gO tOO

SECOt_S S_CONDS

Figure 122. Effects of sideslip angle on steady-state inlet characteristicswith SMCV-- -8 degrees and alpha = 1 degree, flight 2-37.

[84

Pressure above PTI

I I Wingin shaded area No. 2 inletPressure below PTI

in unshaded area Cowl _ I_ _ Ramp,number on, M= 0 85, RB = 7 deg RC = 5 deg

PTI - PTI [l+ [ contour ) _ ] " '- ^ SMCV = -8 deg Looking Aft


Use - if contour in unshaded area

_J _:_:_:i:_:_:_:i:_:_:_:!:!:i

Alpha = 0.8 deg Alpha = 1.0 deg Alpha = 1.0 deg

WIR = 307 pps WIR = 305 pps WIR = 312 pps

IDL = 0.351 IDL = 0.291 IDL = 0.338

IDT = 0.055 IDT = 0.054 IDT = 0.059

Dynamic total-pressure contours

(max IDL scans, filtered to 62.5 Hz, sampled at 360 sps)

Figure 123. - Effectsof SMCVwake on no. 2 inlet dynamictotal-pressurecontours duringsideslip,SMCV = -8 degrees and alpha = 0.9 degrees,flight 2-37.

ALTITUDE i I(K F[ET) l BETA(Or:(;.)

0 Ill|llll, I''lill,I Illll,l,r llllllll Illlltrl Illlll;I50- 2.5 . .

TE4PfEP.ATURF t i I_[L ON[TROLVANE_POSIT10_NT J(TTO-TTSTO) STRUCT W.._E C

COF) O- (:WEG) I

-50- Iltllllll ZlILfI11_ Zlllllll IIlllltt ll,l(lll IllJllll_ "25 _ _!_ {_.0 80 |20 _+0 80 120

SECONDS SECONDS

_21:).7:52

Imo.' io'e lRB = 7 deg, RC = 9 deg RB = 7 deg, RC = 5 deg

PTO PTO

, I III 0 S-- IIIfl IIIIIIiLl I)lllllll Illllllll

..... o0 O"0.2

g I I o,s_'r,o_ I I I 0.2 oIstoRT,o.

0 IIIIIIIIIIIIIIIItllllllllllllllllllltfllllllllllllllllllll{I I iiiiiitlllJflt fll1111111111lllllllllllllllllJllllllfllll_lll j

360 C_tR_CT__IGHTFLO_ 1_60'I , CO.COTEDWEIGHTFLO_ ;

• _.WIR WIR

CPP$) - (PPS) - "

1_0 _llll,llll Illllllll IIIlltll IIIIlllll Ill*tmll _l£1Iltll _60 Illflllll Iltllllll (llllllll {tlffllt tltfl]l] fflllllll0.5

F_ _l_*,_;eA_ NO.s(c.,,_, o.s- , ,FANmsc_ _cHNO.s(c._

O_ IIIIIlll III!llllll Imlllllll I1{111111 IIIJtllll (lllllll 0_ IlZ*lllll iltflrl] i!llllll llfIIIlll flllllll IlmmIItll0 kO 80 120 0 40 80 120

SECONDs SECONDS

Figure 124. -Effects of sideslip angle on steady-state inlet characteristics withSMCV = 0 degrees and alpha = 2.6 degrees, flight 2-37.

186

0- INDI_q'_ A,'_E _ 'ATTACK J

.- HO ALPHA(Dr:6.)

0.5 J-J--LLLJI Illllllll I UJ_.LL I I I _1 II -2 It[

AI_TI TUDE

ALTI'PJDE BL=TA

(K FEET) - (DIEG.) _5-1_ irlllt ,lil,r jiitli ,i!l j0 IlJllllll ililtllll tltlllllf tl[llllll Illllflll

,o_ ,,,,._ "-_----T..... q------r:_-_:-- -1(TTO-TrS_) I- S_UCT_UU.HO0_CONTROLV_E I I

(OF) 0 - (_G) 0 --

._0 f f I I I I I I I I I t I I I I I I ! I / 1111 llZltle:l lillllllll .2_IJ--LL_L_-LL.LL.L_LJ-J_J_J_J-L._JIJ-LI'-ILLJ-._J.

0 20 _,0 60 80 100 0 20 _.0 60 80 100

_, .L._(,ND_< _E_..tlDS

21:12:55

i o., i0,eqRB = 7 deg, RC = 9 deg RB = 7 deg, RC = 5 deg

,,,,,.,: ......I _c,,;)L_t v:,,,'LR;

pT_ PTO

,,"_ . li 4_ L_.LJ-:_ j,_ i i_i, ,i ,[,Jl_;,,

IIOT I I _ IOT -r

" 0 iiiiiiiii l 0 I[ ii• 360- .

! . /- -WIR WIR

- - (PPs) CPPS) "

I_0 I I _ I I I I I I I I I I I I l I [ I I , I t i_0 _llllflll _ I I I I I I I I I I I t [ I l J ] ] i I I I I I I I I : I [ I I I I I

0.5_

__ FAN DISCHARGEHACHNO. SIGNAL(AFT CONTROL) (AFT CONTROL)

iII iI II II_; lllI_IllIlllIll!ll lll''_J'l_l

_I_ 01 t I I I I I I I_011 I I I I I I I14_II I I I I I 116(} I I I' I I I I I_(} I I I I I I I I00 _ I I ' I I I I f 120 _O _O SO I_0

SECONDS SECONDS

Figure 125. - Effects of sideslip angle on steady-state inlet characteristics withSMCV = 20 degrees and alpha = 2.6 degrees, flight 2-37.

187

_- Pressure above PTI@o

co in shaded area I I

Pressure below PTI INo. 2 inlet I Wingin unshaded area I I

,number on,PTI - PTI [l +- L contour ) A ] M = 0.85, RB = 7 deg, RC = 5 deg Cowl RampA

r.. • SMCV = 20 deguse + if contour in shaded area Looking aft


t• .*...,:..,.........:........

_,,_-!...:.!_!

m_-!!

_,"_.i:_:_"/.,,. ,,............ :_.........................._......>.............:..:.:..:.\-_--....-/

I Beta = 0 deg Beta = 1.7 degl I Beta = 3.8 deg I_


WIR = 317 deg WIR = 317 pps WIR = 318 pps

IDL = 0.391 IDL = O.397 IDL = 0.449

IDT = 0.064 IDT = 0.064 IDT = 0.I04



Figure 126. Effectsof SMCVwake on no. 2 inlet dynamic total-pressurecontoursduringsideslip,SMCV = 20 degreesand alpha = 2.6 degrees,flight 2-37.

i.oi-.---ii...............0 _ UJ./_IItltLJ.J./.J.L /.I.U./_-IJ-L/J I .2_tllllltrlllllllllllllllllll

ALTI'/%IDE BETA .(K FEET) (.DEG.)

i ttllfll iiiii111_ fill11111 i -5 I_Lllllll1_Iti_lllllllltlllll

25.... STRUCT'JRALMOOECONTROLVANE _SITI_ ---

+°- ?1CTTO-TrS_)

ATU -- S/4CV 0- -- --(OF) 0- CDEG) ,

-$0. ]etllrll iil11111 tlllllltlll_tleltl I_1111_t Jllllllsl _25_1LL.L.U.LL_ LJ IJ L.L.LZ.L.L.I-LJ_ JJ.Ai I I I I I _J2I I _, JI I t LL.U.L-U-I-LJ"0 _0 80 120 0 k.O 80 120

_."_ 1$: 57

.g[COI,IDSOk:

I.o.1inletl I"o.2+nlotlRB = 7 deg, RC = 9 deg RB = 7 deg, RC = 5 deg

. 1 II ]jlllll]l 1 I, *,r,,ll,, 11_,,, g._ IIIIII11 ][lllj]ll Illll Illl

GIN INDEX STALL t'ARGIN IND(D(

IDL IDt..

++T+ ' '- (PPS) Wl_

" 160- (PPS)

0..5-- FAN D[_E f_ACHNO. $IGt_4. 0.5- FAN D[SC_GE _4ACHNO. S'|Gt,la,L

(AFT CONTROL) I (AFT cOntroL.) !

51_t_t)_,I_S S_.ONDS

Figure 127. - Effects of sideslip angle on steady-state inlet characteristics withSMCg= 0 degrees and alpha = 5.8 degrees, flight 2-57.

189

-II'{)ICATED An;Lf OF ATTACK V

~-"'" ~~ ___ --.r

-L

~I'{)TCATED AI<;~ OF SIDELIP~

I ,1, ,I

-25

AL_(DEG.)

Sl'CV(DEC)

BETA(DEC.)

_._-~.__ .

SECOI'{)S

ALTln.oE

TEMPERATURE(TTO- TTS TO)

-I--~I

I

- " ~'l

- \0 ~oL.LUl.:..llJ.l..U.J.l..U.J.l..U.~.lJ0Ci.J..L'-ll1-L.LU.:..L..O..L:.s.L1o

J.l..U...LC1.JLU.LL.LU.L:.1lJ

20

~IS:00:52

1.0

110

0.550

ALTlTl.OE(K FEET)

050

~TT

(OF)

120so~O

-~..kIGHTFL~

-

]- ,1J'UliFAN D~SCHAAGE ..lcH I-Cl. Slkw.

(AFT CONTROL)

-

_.1.l.1UJ' u_ .

- -- -:I;~lJ~J~j[Pn-tAX-PfMIN]

PTA\IG

O_ULLLLll-l.11J.JL11.lJLLl-LJLLl-LJ..1.LL:.. l Lt LLl .,

560

0.2 _.

1600.5

lOT

WIR(PPS)

SECOI'{)S SEClNlS

Figure 128. - Effects of sideslip'angleon steady-state inlet characteristics withSMCV = 20 degrees and alpha = 5.8 degrees, flight 2-37.

190

SYM TEST p_L_.T/PT 14 ,N..I AI_P_L_ ETA PTO PCNFR-I PCNFR-2 P_NFR-3 PO'IrR-4 I:_tr)UCTION TIME

O 203"/ 6.01 0,87 23059 5.8 0.1 9.'/ 90.04 98.69 82..53 97.06 1/31/'78 2 2 3 18:0:3113 _037 6.02 D.E_ 23015 5.6 1.8 9.6 99.09 99.42 f./_.24 97.35 1/31/78 2 3 2 18:0:51• 203"/ 6.03 0.85 221853 6.2 2.9 9.6 99.13 99.04 6[:1.'/6 9"7.61 1/31/'tB 2. 2 3 18:1:18

.... AIP pressures/PT0 at max IDL - engine l1.040 . i i _ i -

.........................i....................................................i .........................-_..........................-...........................t..........................f ....................................................................................................! ................................................4........................._............................"...........................=:.........................i .......................................................................;::'"": : : . .

........................i.....................................................................................................................................................i........................................................i...........................-- ,.ooo......................"...........:..... ..................................N

o.96o". :i:i.L.....:.:..::::.:.._::]._...i....]..._!.::::&..::ii...........i.:_:i:iLLiiiiiii_ii_if .....L....:._......_.i_ILL...:...:::::::::::::::::::::::::::::::::::::::::::::::::::::::::::iiili i

O. 840 ......................................................._...................._ ..... . .................................................................................................._........................: : s :.. - :- :...........................:............................•............................;.........................._..................................................................................: ...........................,........................... ................_. l " :.

......................................................;...........................'_........................ • .....................................................T...........................'[ .........................T................. T..........................................................._............................_........................._.....................................................i ...........................,............................" .........................." ...........................

o.8oo........................................................i........................_...........].........................................,:......................,;......................_.........................;_............................•: * | . : :" l: : I ....

................................................................................ .:..........................._............................_............................_............................*............................ ...........................-.........................

O.760 ; ! _ ! ! * : :

3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 a 2. 2 2. a 2 2 _ 2 2 2 2 e 2"_l---Probeaaeae21111111J! I000000000099999999998888_'4 3_. I 09 B 76543 e I 098"76543 _ I 098"/ 6543 a I 098 -/6 number

(see figure 104)

AlP pressures/PTO at max IDL - engine 2I.040 _ _ .; i

............................................................................................................•r..........................."...........................f .....................................................-...................................................• ........................ : .........................;..........................,,1...........................,i............................,_........................._ ........................ _,...................... _.....................

,.ooo::..........%'_ ;_:_:_F_.............._f_....._i...........i_:_:J_:_:_:_i-_:_-_-_=_:_:_:-_:_-

iiii iii !-. _'TZ?IIZZf?Z_L..........&LZ22iXl......................j.2ZIZZLZZZ:iZZZZZTZ£

• , . : . : . I .O.800 .........................._............................_............................:............................_.........................'_...................'_......................................................';............ _".............: ; i : : : : "

..........................|............................, ..........................,...........................,...........................,_............................i,............................;,.-_........................, ..........................., ...........................; : ; { : : : : =............................i ...........................;............................;...........................• ..........................._...........................;,..........................._"...........................,i...........................,,7,,..........................

. ! : : [" : " " : • : :

0. 760 : _ ' " _ _ _ ! ' "156666666666666666666666666666666666666663 "_'3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3"3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 386666655555555554444 _ _ 44443333333333_22_=3 4 3 _, I 0 9 8 "/ 6 5 4 3 2. I 0 9 8 'v 6 5 4 3 _ 1.0 9 8 "/ 6 5 4 3 2 I 0 9 8 3' 6

Figure ]29. - IndividualAIP probes, total-pressurerecoveryarray, SMCS vanedeflected20 degrees,Mach 0.85, _-- 6 degrees,flight 2-37.

191

No. inlet

Filtered to 62.5 Hzsampled at 360 sps

No.2 inlet0.10

C0 0.08

+-JL0+-J . 0.06l/l

ICJl"'0

0.lJ't

rc"'0 0.02rc0::: ,

Circumferential distortion

Figure 130.

, ,

Variations in circumferential and radial distortion componentsduring sideslip operation with the SMCS vane deflected+20degrees, Mach 0.85, a= 6 degrees, flight 2-37.

Pres,ure,nsha oabOVearea .... etiI No. 2 inl Wingii

Pressure below PTI

in unshaded area M = 0.86, RB = 7 deg, RC = 5 deg Cowl :_Ramp

,number on, SMCV = 20 degPTI- PTI [l_+ [ contour ) A ]Looking aft



J

":::i::: _ _" IJ _ _ .1 _ 1 _':"I_ _1_ _, I ,t.::"

:i_iii:_:i_-?i!ii_!??!.".o.I. .j I _. j, i:::i:i:

:iii!!_i:::::::_:_ik;. , .,_ V_:j J::::::::::::::::::::::::

_1 :" ":':".._ :' 1_'""1':':":'1:: : :: :" ":'::::::::: : :" ':':::::."": : : ""

11_ '-" ,-.-....... _ .:.:.:....... t- .:.:: ....... _ 4: •

Alpha = 5.8 de 9 Alpha = 5.6 deg Alpha ; 6.0 deg

WIR = 346 pps WIR = 349 pps WIR = 348 ppsIDL = 0.506 IDL = 0.688 IDL = 0.547

IDT = 0.091 IDT = 0.185 IDT = 0.099



Figure 131. Effectsof SMCVwake on no. 2 inlet dynamictotal-pressurecontoursduringsideslip,SMCV = 20 degreesand alpha = 5 8 degrees flight 2-37° y •

Pressure above PTI Wingin shaded area A

Pressure below PTI T

in unshaded area M = 0.86, RB = 7 deg, RC = 9 deg Ramp _ _- _ Cowl

,number on SMCV = 20 degPTI - PTI [l± t contour ) _ ] Looking aft


Use _ if contour in unshaded area

Beta = 0 deg [Beta = 1.8 deg I iBeta = 3.6 deg[



IDL = 0.580 IDL = 0.607 IDL = 0.706

IDT = 0.099 IDT = 0.120 IDT = O.163

Dynamic total-pressure contours(max IDL scans, filtered to 62.5 Hz, sampled at 360 sps)

Figure 132. Effects of SMCVwake on no. 1 inlet dynamic total-pressure contours during

sideslip, SMCV = 20 degrees and alpha = 5.8 degrees, flight 2-37.

For example,local recoverieson a probe-by-probebasis are illustratedin

figure129. Probeswithin a particularrake are connectedby straightlinesand are arranged in order from hub to tip. This particularplot uses sideslipas the independentparameter. SMCS vane wake ingestionis evidentin theinboardinlet at 2 degreesof sideslipand is becoming evidentin the outboardinlet as sideslipis increasedto 3 degrees. Distortioncharacteristicsresult-

-- Lng from scans that produced the maximtmvalue of stall-marginindex are illu-stratedin figure 130. Combinationsof circumferentialand radial distortioncomponentscomputedat 360 samplesper second are shown for selectedaircraftattitudeswith the SMCS vanes deflected20 degrees. Maximumvalues of stall-

margin index are identifiedand comparedagainst inlet design goals (IDL= 1.0).Correspondingtotal-pressurecontoursare illustratedin figures131 and 132.

Maneuverswith the _CS vane deflected20 degrees (figure133) illustrate

the sensitivityof vortex ingestionto angle of attack. At zero sideslip,vortex ingestionis evidentin the no. 2 inlet betweenangles of attackof7 and 8 degrees,consistentwith the maneuver envelopespresentedearlier.(Seefigure 112.) Similarangle-of-attackexcursionswith the vanes held ina neutral positionare shown in figures134 through136 for comparison.

Steady-stateinlet characteristicsduring sideslipmaneuverswith the SMCSvanes deflected13 degreesare shown in figure137. Total-pressurerecoverydecreasesapproximately5 percentwhen sideslipangle increasesfrom 1.7 to

4.0 degrees. Recovery lossesare accompaniedby significantincreasesindistortionlevels. Total-pressurecontours representingmaximtm dynamicvalues of stall-marginindex as a functionof sideslipangle are shown in fig-ure 138, and a well-developed,low-pressureregion is evidentin the hub.Stall-marginindex increasesfrom 0.5 to 0.75 as sideslipis increasedfrom1.7 to 4.0 degrees.

EngineThrottleTransients. Engine throttletransientswere conductedatselectedconditionsto demonstrateengine stall-margincapabilitybeyond that

requiredby the inlet during SMCS vane wake ingestion. Initialtests were con-ducted during flight 2-19 and 2-33. Qualitativedata and operatingconditionsare summarizedin figuresIII and 112.

During flight 2-38, maximumrate throttletransientsbetween IDLE and-- IhFFEPJ_DIATEpower-leversettingswere conductedon the no. 2 engine during

sideslipmaneuversat differentaltitudeswith the SMCS vanes deflected20 and-8 degrees. Steady-stateresultsat nominalangles of attack between 1 and5 degreeswith sideslipout to 4 degreesare shown in figures139 through141.

195

V'II

RC

in 1et I

F~dIS~~~L(AFT CC,nROL)

I I

I J." '"

, No. 2

O...L.Lu.l.Ju.l.J..l.1..1..l.1..1..l.1..1.Lll.l..Lllll.llLLll

0.2,------,----,------.---, ,-__-,

RB = 7 deg,:.o~

pn.E I 1

PTO

I~L

o•8 ..LI-LJ...LLJ...LU-LLJ...LL..LLJ...LLLLl.LLl.LL.LJ..l..L.

:.0 .,..-----r-----,-----,-

o ..LU..l.1..1.l.l J.l.l..J..LUlllJlllJLl..LLi..LLllLllll.l.ll.l..LllllllllJ..l.l"JJ1.l..l..l..lll..

3&0 F====1'==--==r==JORRECTEO WEIGH; ,~C\, -~1

wlR Ll(PPS)

1&0~ " J0.5

.20SECC1'VS

RECO'/ERY

STALL ",I1'VEX

40 80

I No. in 1et

RB = 7 deg, RC = 9 deg

U",L[uT

LL'.u

L

LL

o\ol1JL.LLJ...LLl..LCLU..l.1..1..LJ...L.lJ

':'LTlTLlJE

JJUl' UlJJJJ " .1.Lllll;LJ

FAN DISC>'AAGE .....CH ,t). SIGNAL

(AFT CONTROL)

=----- I Av

JJJJl

I.

1rmI-- D- ~'"'~~ ,

~ , _111'1!,

50

.J.LTIT1JDE(if. r~ET)

?LA(CE~)

1.0

~PrJ

0.3\.0

IOL

1Q.2

lOT

1HIt)

'~lP

(PP5)

1&0O. ;

FAll ~.Q

200

Figure 133. - Effects of angle-of-attack maneuvers on steady-state inletcharacteristics with SMC vane at 20 degrees, flight 2-42.

196

L CDEO /0

O, _111111!11 iiiiiiiii illlflTll |Tlllltll [llltllll IIit1,111 _ I11 SlI _1!111_

• (K FEET) BETA(DEG)

0 iltllllll Iffllllll IlJtllll! IJllllll I_1111[11111_I[11_11111 Illl[tlt! IlrlrtltllL_0 120 _ _'0 80 120SECONDS SECONDS

RB -- 7 deg, RC -- 9 deg RB = 7 deg, RC -- 5 degL.O ,

EE RECCVERY

p'r_l p'r_PTO PT0

0, 8 Irllllllltlllllll I_111111il I11 r I I IlllllllIl LI II III1 Ill It It i1 0, 8 lilJ_lll rlltlllll Illlfllll

0 Illllll flllll f fill IIIII III II I111111111t II III II II IIIIII1?111 I| Q llJlt IIt IlllIIlll I f III III I III II1111fll[I I11111111111111111 ll|

PTAVG

0 i Ill 0 |llltllll) IIIIllllllllll f IIIIll _1111 flflllljljlflllt II IIIIII/

C_RECT_D WEIGHT FLOW , CORRECTEDWEIGHT FLOW

WIR W1R_PPS) (PPS)

0.5 0._ _ _FAN DISCH/_GE ,-_, NO. STGN_L(_,FTC0NTI_(_,.) (AFT CONTI_OL)

0 Ill|Ill !|lt|f tl|llll! !1|11111 I|11111|1 |1111|1 0 _ I IllJl[I IIIIIIIIJ till|Ill

" IPl..A . PL_

(DEG) (D_G)

0 oIllllrtll IIIItllll Illlltlll Illllll Jllllllll ItllltlllJ 0 _l'_lllll:lllllllllllll_o_.0 0 %20SECONDS sEcOflDS

Figure 134. - Effectsof an angle-of-attackmaneuver on steady-stateinletcharacteristicswith no. 1 engine at intermediate,flight 2-42,

SMCS vane deflectionangle = 0 degrees.

197

IRECOVERY

STALL /'o4ARGIN It-VEx

20

INo.2 inletl

7 deg, RC = 5 deg

I~DICAiE:J .G.l'GL:: 'JF ~TT~CK

..r---..~ ~ lr---

u' .,

2 DISTORTlCJ'.l[PTKAX-PT><INj" PTAVG

1-

, ,

~QRRECTE:> ~ IGHT FLl

5 FAA! OISCHARGE KOoCH NJ. StGN4L(AFT CONTROL)

. !

0' ,

0 F'J1oIEK LEVER fol-GLE

I0

..

1-J

5.,----,------;,..---:-:-::-0------.---,----,{'rCArED""{;L~ ':F ~~1-2Er_SL-I~P_ _+_~__

V

20

II

SECC/-DS

ol60

oliJ.1.l.lll.1.tl.u..J.l..lll.1.Li.J.Jl.UJl..U.Li.J.JLL!.ll..U.J..!..ull.1.UUlll.l.ll.U.l.-'-L

O.

160

O.

lOT

RB =1.0

PTlPT~

0.8

1.0

lDL

eEiA(O~G) k_

WIR(PPS)

PLA(OEG)

GOSECOf,()S

ALTITl.OE

20

~No. inletl

7 deg, RC = 9 degRB~ RECOVERY

~

STALL "1ARGIN [N:lEX

U I~

llll.l.llllu.lllU

IJ .-..-""

·LJ OrSTCRTION

rTKAX-pT><IN]P~iVG

tJI I

C:)RRECTED wE IGHT FLOW

"

F"'N OISCJ-'"Cl.RGE MACH ~. SIGf'.4A.L(AFT CeNTROL)

.~ .f\.A

7.1

1.0

~" ...----r--.....-PQW:-::::ER-L--:~-E"".-t_>-G,...L-E-.--- .....---,

oloO

7

n.2

160

7.5

lOT

ICL

~l~

rpP<j)

1<0 :.r:[un[~_==::-RU'eER

~ u ' ULUllLUlJUlJ..Ll..l..Ll..l..llJ..l..llJ..aJ

-,--c==c:-----r--,------,

'<.T1n;DE'OEfr ... <'"<::::T:

° ll.U llUJ 1 LLlll_l.lllllll.l.JJ.ULLi.J.J..Ll..l..Ll..l.ili.L1.l.l.ll.W

~11:l2:lO

Figure 135. - Effects of an angle-of-attack maneuver on steady-stateinlet characteristics with no. 2 engine at intermediate,

flight 2-42, SMCS vane deflection angle = 0 degrees.

198

\.0 1===+==::[::3""~C~H ~'U'~eE3:'=+===F==l

fAD ').;

ALTlruoE(K FEET)

1+--17:n:2&

L,,~

-< col,> "C" 1---1--,f---+---"-----+-- -, !

I~L

,0

deg

in 1etl

RC = 97 deg,RB

~ J- IRECC"C~Y

lu ,I

"~'l3±JLA~ ~.• o "- "'"

DISTORTION[PT'-'A)(-~""Wll

PT;'VG j

a

IC:JRRE>::::~ rWETSI-" .:'L1W

a ~.

\ FAN 'JISCI"¢.RGE "'ACH f.IJ. SIGNAL(AFT CCt,TROL)

!~

'J "

POtiER L::~E~ ..:.tGLE

If\

",.

36

~ 6a.

O.B\.0

200

a0.2

\.0

IDL

lOT

~PTa

WIR~PPS)

FAN ff

SECCtJ)S

Figure 136. - Effects of an angle-of-attack maneuver on steady-stateinlet characteristics with no. 2 engine at intermediate,

flight 2-42, SMCS vane deflection angle = 0 degrees.

199

Figure 137. Effectsof sideslipangle on steady-stateinlet characteristicswith SMCV = 13 degreesand alpha = 3.3 degrees, flight 2-38.

200

Pressure above PT I I I _Wing

in shaded area No. 2 inlet

Pressure below PTI

in unshaded area M = 0.85, RB = 7 deg, RC = 5 deg Cowl -_4_>- Ramp

,number on SMCV = 13 degPTI - PTI [I+_t contour ) A ] Looking aft

i

Use + if contour in s'haded area


I lBeta = 0 deg Beta = 1.7 deg Beta = 4.0 deg



IDL = 0.514 IDL = 0.504 IDL = 0.750

IDT = 0.090 IDT = 0.099 IDT = 0.163

Dynamic total-pressure contours i


Figure 138. -Effects of SMCVwake on no. 2 inlet dynamictotal-pressurecontours during•o'_ sides-lip, _ICv'= 13 degrees and alpha = 3.3 degrees, flight 2-38.

&0

SECCNlS

20

-aDlCATfD At-&L~ OF ATTACK

"- ~ V'-- '- ~-v-v

_1-ll.LLJ..ll.I.JL.U.1.LU.LLJ..WJCLl..l.LU.LLJ..ULWJ.LU.LUULWJl.U..1.LU.ll.LJ

ALPHA(OEG.)

BETA(OEG.)

&0SECCN>S

o 20

~20:00;4b

[1MAC.H N..f"8ER

-- -_.

-ALTlruoE

-

LLLU

0.1\0

1.0

MO

:"I.TlflKJf(r FEET)

RB

in 1etl7 deg, RC = 9 deg RB

INo. 2 in 1e tl7 deg, RC 5 deg

&0

STALL MARGIN I/{)f)(

20

WJ"(PPS)

IOL

PTIPTe

lOT

PLA(DEG.)

&04020

:rRECOVERY r

'\l/\-vv -v

STALL MARG IN IIllEX

) 1"4 V -"'y- rv-

- U

r--

JIV'I n..,.~ '\OIJ:I~-~ [PTl<AX-PTMIN]

PTVG

-I--- or

~RECTED 11GKT FLJ-

,,"~LJ,~_ JJUJ1..l.l-l-

FAN DISCt-'ARGE r-oACH tD. SIGNAL(AFT CONT"OL)

- Iv-.

--.---

PCWER LEVER AAGLE

.r--t-------.

0o

1.0

o0.2

o1&0

0.81.0

IOL

lOT

~PTO

WI'(PPSI

1&00.1

FAN ¥

0200

PLA(;lEG. )

SE(CIllS SECCN>S

Fib~re 139. - Steady-state lnlet characteristics during no. 2 enginetransients, sMcv ~ 20 degrees, alpha = 5.5 degrees,and beta = +4 degrees, flight 2-38.

202

S-r----,-----r----,-----,----,

", ]~--'~t[l~::~>]""""g,~','~,blJ

a 20 40 60 80 100

I SECOI'()S

f-----21:SI:S8

INo. lin I e tl

.oLPHA(DEG.)

BETA(DEG.)

8!>Il!CATED Al'GLE OF ATIACK

a-2- ."

I1>IliCATED Al'GLE OF SIDESLIP

I20 80

SECOI'()S

INo.2 inletl

lOa

o ' . , "-Ll.LU..LU..LU..LU..u..L.u..L.u..L.w

200_

l:- """Ell LEYOl AN::LE

- r-r-'-Li~- - --\.-, r'-

_ ll1.-t ' 'I

'~IR

(PPS)

FAN ~p

PeA(DEG. )

20

SEem,DS

60 80 lOa

IDT

WIR(PPS)

F/>I< -¥

PLA(DEG.)

SECCKJ5

Fi~lre 140. - Steady-state inlet characteristics during no. 2 enginetransients, SMCV = 20 degrees, alpha = 3 degrees, andbeta = +4 degrees, flight 2-38.

203

240160

I--l rulm~DiATEIDLE

H""tE.,;: Lt.VF.R MoGLE

"\. .J rvREJVERY

.. ~.

-

-5. ~. 160 240

SECDt<lS

20 deg -8 deg •INo. 2 in 1etl

RB 7 deg, RC =-5 deg1.0

PLA(oEG.)

240

BETA(OEG.) 0

240

ISMCV

160

SECQr-()S

SEcoms80

FAN DISCtAAGE r-tACH N>. SIGNAL(pET CONTROL)

- I.

--t Pl"MER LEVER A1'C.LE

-lrJI~~\LJlfL-.-.-lJU

_Wlll.llll ,

AL_(oEG.)

1.0- .._- u_ r----.---~--..._--__,___--__,

o'00

o 0

o

l---21: 40: 48

'No. 1 in 1etlRB 7 deg, RC = 9 deg

1.0

~ PTIPTO ?TO

0.'1.0

101. 101.

00.2

lOT lOT

PLACCEG.)

MO

WI"(PPS)

AL TIT\X:lE(<FEET)

Figure 141. Steady-state inlet characteristics during no. 2 enginetransients, ~1CV = 20 degrees and -8 degrees,alpha = 1 degree with positive sideslip, flight 2-38.

204

SU_V_RY

B-I experiencefrom air-inductionsystem flight tests with vorticesgeneratedby the SMCS vanes is summarizedwith the followingobservations:

_ (i) Agreementbetweenflight-testand wind-tunneltest resultswasgenerallygood. Small differenceswere noted in maneuvers (combinationsofangles of attack and sideslip)that resultedin wake ingestion.

(2) Inlet total-pressurerecoveryand distortioncharacteristicscan beadverselyaffectedby wake ingestion,and considerationof these factorsplaysan importantrole in establishinggood inlet/enginecompatibility.

(3) Tests were conductedwith the vanes deflectedand held at a staticconditionand also with the vanes driven plus and minus full-scaledeflectionat a frequencyof 1 Hz. Tests with oscillatingvanes verified SMCS vanewake/vortexingestionin flight. Tests with staticvane deflectionsproducedresuitsconsistentwith oscillatingvanes and were less objectionabletothe crew.

(4) SMCS operationhas not resultedin any flightrestrictions. Dynamicvalues of stall-marginindex remainedwell within allocationsfor all conditionsinvestigated,and there were no flight incidentsrelated to the conductof theSMCS tests.

(5) Effectson total-pressurerecoveryand engine-facedistortionweregenerallyrestrictedto windward sideslipoperationcombinedwith leading-edge-upvane deflections. The 13-degreevane deflectionangle produced effectssimilarto the 20-degreedeflectionangle on inlet performance.

(6) Sideslipangles resultingin wake ingestionare smallerduring opera-tion at higher anglesof attack due to the increasedoutwash from the fuselageforebody. Also, the range of sideslipangles resultingin wake ingestiondecreasesas angle of attack increases. No wake ingestionwas encounteredabove 7.5 degreesangle of attack.

S_Y OF SMCS FLIGHT TEST.RESULTS

" The objectivesof the SMCS flight-testprogramwere fourfold: (i) obtainspecificdynamicresponse data to validateanalyticalmodels of the aircraftand controlsystems, (2) determinedetail SMCS performancecharacteristics,(3) determineimpactof SMCS on handlingqualities,and (4) determineimpactof SMCS on operationalcapabilitiesof the aircraft.

205

l'ouraircraftwere built under the B-I contract;A/C-I and A/C-2 wereused for detailedengineeringand developmenttests while A/C-3 and A/C-4were used for operationaltests. As of 16 March 1979, B-I SMCS-relatedtesttime was as Follows:

A/C-I 13.6 hr

A/C-2 3.3

A/C-3 116.5

A/C-4 0.3

Total 133.7 hr

A considerableamount of the flight-testdata has alreadybeen discussedin associationwith the topicsof previousparagraphsof this report. It istileintentof this sectionto present flight-testdata not touchedupon by thespecifictopics coveredearlier.

_S PERFORMANCE IN TIME-HISTORY-DATA FORMAT

Accelerationtime historiesare perhaps the most dramaticmeans of demon-stratingthe ride qualityproblem solvedby the SMCS. In the vertical axis,the syn_etricfirst-fuselagemode at approximately3 Hz is the big motionproducerat the nose of the aircraft due to turbulence. In the lateralaxis,tileantisymmetric(lateral)first-fuselagemode at approximately5 Hz producesthe largestresponseto turbulence. Figure 142 illustratesthese motionswith typicaltime-historyplots of verticaland lateralaccelerationat thenose of tileaircraftat the SMCS vane location. The aircraftwas flyingthrough low-altitudeturbulenceat M = 0.75with the SMCS off at the timethat these recordswere taken.

Figure143 shows data similarto that of figure142; only during thisflighttest, the SMCS was operatedto determineits effectiveness. The datasl_cn were recordedon a-flightat M = 0.70 where the B-I was flying ataltitudesof 305 to 610 meters (i,000to 2,000 feet) above the terrainin thelocal EdwardsAir Force Base area. Considerablelight to moderateturbulencewas presentnearly continuously. The SMCS was turned on with both the vertical -

and lateralcockpitgains setat 1.5. To demonstratecomparativeaircraftper-formancewith and without SMCS operating,severaltime periodswith the SMCSon and off were recorded. Figure 143 is typicalof these data. The pilotstationverticalaccelerometerwas not operativeduring this flight,so thereadingof the vertical accelerometeron the radome is shown.

2O0

A2008

Vertical acce] 3 Hz first fuselage symmetric

at FS 572 (225) (vertical) mode

+0.129 g

•1.0 g---0.129 g

_k

+0. 132 g

0 g---0.132 g

-

A20I ILateral acce] 1 5 Hz first fuselage antisymmetric (lateral) mode i

i sec

SCAS on, SMCS off

M = 0.75, alt = 152.4 m (500 ft)

Figure 142. - Typicaldynamicresponse near crew station dueo

to turbulenceduring low-altitude,high-speedflight.

F,;COo

SMCS off SMCS on

Radome uE "- _ Radome E ._

vertical ._ vertical _ _ [velocity _ velocity "_._A2027 r-,_

Fs56(22)_ _

Pilot- us Pilot- u m

stationlateral _ station m c_.• lateral °i_

velocity _JL _ +_A2011

FS 747 (296) Right ° I _ ^ . . ^ Avane F-,

dell ,_i

Left °

vane

defl

Figure 143. SMCS performancein turbulenceM = 0.70, _t = 305 m (I,000ft) AGLA= 65°•

!

Verticaland lateralmotion at the front end of the aircraftwith the

SMCS off are shown by the first two time historiesof figure 143. As indi-cated, the primarymotion in the vertical axis was the first fuselagebendingmode at approximately3 Hz. The lateralmotion was composedof whole-vehiclemotion near 1 Hz and the first-fuselageside-bendingmode motion of approxi-

_ nmtely 5 Ilzsuperimposed. When the SMCS was operated,as shown in the nexttwo plots, considerableattenuationof the 3 Hz motion was achieved. Verylittlemotion of the aircraftat lower frequenciesappearsto be present. Theeffectof the SMCS on the lateralaxis motion was not as dramaticas on the

verticalmotion, but the 5 Hz motion was partiallysuppressed. The whole-vehicle lateralmotion was not attenuated. It is to be recalled,however,that the SMCS is designedto attenuatestructuralmode responsewithoutadverselyaffectingw|mle-vehiclemotion (handlingqualities). The last twoplots in figure143 show the SMCS right and left vane motion during the timethat the SMCS was operating. As shown,the maximumvane deflectionsseldomexceeded 16 degrees,whereas, +20 degreeswere available. Both the 3 Hzvertical and 5 Hz lateralstructuralmotion can be seen to drive the vane

deflections;the largestcomponentis due to the verticalmotion.

Anothermeasure of the SMCS effectivenessis the amount of structural

dampingthe system is able to provideto the key fuselageresponsemodes. Theupper left-handplot in figure 144 shows that it was possible to excite thes_unetricfirst-fuselagebendingmode with a sharp horizontal-tailinput pulse.;:tomthe time historyof the vertical load factor at the SMCS vane location(afterthe horizontal-tailpulse was removed),it was possible to extractthestructuralmode dampingratio, _. Figure 144 shows a plot of _ obtained inthismanner versus SMCS verticalgain setting _as set in the cockpit). A

typicnlnominalgain settingof 1.5 is indicated. The upper right_handplotshows the time-historyresponseof the normal load factor with the verticalgain at this setting. Figure 145 demonstratesthe vertical SMCS performanceover a wide range of vehicleweights.

Attemptswere made to excite the antisymmetricside-bendingmodes withsharp lower rudderpulses and to extract structuralmode dampingratios.

- This did not prove to be a successfultechnique. A differentexcitationtechniquewas used. As has been mentionedin an earliersection,A/C-Iand A/C-2 of the test B-I aircrafthave syst_nsinstalledallowingthe SMCSvanes to be oscillatedat variousamplitudes (A) across a range of frequenciesup to i0 Hz. This capabilitywas used in an attemptto extract structuralmode dampingfor lateral side-bendingmodes. The techniquewas to selectaresonantfrequencyand amplitudesufficientlylarge so asto providetransientdata when the forcingmotion was cut off. The data of figure 146 shows thatit was possibleto excite the 5 Hz first-fuselagelateralbending mode rathercleanly. .Exceptfor SMCS off response,the decay responsecould not be usedto extractthe structuralmode-dampingratio because of high-frequencymode

209

SMCS off SMCS on, vert gain 1.5

! I t I, I ! _ I

0 ] 2 3 0 l 2 3

nz Time- sec nz /'_ Time- sec

FS 572 FS 572

(225) (225)

n

z nzFS 2649 .._ FS 2649

(IO43) (I043)

LH 6 cv J_

RH _ cv /__

j-0.15

Structural I

mode O.lOtime history

damping 0.05ratio plots

0 , , , , I . • _ , • I0 l 2

SMCS vert gain setting

A = 650 , M : 0.85, alt = 914 m (3000 ft)

Figure 144. - First fuselagesymmetricstructuralmode dampingfrom horizontaltail pulse excitations.

oM = 0.85, alt = 18i9 m(6000 ft), wt = 131 544 kg (290 000 Ib)

, 0 M = 0.85, alt = 914 m (3000 ft), wt = I16,575 kg (257 000 Ib)

O M = 0.85, alt = 914 m (3000 ft), wt = 123 969 kg (273 300 Ib)

OM = 0.85, alt = 914 m (3000 ft), wt = 144 698 kg (319 000 Ib)

M = 0.85, alt = 152 m (500 ft), wt = 127 008 kg (280 000 Ib)

A= 65°

Structural mode frequency,3 Hz

0.20

I

IO. 15 f

structural Imode 0.l0 I

damping Iratio il

Nominal

0.05 Settingg

II

" 0 I I I I I0 0.5 1.0 1.5 2.0 2.5

SMCS vertical gain setting

Figure 145. Effect of SMCS verticalgain settingon first fuselagesymmetricstructuralmode dampingat variousaircraftweights.

211

Lateral acceleration Exciter --

at SMCS vane FS 572 (225) off SCAS off, SMCS offI

'±0.J8 II

A=150 I

f=5 Hz I SCAS on SMCS off -Ii

+-0.20

A=150 i

I SCAS on, SMCSiI Lat gain 1.0

A=I50 II SCAS on, SMCSII Lat gain 1.5

+0.15 gl I

A=200 II

I SCA_Son, SM_C-SI! Lat gain 2.75

+-O. 18 g_I

A=300 I _ l sec-_

A= 65°, M = 0.85, Alt. = 762 m (2500 ft)

Figure 146. First fuselageantisymnetricstructuralmode dampingfrom forced SMCS vane oscillations.

212

contamination. The data do indicate that the SNCSis effective in eliminatingstructural response motions at the lateral system gains tested.

SMCS PERFORMANCEIN PSD-DATAFORMAT

in additionto time-historydata, anotherconventionalway of lookingat

ride qualityperformanceand the effectof the SMCS on performanceis in theform of PSD plots of load factorsat the pilot station. Figures147, 148,and 149 show typicaldata of this type for the vertical and lateral loadfactorswith SMCS off and on. Becauseof other test requirements,none of

the B-I flight-testaircrafthad a gust boom installed. Approximateverticaland lateralgust intensitieswere estimatedusing nose-boomangles of attackand sideslipangles in order to providenormalizingfactors for the PSD data.Gust intensitiesfor the data shown were estimatedat 1.22 to 1.52 meters per

second RMS (4 to 5 feet per secondRMS).

l:igure147 shows that the approximately3 Hz first-fuselageverticalbendingmode, previouslyshown in the time-historyplots of figures 142 and143, ks the main contributorto the verticalmotion at the pilot station. Asin figure143, the data of figure147 demonstratethat the SMCS is veryeffectivein reducing the pilot vertical load factor.

PSD plots of the pilot stationlateralload factorare shown in figures148and 149 for the SMCS off and on conditions. Insteadof a consistentsingle-

peak responseas in the vertical case, the lateralresponse exhibitstwodifferenttypes of responses,dependingon fuel loading. Figure 148 illustratesa single-peakresponse as seen in the time-historyplot of figure 142. Fig-gure 149 shows the other common responsewith two peaks between 4.5 to 6 Hz.In this latter instance,the time,historydata appearmore random than thelateralaccelerationtrace of figure 142. These data were taken with the SMCSforwardsensor package in its originallocation. The SMCS is seen to signifi-cantly reducethe main responsepeaks but tends to excite some of the higherfrequencymodes in the immediatevicinity. It was this couplingthat led to

. the sensor relocation study discussed in detail earlier.

-- SMCS AND HANDLINGQUALITIES

One of the design goals for the SMCS was not to interferesignificantlywith basic handlingqualities. The impactof SMCS operationon the B-I handlingqualitieswas determinedfrom horizontal-tailand rudder-doublettransientresponses. The resultsof these tests are shown in figure 150. The wingswere at the 65-degreesweep position,Mach 0.85, and altitude1524 meters(5000 feet). During these tests, the CG locationwas varied.

213

SCAS only

SCAS + SMCS (cockpit gain set 1.5)

--4

0.00025

0.0025

0.00020

0.0020 _nz

0.00015

g2 0.0015' g2

(m-_sec)2 (ftlsec) 2tad/see

rad/sec 0.000100.0010

0.0005 0.00005

0 - 0 _0 lO 20 30 40 50 60

Frequency - rad/sec

I, I I I I I I I I I I

0 l 2 3 4 5 6 7 8 9 lO

Frequency - Hz

Wt = 125 194 kg (276 O00 Ib), CG at .35 _w

= 65°, M = 0.85, alt _152 m (500 ft)AGL

Figure 147. - Vertical SMCS performance in turbulence as shown by power spectraldensity of vertical load factor at pilot station, FS 747 (294).

214

SCAS only.... SCAS + SMCS (cockpit gain set 1.5)

0.40 xlO--3

"- --L

0.35 xlO

0.35-

0.30_ny

o.30- -2Jg2

(m/sec)2 O.25rad/sec g2

0.25 -(ft/sec)2rad/sec

0.20

0.20 -

0.150.15-

0.100. I0 -

i¢'_\

I \€'\

0.050.05 - /\\ ¢/ _., , \\

O- 0 _ -_ _-. -_-0 10 20 30 40 50 60

f

Frequency _ rad/secI I I I I I I I I I I

-. 0 1 2 3 4 5 6 7 8 9 10

Frequency-, Hz

Wt = 125 194 kg (276 000 Ib), CG at .35 _-w

A= 65°, M = 0.85,alt_ 152 m (500 ft)AGI

Figure 148. - Lateral SMCS performance in turbulence as shown by power spectraldensity of lateral load factor at pilot station,

FS 747 (294), single-peak response.

215

SCAS only

SCAS + SMCS (cockpit gain set 1.5)

0.12

0.12 m xl0 -4

xl0 -3

0.I0

0. I0

0.080.08

g2

(m/sec)2 g2 !I 13

rad/sec (ft/sec)2 II

rad/sec I1t_

0.04 0.04 \/ _

/ 1

O. 02 O. 02 I I r_\ / "x

../ / I,","x

0 _ OJ0 l0 20 30 40 50 60

Frequency- rad/sec

i I I I I 1 I I I I I0 1 2 3 4 5 6 7 8 9 I0

Frequency - Hz

Wt = 128 369 kg (283 000 Ib), C.G. at 0.46 Ew

A= 65° ' M= 0.70, alt _305 m (1000 ft)AGL

Figure 149. - Lateral SMCS performance in turbulence as shown by power spectraldensity of lateral load factor at pilot station,

S 747 (294), double-peak response.

216

25" °Dutch roll mode 2.0

frequency - 1.5rad/sec

l.O

.5

0 I I I I I0.2 0.3 0.4 0.5 0.6

C.G - MAC

5-Q Q

Short-period mode --_--B

frequency - 3

rad/sec

21 _SCAS on, SMCS off--O--SCAS on, SMCS on

!

0 I I I I I0.2 0.3 0.4 0.5 0.6

C.G.-MAC

Wt = I17 482 kg(259 0001bs) to 128 822 kg(284 O001bs)

A = 65 °, M = 0.85, alt = 1524 m (5000 ft)

Figure 150. - SMCS impacton short-periodand dutch roll frequencies.

217

In the longitudinalmode, short-periodmode frequencyand dampingratiowere extractedfrom the transientfollowingthe pitch doublets. A slightreductionin the short-periodmode frequencywas observedwhen operatingwiththe SMCS on (figure150) for all CG positions. No significantchange in theshort-periodmode damping ratio was observed.

In the lateral-directionalmode, the Dutch-rollmode frequencyand damping --ratio were extractedfrom the transientfollowingrudder doublets. Resultsobtainedfor the lateral-directionalhandlingqualitieswere similarto thosefor the longitudinalhandlingqualities. The Dutch-rollfrequencyis slightlyreducedby the SMCS at all CG _ositions. (See figure150.) Again, no signifi-cant effectof SMCS on Dutch-rolldampingcould be detected.

In order to determinewhether SMCS would interferewith aircraftmaneuver-_ng,a roller-coastermaneuverwas executed with the SMCS off and on. Evalua-

tions of recordeddata of the SMCS vane deflectionsshowedno significantmotion;maxin_mvane deflectionsrecordedwere less than 2 degrees. This wasdeterminedto have a negligibleimpacton the maneuveringof the aircraft.Subsequentto the previouslydescribedtests, pilots have reporteda slightincreasein stick force required in terrain followingwith the SMCS on overthat with the SMCS off.

SMCS HICH-GAINTESTS

B-I A/C-I and A/C-2 were used in the high-gaintests to demonstratesyst_gain marginsover the expectednominal gains. It had been initiallyplannedtodo all of the testingof the relocatedforwardSMCS sensor packageon A/C-Ionly, includingthe high-gaintests under discussion. However, after only afew of the tests associatedwith the relocatedforwardSMCS sensorpackagehad been completed,A/C-I went into layup for modifications. In order tocontinuewith testing,the forwardSMCS sensor packagewas relocatedonA/C-2.

A summaryof the high-gaintests on A/C-I and A/C-2 is presentedintable X. At nominalgains, there appearsto be no differencein the SMCS

performanceon A/C,1 or A/C-2. At high gains, however,A/C-2 wan limitedbya 35 Hz limit cycle which was not evident in similarA/C-I data. These data _-show that A/C-I and A/C-2 have nearly a factor of two gain marginsoverexpectednominalsettingsfor the ride qualitydesign point of Mach 0.85 at

low altitudes. It is at Mach 0.55 at low altitudeson A/C-2 in a lightweightconfigurationthat the 35 Hz limit cycle preventsobtainingthe expectednominalgains. From the lightweightconfigurationdata of flight1-41 obtainedat Mach 0.85 at high altitudewhere the dynamic pressure schedulegain in the

218

• o i I

TABLE X. - SMCS HIGH-GAINTEST_ SUMMARYi

WingHeavy weight Light weight

sweep M = .85,low alt M = .85 low alt(degrees)

65 Vertical Lateral Flt Vertical Lateral Flt

gain gain no. gain gain no.

1.9 2.2 1.9 2.2

expected expected expected expectednominal nominal nominal nominal

4.8 3.9 2-23 4.5 6.0 1-48

35 Hz limit cycle no 35 Hz3.0 6.0 a2-25

35 Hz evident

Heavy weight Light weightM = .55, low alt M = .55, low alt

55 Vertical Lateral Flt Vertical Lateral Fltgain gain no. gain gain no.

3.0 3.7 3.0 3.7

expected expected expected expectednominal nominal nominal nominal

4.5 4.7 2-23 3.0 3.1 2-23

35 Hz limit cycle 35 Hz limit cycle

S.0 5.5 bl-50no 35 Hz

ia,Repeat of 1-48 runsbRepeatof 2-23 runs

SMCS is about the same as at the low-altitudemach @.55 flight condition,itwas expectedthat a verticalgain of six would be obtained. However, as thetable shows,this value was not obtained.

These inconsistenciesin A/C-I and A/C-2 data led to repeatingwith A/C-2a set of previouslyrun A/C-I tests and vice-versa. These tests were completedon flights 2-25and 1-50. Analyses of these data indicatethat the 35 Hzgain limitationsare unique to A/C-2. Ground tests were conductedon bothaircraftin an attempt to identifythe causes of the 35 Hz limitationsofA/C-2; these tests were unable to isolatethe causes of the 35 Hz.

CREW EVALUATIONSOF SMCS EFFECTS

The bulk of the SMCS operationalsuitabilitytests were conductedon B-I

A/C-3. Ten B-I flight crewmenparticipatedin the tests: four pilots (PLT),three flight-testengineers (FTE),and three OffensiveSystemOperators (OSO).The crews flew regularlyscheduledmanual terrain-following(MFF)and automaticterrain-following(ATF)flight-testmissions with the SMCS on and off. Crew-

memberswere instructedto maintain awarenessof comfortand personalper-formanceduring the various TF missions. Each crewmemberwas instructedtocomplete a test questionnaireto documenthis evaluationof the B-I ride andeffectsof this ride on his performance.

The curves in figure 151 through158 summarizethe subjectiveresponsesto questionson overall ride quality and the effect of turbulenceon: flightpath/nonflightpath controltasks; readabilityof instrumentsand displays;reaching/usingcontrols;crew fatigue,motion sickness,and physicaldiscom-fort. These data are in the form of compositeresponsesfrom all crewmembers.

Overall Ride Quality

Subjectswere asked to rate ride qualityduring TF flight as a functionof SMCS on and off for the followingconditionsof turbulence: smoothairand light,moderate,and heavy turbulence. The combinedratings for all Lsubjectsare shown in figure 151. The data reveal the following:

(i) Ride quality ratingsdecrease (ride qualityworsens) as turbulenceincreases.

(2) In smooth air, use of the SMCS results in little improvementinride quality.

(3) Use of the SMCS improvesride quality in light,moderate,andheavy turbulence.

220

N=7Poor- \ _ SMCS off

a"a,

c_ N = lO s N = lO

'_ Fair- SMCS on 4 PLTCD

3 FTE

_. 1,/ 30SO4J _ p

m Good- ,,$II N = 7 N = 7• _ •_S S

• cr ." 4 PLTN= I0

-o 2 FTE

Excellent -- l OSO

N = number of test

I I I I participantsSmooth Light Moderate Heavy

Turbulence rating

Figure 151. Ride quality ratings for varying degrees of turbulence.

Turbulence effect ratingN = 4 PLT

_orkload Performance

effect effect

Unable to perform task-

Large-- i _

Max effort Major--_. SMCS off

Moderate-- _

Moderate-- __ SMCS on

• Incr effort Minor--None --

- None [ None--t

I I ILight Moderate Heavy

Turbulence rating

Figure 152. - Effectsof turbulenceon flightpath controltasks.

221

Turbulence effect ratingN = lO

Workload Performance 4 PLT

effect effect 3 FTE .

Unable to perform task- 30SO

I Large - N = 7

Max effort Major-- N = 7 4 PLT

2 FTEModerate- N = I0 _-" SMCS off 10SO

{Moderate--_ _" N_ SMCS on_s.. _ '

Incr effort Minor-- f

None- = 7

NoneI None - N = lOI.

I I ILight Moderate Heavy

Turbulence rating

Figure 153. Effects of turbulence on tasks other than flight path control.

Figure 154. Effectsof turbulenceon readabilityof instrumentsand displays.

222

N=IOExtreme --

4 PLT

3 FTEc N=7

-*_ Moderate- _,,_ SMCSoff 3 050

11 N = 7

N = lO 11.1_ SMCS on

_./// ..__/_ _. 4 PLT\._u Slight-- X_ /////- \ 2 FTE

_ .//" % ..= 7 I OSO

None -- N = 10

I I I ISmooth Light Moderate Heavy

Turbulence rating

Figure 155. - Effects of turbulence on reaching/using controls.

Extreme-- N = 10N=7

c \ 4 PLT

"_ -_ ...- SMCS offm _-" 3 FTE

Moderate-- ._ SMCS on 30SO4-J

u N= I0 N=7

_- 4 PLTSIight --

. o N = 7 2 FTE10SO

"_ N= 10None --

I I I ISmooth Sight Moderate Heavy

Turbulence rating

Figure 156. - Effects of turbulence on crew fatigue.

223

c_ N= I0c Extreme-

SMCS off 4 PLTrO

SMCS on 3 FTE4-'

PIoderate- 30SOtl..,,

_- N=7iz1

N = lO 4 PLT

Slight- 2 FTEu 10SOt/3

r-

.o None -4J

0

I I I lSmooth Light Moderate Heavy

Turbulence rating

Figure 157. Effects of turbulence on tendency for motion sickness.

_- N = lO._Extreme --

fDL SMCS off 4 PLT

SMCS on 3 FTEU

_- Moderate- 30SO

N=74--1L.

o 4 PLT

OE Slight -- _ _ _k_1, 2 FTE'_ "f _'_ I oso ":g

None- N = 7 "U

U3 -->..¢-" I I I I

Smooth Light Moderate Heavy

Turbulence rating

Figure t58. El;fectsof turbulence on physical discomfort.

224

One subjectadded the followingcommentto his ratings: "In smoothair,SMCS is not needed. Standardof comparisonis the F-Ill which is excellent."(OSOA)

Flight-PathControlTasks

Pilotswere asked to rate the effectsof turbulenceon flight-pathcontroltasks during TF with SMCS off and on. Ratingswere to be made with referenceto a turbul_ence-effectrating scale. The combinedratingsfor all subjectsare shown in figure152. The data reveal the following:

(i) The effortto performflight-pathcontroltasks and the negativeeffect on task performanceincreaseas turbulenceincreases.

(2) Greatereffort is required,and the negative effect on subjectperformanceis greaterwith SMCS off than with SMCS on; i.e.,workload isless When SMCS is being used.

Non-Flight-PathControlTasks

All subjectswere asked to rate the effectsof turbulenceon non-flight-path controltasks during TF with SMCS off and on. Ratingswere made to aturbulence-effectrating scale. The combinedratingsare shown in figure153.The data reveal the following:

(I) The effortto performnon-flight-pathcontroltasks and the negativeeffecton task performanceincreaseas turbulence increases.

(2) Greatereffort is required,and the negative effecton subjectperformanceis greaterwith SMCS off than with SMCS on; i.e., workload is lesswhen SMCS is being used.

One subjectadded the followingcogent, "Operationof equipmentdifficult-particularlyCITS." (FFE C)

Readabilityof Instrumentsand Displays° -

All subjectswere asked to rate the degreeof difficultythey experiencedin reading instrunentsand displaysduring TF flight as a functionof SMCS offand on for four conditionsof turbulence: smoothair and light,moderate,and

225

heavy turbulence. The combined ratingsfor all subjectsare shown infigure154. The data reveal the following:

(1) The difficultyto read instrumentsand displays increasesasturbulenceincreases.

(2) Difficultyratingsare higher (readabilityless difficult)with SMCS ""on than with SMCS off.

One subjectadded the followingcon_nent:"'E' scope always extremelydifficultto read because of location - turbulenceno factor."

Reaching/UsingControls

All subjectswere asked to rate the difficultyin reachingcontrolsorin performingcontrolactionsduringTF flight as a functionof SMCS off andon for four conditionsof turbulence: smooth air and light,moderate, andheavy turbulence. The combinedratingsfor all subjectsare shown in fig-ure 155. The data reveal the following:

(i) The difficultyto reach and use controls increasesas turbulenceincreases.

(2) As turbulenceincreases,it is easier to reach and use controlswithSMCS on than with SMCS off.

Five of the subjectsadded a conlnent,as follows:

(i) "In manual TF, pitch stick force too heavy." (PLT C)

(2) "Ratingis based primarilyon use of central integratedtestsystem (CITS)."(FTEA)

(3) "CITS controlssomewhatdifficultin turbulence." (FTEB)

(4) "Operationof CITS is difficultat best. SMCS helps." (FTE C)

(5) "Some of the ratingsare due to the locationsof the controls."(OSO C)

226

Crew Fatigue

All subjectswere asked to rate the degree to which TF flight introducedany specialtendenciestoward fatigue. Ratingswere requestedfor the SMCSoff and on modes for four conditionsof turbulence: smoothair, and light,

_ moderate,and heavy turbulence. The combinedratings for all subjectsareshown in figure 156. The data reveal the following:

" (I) Fatigue effectsincreaseas turbulenceincreases.

(2) There is little differencein fatigue-effectratingswith SMCS offversus SMCS on except for the moderate turbulencecondition. For moderateturbulence,the fatigue-effectsrating for SMCS on is higher (less fatigueeffect)than for SMCS off.

Three of the subjectsadded a con_nent,as follows:

(i) "Ratingsmade for AUFO TF." (PLTA)

(2) 'Turbulenceis not the primary factor in producingfatigue - TFitselfproducesa high level of fatigue." (FTEB)

(3) "Turbulencedefinitelyincreasesfatigue." (FTEC)

Motion Sickness

All subjectswere asked to rate the degree to which TF flight introducedany specialtendency for motion sickness. Ratingswere requestedfor theconditionsof SMCS off and on for four levelsof turbulence: smoothair and

light, moderate,and heavy turbulence. The combinedratings for all cre_nenare shown in figure 157. The data revealthe following:

(I) The compositedata show little tendencyfor motion sicknessfor allturbulenceconditionsfor both SMCS off and on modes. FTE were more affected

than pilots.

: One subjectcommentedthat: 'Turbulenceis not a major influence. The rough-ness of terrainhas more impactwhen you are stuffedin that 'blackhole' withno windows." (FTEC)

227

PhysicalDiscomfort

All subjectswere asked to rate the degree to which TF flight introducedany specialtendenciestoward physicaldiscomfort. Ratingswere requested:forthe conditionsof SMCS off and on for four levelsof turbulence: smooth

air and light,moderate,and heavy turbulence. The combinedratingsfor allcrewmen are shown in figure 158. The data'revealthe following: -"

(I) Physicaldiscomfortincreasesas turbulenceincreases;althoughatheavy turbulence,the effectsare rated only as slight.

(2) Less physicaldiscomfortis reported for light,moderate,and heavyturbulencewith SMCS on than with SMCS off. For smooth air, there is nodifferencein physicaldiscomfortratingsfor the SMCS off versus SMCS onmodes.

One subjectcommentedthat: "Physicaldiscomfortresults from feelingsofirritation,aggravation,and anxietyproducedby the rough ride and highworkload,hard to pin down further." (PLTD)

AdditionalRide QualityObservations

All subjectswere asked to add any additionalobservations(not coveredby the ride qualityquestions)concerningride qualitycharacteristicsorSMCS effectsduring TF flight or any other characteristicsrealtingto TFwhich have a bearingon crew comfortand efficiency. Comments includedthefollowing:

(i) "SMCS is very effectiveand required in the B-I to aid the flightcrew in performanceof the TF task." (PLTA)

(2) "Essentialfor effectiveB-I MrF, desiredfor effectiveB-I ATF."(PLT C)

L

(3) "Turbulencemakes this aircraft hard to stabilizeon a bank angle, --adds to an alreadyhigh workload." (PLT D)

(4) "Ride qualitieswith SMCS are definitelybetter than without - thedifferenceI don't think is reallyable to be seen in the layout of thisquestionnaire." (FTEA)

(5) "Lack of outsidevisual referencecontributessignificantlytodisorientationand motion sickness. Movementabout the crew compartmentishamperedby TF, especiallyin turbulence. Any movement in the crew compartment

228

also contributesto disorientationand motion sickness. There is a significantimprovementin the ride with SMCS on." (FTEB)

(6) "I considerSMCS essentialfor long term TF flight." (OSOA)

(7) "Due to part of the controlsand displayslocation,it is difficultto accomplishsome weapon-orientedtasks under turbulentconditions. Thesensitivityof both the navigation (NAV)panel and Stores ManagementSystem(SMS)FWD/REVswitchmake it difficultto use under turbulentflight condi-tions." (OSOB)

(8) 'The discomfortsin the OSO stationare are severein light ormoderate turbulencewith SMCS off due to the fact that we have no outside

reference." (OSOC)

HandlingQualities

Pilots were asked to describethe effects,if any, of SMCS activationonaircrafthandlingqualitiesduringMTF. Commentsincludedthe following:

Pitch Control

(I) "A/C ride is smootherwith SMCS on and thereforecontrolis easier."(PLT A)

(2) "Assists by damping." (PLT B)

(5) "COUld have some detrimental effect because rapid control inputsare apparently countered by SMCS -- not considered a problem to date, moreevaluation required." (PLT C)

(4) "Increases pitch forces, se_ns to slow aircraft response." (PLT D)

Lateral/Direct ional Control

(1) "A/C ride is moother with SMCSon and therefore control is easier."(PLT A)

(2) "Assists by damping." (PLT B)

229

APPENDIX

NONENCIATURE

This reportwas the resultof the contributionsof a number of authorsand each has used nomenclatureunique to his particulardiscipline. In order -_to help the reader to quicklylocate a given symbol,this sectionhas beenorganizedso that a general section,is presentedfirst followedby nomenclatureassociatedwith three sectionsof this reportwhich are especiallyheavy inspecializednomenclature. These sectionsare: "FlexibleAircraft Equationso[ Motion," "Impactof SMCS on SelectedLoads,"and "_S Vane Effect on Inlet/Engine Characteristics."Under this system,similarsymbolsoften have differ-ent meanings;the reader is cautionedto identifysymbolswithin the contextoftheir use.

GENERAL

A amplitudesettingof B-I oscillatingsystem for the SMCS vanes

A/C aircraft

alt altitude

ATF automatic terrain following

AUTO automatic

BP butt plane

b[_F referencelength

mean aerodynamicchord

CITS centralintegratedtest system

CG center of gravity

cm centimeters

deg degrees

El bending stiffness

[ frequency,cycles per second

F, Flex subscriptdenoting flexible

230

F1t flight

ft feet

];l'Ji flight test engineer

[:/R flexible-to-rigidratio

[

]R flexible-to-rigidratio of bracketedparameter[

FS fuselage station

g Accelerationof gravity

CJ torsionalstiffness

GVT ground vibrationtest

crew sensitivityindex; subscriptZ denotesvertical axis,li() Y denoteslateralaxis

Hz hertz (cyclesper second)

hr hours

in. inches

_K _bREFk reducedfrequency,_ , V

o o

K SMCS gain

Khp SCAS gain scheduledwith altitude

kg kilogram

- KMNn yaw SCAS lateralaccelerometergainY

kq pitch SCAS gym gain

k_ SMCS gain scheduledwith dynamicpressureq

231

K_ r yaw SCAS gyro gain

Knz pitch SCAS normal accelerationgain

ib pounds

L-gain lateralSMCS gain

Lh dimensional,force along Z-axisdue to plungingmotion h,+ down

Le , dimensionalforce along Z-axisdue to pitchingmotion o,+ downJ

[ALO]Real [L@I Real V° [Lh]Imag

Lni dimensionalforce along Z-axis due to structuralmodegeneralizedcoordinatemotion hi, + down

Lwg dimensionalforce along Z-axisdue to verticalgust velocityVCg,+ down

L_ dimensionalforce along Z-axis due to control surfacemotion _,+ down

m meter

M Mach number

MrF manual terrainfollowing

Mh dimensionalmoment about Y-axis .dueto plungingmotion h,+ nose up

Mo dimensionalmoment about Y-axis due to pitchingmotion 6,+ nose up

Mni dimensionalmoment about Y-axis due to structuralmode gen-eralizedcoordinatemotion hi, + nose up

232

Mwg dimensionalmoment about Y-axis due to verticalgust velocityw , + downg

N newton

N ntmber of test participants

Qih dimensionalgeneralizedforce in structuralmode i due to

plungingmotion h, + for ni increased

Qi@ dimensionalgeneralizedforce in structuralmode i due to

pitchingmotion @, + for Ni increased

[AQi_IRoal IQi0]Real-(_) [Qih]imag

Qini dimensionalgeneralizedforce in structuralmode i due tostructuralmode generalizedcoordinatemotion hi, + for niincreased

Qiwg dimensionalgeneralizedforce in structuralmode i due toverticalgust velocityWg, + for Ni increased

Qi6 dimensionalgeneralizedforce in structuralmode i due tocontrol-surfacemotion _, + for ni increased

SL sea level

V-gain vertical SMCS gain

WL waterline

dampingratio

@ pitch angle about elastic axis

A sweep angle of leading edge of lifting surface

a Vertical gust intensityderived from angle-of-attackvane_Vmeasurements

OBv lateralgust intensityderivedfrom sideslipangle vanemeasurements

elasticaxis-bendingslope

233

•() power spectraldensityof subscriptparameter

frequency,radiansper second

RIDE QUALITYEQUATIONSOF MOTION RELATED

qo 1/2 pV2, dynamicpressure

p density of air

AP

_o pressurecoefficient

Vo resultantvelocity of the CG

componentof resultantvelocity;subscriptdenotesaxisV r ) along which componentacts

Wg verticalcomponentof gust velocity

Vg lateralcomponentof gust velocity

Sw wing area

bw wing span

_w wing mean aerodynamicchord

Z vertical deflection

Y side deflection

Zx' Zy' _z distancealong the x, y, and z-axis,respectively

- - distancefrom vehicle CG to control surfacek hingeline -_' z( ) (+ aft), subscriptidentifiessurface

distance from control surfacehingelineto surfaceCG_' _( ) (+ aft), subscriptidentifiessurface

distancefrom X-axis to surfaceCG in-planeperpendicularto) plane of symmetry (always+), subscriptidentifiessurface.

_°F perpendiculardistancefrom CG to thrust axis; + down

234

W airplane weight

m airplane mass

mass of control surface,subscriptidentifiessurfacem, m()

M. the ith mode generalizedmass,1

f f f dX®dzy,z) Am(x,y,

Ix, Iy, Iz moment of inertiaabout the X-, Y-, and Z-body axis, respectively

I product of inertia; positive when the principal X-axis isxzbelow the body axis at the nose of the vehicle

IR engine rotor moment of inertia

I( )[L moment of inertiaabout hingeline; ( ) subscriptidentifiessurface

n( ) load factor;subscriptdenotesaxis along whichcomponent acts

Euier azimuthangle

@ Euler pitch angle

Euler roll angle see figure 6

X, Y, Z body-axiscoordinates

Xe' Ye' Ze earth-axiscoordinates

11 altitude (+ up from sea level)

235

p rollingrate about X-body axis •

p rollingaccelerationabout X-bodyaxis

q pitchingrate about Y-body axissee figure 2

q pitchingaccelerationabout Y-body axis

r yawing rate about Z-body axis

r yawing accelerationabout Z-bodyaxis

_R rotationalrate of engine rotor relativeto airframe

_. natural frequencyof ith modei

forcingfrequency

angle of attack;angle betweentheprojectionof the resultantvelocityvector on the XZ-planeand the X-body

(referenceaxis) _ see figure 5|sideslipangle; angle betweenthe Iresultantvelocity vector and the Jplane of symmetry XZ

I" control-vane dihedral angle

6 con.trol-surfacedeflection;positive deflectionproduces

( ) positive force (+CN,+Cy)(seefigure 3); subscriptidentifiessurface

control-surfaceacceleration;positive in the sense that 6( )6( ) is positive;subscriptidentifiessurface ,.

6H rolling tail controldifferentialdeflection,+ deflectionproduces +Cz

g accelerationof gravity

gsi structuraldampingconstant,mode i

236

n. deflectionof the ith normalizedstructurali mode at normalizationpoint

n. rate of change of the ith mode at pQint ofi normalization

_. accelerationof the ith mode at point ofi normalization > see figure 4

¢.( ) the i normalizedmode shape; i.e., ratio ofi local deflectionto deflectionat normalizing

point (nondimensional); ( ) superscriptdenotes location

_! )' slope of the ith normalizedmode; ( )i superscriptdenoteslocation

T

_! ) fuselagetorsionalangle, ( ) superscriptdenotes location1

F force

Z aerodynamicforce in Z-direction

N aerodynamicnormal force (N = -Z)

L aerodynamiclift force (L_N for small _)

C aerodynamicchord force (C = -X)

D aerodynamlcdrag force (D_ C for small _)

X aerodynamicforce in X-direction _ see figure 3

Y aerodynamicforce in Y-direction

T thrust (T = X)

L aerodynamicrollingmoment about X-axis

M aerodynamicpitchingmoment about Y-axis

N aerodynamicyawing moment about Z-axis

237

Nwg normal force due to unit vertical gust velocity

Mwg pitchingmoment due to unit vertical gust velocity

Q[ generalized force in ith mode fffF(x,y,,) i(x,y,z)dxdydz ..

generalized force in structural mode i due unit verticalQiwg gust velocity

Yvg side force due to a unit lateral gust velocity

rollingmoment due to a unit lateral gust velocityLvg

Nvg yawing moment due to a unit lateralgust velocity

Nni yawing moment due to mode i deflection

Yni side force due to mode i deflection

M_i pitching moment due to mode i deflection

Nni yawing moment due to mode i deflection

Lni rollingmoment due to mode i deflection

Qij generalizedforce in mode i due to mode deflectionj

C chord-forcecoefficientCC = Swqo

DCC= --chord-force coefficientdue to control-surfacedeflection,

CC6( ) a6()subscriptidentifiessurface

N . normal-forcecoefficientCN = %%

OCNCNf Off normal-force curve slope

238

8CNnormal-forcecoefficientdue to downwashlag and

CN_ = _Cw verticalacceleration

8CN

CNq = -/q Cw_ normal-forcecoefficientdue to pitch rate

8CN

CN__ = a/Q_.2\lwl normal-forcecoefficientdue to pitch acceleration\4V2/

8CN

CNni = 8Ni normal-forCedeflectioncoefficientdue to structuralmode

8CN normal-force coefficient due to structral mode

CNni = 8_, deflection rate

8CN normal-forcecoefficientdue to controlsurface

CN6 = 86( deflection,subscriptidentifiessurface() )

8CN normal-forcecoefficientdue to controlsurface

CN6( =) On deflectionrate, subscriptidentifiessurface• ()

M

Cm = °w_w_=-o pitching-momentcoefficient

8CmCm_ = 8_ pitching;moment curve slope

239

_Cm pitching-momentcoefficientdue to downwashlagCmt_ = &I \ and verticalacceleration

3Cln __

Cmq = __qC-w_ pitching-momentcoefficientdue to pitch rateH I

-\2Vo!

aCm

Cmq = /_F.,2\ pitching-moment coefficient due to pitch acceleration

_Cm pitching-moment coefficient due to structuralCm_i = 3ni mode deflection

3Cm pitching-momentcoefficientdue to structuralmode

i _ deflectionrate

Cm pitching-momentcoefficientdue to control surface

Cm6( =) 36( ) deflection,subscript identifiessurface

8Cm pitchingmoment due to control surfacedeflectionrate, subscriptidentifiessurface

) )°

= a subscript_, B, etc, as shown indicatesthat()(_) bracketedparameter is a nonlinearfunctionof the

subscriptvariable

Qi generalized-forcecoefficientin the ithCni =

Swq° mode

240

C . = __8C_i_ generalized-forcecoefficientdue to angle of attack_i_ 8_ in ith mode

8Cni generalized-forcecoefficientdue to downwashlag

C_i& = 81_Cw_1-----'_ and vertical acceleration

VV°I

aCni generalized-forcecoefficientdue to pitch rate

C_iq = /qC-w_ in ith mode

3Cni generalized-forcecoefficientdue to pitch

Cni_l -tqcw21 acceleration in the ith mode

_CNi generalized-forcecoefficientdue to the jth mode

C_I_j" = _n]. shape in the ith mode

3CNi generalized-forcecoefficientdue to rate of change

C_i_J -- (-_o)

"" fli of jth mode in ith mode

aCni generalized-forcecoefficientdue to control-surfacedeflection in ith mode, subscript identifies control

CrJi_( ) a6( ) surface

aCNi generalized-forcecoefficientdue to control-surface

=) a6( ) k deflectionrate in ith mode, subscriptidentifiescontrolforce

YCy -- - side-force coefficient

qoSw

aCy

Cy_ = _s side-forcecoefficientdue to angle of sideslip

241

_Cy side-forcecoefficientdue to rate of change ofCy_i = O/_bw\ angle

12---_°_ sideslip (lateralacceleration)

CyBqi = 3CyB side-forcecoefficientdue to sideslipangle caused -_by symmetricstructuralbending (dihedral)i

_Cy

Cyr = side-forcecoefficientdue to yaw rate01 l\2Vo!

2%

CYr = a/_b,.2\,___I side-forcecoefficientdue to yaw acceleration\4Vo2 /

_Cy

Cyp = /pbw \ side-forcecoefficientdue to roll rate

_Cy

Cy_ = a/_bw2\| | side-forcecoefficientdue to roll acceleration\4Vo2/

_Cy side-forcecoefficientdue to control-surfaceCy6 =

( ) _6( ) deflection,subscriptidentifiessurface

_Cy side-forcecoefficientdue to control-surface

Cy_( = _() ) deflectionrate, subscriptidentifiessurface _

242

J

N

= Swbwqo yawing-momentcoefficientCn

_Cn

C_fl = _ yawing-momentcoefficientdue to sideslipangle

_Cn yawing-momentcoefficientdue to rate of changeCnfi =

_bw) of sideslipangle (lateralacceleration)

;CnB yawing-moment coefficient due to sideslip angle

Cn_ni _i caused by symmetricstructuralbending (dihedral)

_Cn

Cnr = /rbw\ yawing-momentcoefficientdue to yaw rateI I

_CnCn. = yawing-momentcoefficientdue to yaw acceleration

r i_bw2_

_Cn

Cnp = 'Jpbw\lI yawing-momentcoefficientdue to roll rate\2Vo/

_Cn

Cn_ = D/pbw2_ yawing-momentcoefficientdue to roll acceleration

\4Vo J.

_Cn= yawing-momentcoefficientdue to control surface

Cns( ) 26( ) deflection,subscriptidentifiessurface

243

aCn yawing-momentcoefficient due to control surface(:n_ =

( ) a_( ) deflection rate, subscript identifies surface

LCI = rolling moment

Swbwqo --

aCl

Cz_ = aB rolling-momentcoefficientdue to sideslipangle

aC_ rolling-momentcoefficientdue to rate of change --C_ =

81_bwl of sideslipangle (lateralacceleration)\ -vo!

aCz8 rolling-momentcoefficientdue to sideslipangle

C_Bni = ____aNi caused by symmetricstructuralbending (dihedral)

aC_

C_r = /rbw\ rolling-momentcoefficientdue to yaw rate

_C_

C_r = /+h2\ rolling-momentcoefficientdue to yaw acceleration

aC

Cgp = o/pb\|___Z_.lrolling-momentcoefficientdue to roll rate

ac= _

Cg_ /nh2\ rolling-momentcoefficientdue to roll acceleration

i wj8\4V02/

244

aC£ rolling-momentcoefficientdue to control surface

C_6( ) a6( ) deflection,subscriptidentifiessurface

aC_ rolling-momentcoefficientdue to control surface

_ C_( ) an( ) deflectionrate, subscriptidentifiessurface

aCNi generalized-forcecoefficientdue to sideslipangleC_i = a6 in ith mode

8

_CNi generalized-forcecoefficientdue to sideslipangleCni" =

6 l_bw_ rate of change in ith mode

aC.ni= 6 generalized-forcecoefficientdue to sideslipin the

Cni6 an. ith antisymmetricmode due tobending in the jthnj J symmetricmode

aCni generalized-forcecoefficientdue to yaw rate in

Cnir =Irbw_ ith mode

aC_i generalized-forcecoefficientdue to yaw accelerationCni" = • 2_ in the ith mode

r rbw

_ aCni generalized-forcecoefficientdue to roll rate in

CHip =pbwl ith mode

\2Vo!

_C_i generalized-forcecoefficientdue to roll acceleration

Cni_ = i_bw2 _ in the ith mode

a t[ 14Vo2 I

245

qoSwKs =

mVo-

qoSw_wKq = ly

qoSwK# = mVo

qoSwbwKr =

Iz

qoSwhwKp =

Ix

qoSwKni = Mi

i /1-

i, j used as subscriptto identifystructuralmode

k subscriptidentifyingcontrol surface

II subscript identifying horizontal-tail control surface

r subscriptidentifyingrudder

cv subscriptidentifyingstructuraln_de controlvaner

R subscriptindicatingreal part

I subscriptindicatingimaginarypart

o subscript indicating trim value

tL hingeline

I;RL fuselagereferenceline

246

MAC mean aerodynamicchord

( )TF aerodynamictransferfunction

- LOAD EQUATIONSOF MOTION RELATED

a accelerationl

A gust response factor Oo/Ow

b turbulencefield parameterdenotinggust intensity

IB ] phased loadingconditions,one column per condition,each

elementproportionalto Pi-3 °" _"i 3

IE J expectedload values for each load it_n

{F} forces at each SIC point: real or in the frequencydomain,complex

g structuraldampingparameter

Ih I real generalizedcoordinates

H(_) frequencyresponsefunctionfor a load item

II*(_) denotes complexconjugateof H(_)

Ill(w)] frequencyresponsefunctionsof load items, one row per itemfor i01 frequencies

j #T

- [LOADS ] loads; shears,moments, and torques;real or in the frequencydomain,complex

LOAD geometryto compute shears,moments,and torques,one rowGEOM per load item, one column for each SIC load point

IDAD I load generationmatrix, one row per load item, one columnGEN 3 for each normal elasticmode

EM_ mass matrix

247

[_M] generalizedmass matrix for rigid-bodymotions

[M_ generalizedmass matrix = [M][_],rigid and elasticmodes -_[¢]T

\

N number of crossings of zero per hour with a positive slopeo

N(y) averagenumber of level crossingswith a positive slopeequallyor exceedingy per hour

P turbulencefield parameterdenoting proportionof time inturbulence

q I complexgeneralizedcoordinates

QM] generalizedaerodynamicforces due to modal motion

Qc] genezalizedaerodynamicforces due to unit deflectionsofcontrolsurface,column 1 for horizontaltail and c61umn 2for the mode controlvane

_iI generalizedforces due to a unit sinusoidalgust

R] rigid-bodymode-modifyingmatrix

S Laplace operator

S_ phased loadingconditions,one column per condition,eachcolumn scaled so that the diagonal elementequals the corre- .-'-spondingexpectedload value

[ Sl C] structural flexibility influence coefficients

ISLOADS I phased loading column from [S],shears,moments, and torques

IT .] transformationmatrix,for controlsystem feedback,relatingcontrol-surfacedeflectionsto the generalizedcoordinates

V velocity

248

_H' _cv control-surfacedeflections;subscriptH for horizontalstabilizer,subscriptV for mode controlvane

mode deflection

L[ ¢]J mode shapesby columns, superscriptdenotestype of modes;RBM for rigid-body,and E for normal elastic

_w' _o root mean-squarevalues of gust velocity and output itemresponse,respectively

oi, o. root mean squarevalue of responseitems i and jJ

_(_) gust power spectrum,normalizedon unit gust intensity,o2W

(_) load item output responsepower spectrumo

pitch rate

o.. correlationcoefficientexpressingthe degreeof statisticalij linear dependencebetween load item i and load item j

frequency

structuralnormal elastic-modefrequency

] squareor rectangularmatrix;where i = number of rows, andi,j j = number of colunns

E _ diagonalmatrix

cQlumn matrix

,_. L _] row matrix

[_T matrix transpose

[] matrix inverse

249

ENGINE/INLET RELATED

A/l) analog to digital

AIP inlet/engine aerodynamic interface plane

AIS air induction system

Aduct duct area, psi/count

CAL calibration

CBW constant bandwidth

CTS counts

CTSCAL counts output during calibration step

CTS_4EAN mean value in counts during operate step

CTSOP counts output during operate step

CTSZ counts output during zero step

db decibels

I!CI-8 inlet configuration identification

FS full scale

HP high pass

IDC circumferential distortion component _.

IDL engine stall-margin ratio (function of IDC and IDR)

IDR radial distortion component

ips inches per second

K i000 feet

LP low pass

25O

NF engine fan speed

N2 nitrogen

P pressure

_P pressure difference

PCM pulse-codemodulation

PDYN Instantaneouspressure (dynamic)

PLA engine power lever angle

PPCM time-averagedpressure (steadystate)

pps pounds per second

psi pound per square inch

PT total pressure

PTI average total pressureat inlet/engineAIP

VrI local total pressureat inlet/engineAIP

PTO free-streamtotal pressure

FIIv_X maximum total pressure at inlet/engineAIP

VFMIN minimum total pressure at inlet/engineAIP

RB first movable inlet ramp angle

RC secondmovable inlet ramp angle

R2 referencepressure

SMCV structuralmode controlsystemvane

sps sampleper second

VCO voltagecontrolledoscillator

WIR correctedengine airflow

251

X2 amplifier (times 2)

X6 amplifier (times 6)

ZOC a three-way value (refer to page 161)

t_

Z_ fractional increment on total pressurecontour, AIP

6002 reference pressure, absolute units

6631 calibration pressure, absolute units -_

252

REFERENCES

I. Wykes, John H.; Borland,ChristopherJ.; Klepl,Martin J.';and MacMiller,

_ Cary, J.- Design and Developmentof a StructuralMode Control System,NASA CR-143846,October1977.

2. Etkin, Bernard: Dynamicsof Flight. John Wiley and Sons, Inc., New- York.

3. Bisplinghoff,RaymondL.; Ashley,Holt; Halfman, RobertL." Aeroelasticity.Addison-WesleyPublishingCompany, Inc., Reading,Massachusetts.

4. Landahl,M. T.: GraphicalTechniquesfor AnalyzingMarginallyStableSystems. Journalof Aircraft,September-October1964.

5. Rustenburg,John W.: Developmentof TrackingError FrequencyResponseFunctionsand Aircraft Ride QualityDesign Criteriafor Vertical andLateralVibration. ASD-TR-70-18,January1971.

6. Stenton,Thomas E.: TheoreticalFrequencyResponseFunctionsand PowerSpectraof the XB-70 Responseto AtmosphericTurbulence. NASA CR-1621,August 1970.

7. Dugundji, John: On the Calculationof NaturalModes of Free Free

Structure. Journalof the AeronauticalSciences,February1961.

8. Flight ControlSystem Description,Rockwell International,North American

Aircraft Division,E1 Segundo,TFD-71-807_June 1971 (RevisionAugust19753.

9. Austin, WilliamH., Jr.: Developmentof ImprovedGust Load CriteriaforUnited StatesAir Force Aircraft. SEG-TR-67-28,September1967.

i0. Fuller,J. R.; Richmond,L. D.; Larkins,C. D.; and Russell,S, W.:Contributionsto the Developmentof a Power SpectralGust Design Procedurefor Civil Aircraft. FAA-ADS-54,January1966.

253

J

1. ReportNo. 2. GovernmentAccessionNo. 3. Recipient'sCatalogNo.NASA CR-144887

4. Title and Subtitle 5. Report Date

January. 1980ANALYSES AND TESTS OF THE B-1 AIRCRAFT 6. PerformingOrganizationCodeSTRUCTURAL MODE CONTROL SYSTEM

7. Author(s) 8. PerformingOrganizationReportNo.John H. Wykes, Thomas R. Byar, Cary J. MacMiller, and NA-79-405David C. Greek

10. WorkUnitNo.9. PerformingOrganizationNameandAddress

Rockwell InternationalNorth American Aircraft Division 11. Contractor GrantNo.

815 Lapham Street NAS4-2519El Segundo, CA 90245 "

13. Typeof ReportandPeriodCovered12. SponsoringAgencyNameandAddress

Contractor Report - Final

National Aeronautics and Space Administration 14. SponsoringAgencyCodeWashington, D.C. 20546

H-1109

15. SupplementaryNotes

NASA Technical Monitors: Jim MeKay and Larry Felt, Dryden Flight Research CenterNASA Program Manager: Jack Nugent, Dryden Flight Research Center

16. Abstract

An 18-month program was conducted to compile and document for publicationinformation pertaining to analyses and flight tests of the B-1 Structural ModeControl System (SMCS). This is the second phase of a c_ntinuing effort;results from the first phase study are documented in De_;ign and Developmentof a Structural Mode Control System, by John H. Wykes, Christopher J. Borland,Martin J. Klepl, and Cary J. MaeMiller (NASA CR-143846, October 1977). This

report covers the_following topics: _'_

(1) Flexible aircraft equations of motion

(2) Description of flexible aircraft analyses model

(3) Comparison of analyses and flight-test performance results of the SMCS

(4) A summary of the study of the forward SMCS sensor package relocation

(5) Truncated analytical models used in simulation effort

(6) An analysis of the SMCS vane interference effects

(7) Impact of SMCS on selected loads

(8) Flight-test results of the SMCS vane effects on inlet/engine characteristics

r (9) Summary of SMCS flight-test results

17. Key Words(SuggestedbyAuthor(s)) 18. DistributionStatement

Structural mode control systemLarge flexible aircraftAnalyses and tests

STAR category: 05

19. SecurityClassif.(ofthisreport) 20. SecurityClassif.(ofthispage} 21. No.of Pages 22. Price*Unclassified Unclassified 268

*For sale by the National Technical Information Service, Springfield, Virginia 22151

i"

,t

D

3 1176 00156 1985 - ntrs.nasa.gov

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