Xv-8a Flexible Wing Aerial Utility Vehicle

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AD460405

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TOApproved for public release, distributionunlimited

FROMNotice: All release of this document iscontrolled. All certified requesters shallobtain release approval from Secretary ofDefense, Attn: ARPA-AGILE, DefenseAdvanced Research Projects Agency, DARPA,3701 North Fairfax Drive, Arlington, VA,22203-1714.

AUTHORITY

DARPA ltr, 20 Nov 2001

THIS PAGE IS UNCLASSIFIED

USATRECOM TECHNICAL REPORT 64-74

XV -SA flEXIBLE WING AERIAL UTILiTY vEH~ClE

FINAL REPORT

,

By

F. Landgraf

P. F. Girard

U. S. ARMY TRANSPORTATION RESEARCH COMMAND

FORT EUSTIS, VIRGINIA

CONTRACT DA 44-177 -AMC-87 4{T)

RYAN AERONAUTICAL COMPANY

SPONSORED B'1:

ADVANCED RESEARCHPROJECTS AGENCY

DDC Availability Notice

All distribution of this report is controlled. Qualified DDC usersshall reque; . through Secretary of Defense, ATTN: ARPA-A( ILE,Washington, D. C. 20310.

Release or announcement to the public is not authorized.

Reproduction Limitations

Reproduction of this document in whole or in part is prohibited exceptwith permission of the Secretary of Defense.

Disclaimer

The findings in this report are not to be construed as an official

Department of the Army position, unless so designated by other author-ized documents.

When Government drawings, specifications, or other data are usedfor any purpose other than in connection with a definitely relatedGovernment procurement operation, the United States Governmentthereby incurs no responsibility nor any obligation whatsoever, and the

fact that the Government may have formulated, furnished, or in any waysupplied the said drawings, specifications, or other data is not to beregarded by implication or otherwise as in any manner licensing the

holder or any other person or corporation, or conveying any rights orpermission, to manufacture, use, or sell any patented invention thatmay in any way be related thereto.

'4'

Disposition Instructions

Destroy this report when it is no longer needed. Do not return it tothe originator.

-- -. :

HEADQUARTERSU S ARMY TRANSPORTATION RESEARCH COMMAND

FORT EUSTIS, VIRGINIA 23604

This report is a summary presentation of the aerodynamic analysis,stress analysis, structural tests, and initial flight evaluationof a Flexible Wing Aerial Utility Vehicle that is designated theXV-8A. The XV-8A was designed and built by Ryan Aeronautical Companyin accordance with the requirements of Contract DA 44-177-AMC-874(T),which was initiated by the U. S. Army Transportation Research Commandand funded by the Advanced Research Projects Agency.

The conclusions reached in this report are concurred in by thisCommand. Based on these conclusions, the aircraft has been modified,and a follow-on flight test program has been conducted to determinethe performance and handling qualities (reference USATRECOM Technical

Report 64-55).

This Command gratefully acknowledges the assistance provided duringthis project by the Airborne Operations Section, Yuma Proving Ground,Yuma, Arizona.

I

Il

Task ID121401A14172Contract DA 44-177-AMC-874(T)

USATRECOM Technical Report 64-74

February 1965

XV-8AFLEXIBLE WING AERIAL UTILITY VEHICLE

REIP ORT NO. 65B003

This research was supported by the Advanced Research ProjectsAgency of the Department of Defense and was monitored by theU.S. Army Transportation Research Command (USATRECOM)under Contract DA 44-177-AMC-874(T).

Prepared byRYAN AERONAUTICAL COMPANY

SAN DIEGO, CALIFORNIA

A'.r

U.S. ARMY TRANSPORTATION RESEARCH COMMANDFORT EUSTIS, VIRGINIA

- "- - - - " -J

ABSTRACT

This report entitled XV-8A FLEXIBLE WING AERIAL UTILITY VEHICLEis a final report based upon the Contractor's abridged Report No. 64B08dated August 1964 which contained 96 pages, 29 illustrations and 13 tables.The contract number was DA 44-177-AMC-874(T), ARPA Order 294-62Amendment No. 3, Unclassified. This final report and the abridged pre-liminary report discussed procedures and accomplishments of the design,fabrication and test program of the XV-8A for the following phases:

1. Investigation of the feasibility of construction of a Flexible WingLight Utility Vehicle that was simple to operate and capable of transport-ing a 1,000-pound payload for a distance of 100 miles at a speed of 50miles per hour.

2. The design, building and testing of two vehicles.

3. The structural tests at the Contractor's plant and the flight tests atthe U.S. Army Proving Ground, Yuma, Arizona.

4. The conclusion that the vehicles are feasible, and that further testsare recommended.

iii

____ ___ ____ ___ ___------------------flfjjj

-..

FOREWORD

The design, fabrication, and test program discussed in this report wasconducted under the provisions of Contract DA 44-177-AMC-874(T)between the U.S. Army Transportation Research Command and the RyanAeronautical Company.

The vehicle described herein is a second generation outgrowth of a flex-ible wing manned test vehicle previously developed by Ryan AeronauticalCompany. Tests on the original vehicle were made under the provisionsof Contract DA 44-177-TC-721 as reported in TCREC Technical Report62-25.

Structural tests described herein were conducted at the contractor'splant at San Diego, California. Taxi tests and flight tests were conductedat the Yuma Proving Ground, Yuma, Arizona.

v

CONTENTS

Page

ABSTRACT iii

FOREWORD v

LIST OF ILLUSTRATIONS ix

LIST OF TABLES xi

LIST OF SYMBOLS xii

SUMMARY 1

CONCLUSIONS 2

RE COMMENDATIONS 4

DESCRIPTION 5

AERODYNAMICS ANALYSIS 25

STRESS ANALYSIS 44

STRUCTURAL TESTS 53

FLIGHT-TEST EVALUATION 76

MAINTENANCE REQUIREMENTS 92BIBLIOGRAPHY

93

DISTRIBUTION 95

vii

...- ....._._._!glo w

ILLUSTRATIONS

Figure Page

1 In-Flight View, Flexible wing Aerial Utility Vehicle 11

2 Final Assembly Model 164 13

3 Aerial Utility Vehicle Platform Assembly 15

4 Wing Assembly and Installation 17

5 Main Landing Gear and Brake Assembly 19

6 Fuel System Installation 21

7 Controls Installation 23

8 Performance Data 27

9 Stability and Control Data 30

10 V-n Diagram for Symmetrical Flight 46

11 Reierence Data for Weight and Center of 51Gravity Calculations

12 Loading and Burn-Off Curve 5213 Whiffle Tree Configuration 54

14 Static Test Setup 55

15 Keel Deflection 55

16 Cockpit Con trols Proof Loading Test Configuration 58

17 Control Column, Control Wheel and Brake Pedal 59Deflection

18 Keel Deflection Under Load 59

ix

ILLUSTRATIONS (Continued)

Figure Page

19 Deflection Point and Strain Gauge Location, 81Ruddervator

20 Shot Bag Configuration, Ruddervator 61

21 Ruddervator Failure 62

22 Ruddervator Failure, bhot Bags Removed 62

23 Ruddervator Deflection Data 63

24 Drop Test Configuration 66

25 Landing Gear Test Setup and Axle Failure 68

26 Landin& Gear Test, Strain Gauge Location 74

27 Larding Gear Test, Load Vs. Strain 74

28 Landing Gear Test, Load Vs. Deflection 75

29 Performance Data, Level Flight 85

x

TABLES

Table Page

1 Longitudinal Dynamic Stability Data 41

2 Lateral-Directional Dynamic Stability Data 42

3 Summary - Critical Loads and Factors of Safety 47

4 Weight and Center of Gravity Calculations 49

5 Actual Weighing Results 50

6 Weight Adjustments 50

7 Wing and Platform Test Data 57

8 Keel Deflection Load Results 60

9 Drop Test Configuration No. 1 69

10 Drop Test Configuration No. 2 70

11 Drop Test Configuration No, 3 71

12 Drop Test Configuration No. 4 72

13 Drop Test Configuration No. 5 73

xi

________

I,

,,. - . ,,2I

LIST OF SYMBOLS

BHP brake horse power

cg center of gravity

CA axial force coefficient,

C drag coefficient, DDqS W

C drag coefficient at zero lift

C cycles to damp to 1/2 amplitude

8chCh hinge moment coefficient due to angle of attack, -"-

C K wing keel length, feet

CL lift coefficient, -

Cq rate of change ofclift coefficient with pitch velocity per radian

q

C rate of change of lift coefficient with elevator deflection d- LL 6dO'

S per degree

C rate of change of rolling moment coefficient with side slip

angle, d-C per radian or degree

MC pitching moment coefficient, - -

m

I xii

_ _ __I _ a r

dCmC rate of change of pitching moment with lift coefficient, -d L

L per degree

C pitching moment coefficient at zero lift

C pitching moment coefficient due to thrust

C rate of change of pitching moment coefficient with angle ofm dCmM attack, dC- , per degree

C rate of change of pitching moment coefficient with elevatordC

e deflection, dm , per degreed e

CN normal-force coefficient, NN

qS wdC n

c n rate of change of yawing moment with side slip angle,

a per radian or degree

C R specific range, number miles/pound of fuel

dCyC rate of change of side force with sideslip angle, _ I

13 per radian degree

D drag, pounds

F control stick, force, poundsS

force along X axis, poundsX

FS fuselage station, inches

GW gross weight, pounds

32 ftg acceleration of gravity, -2-sec

xiii

... .. ...- U

HM hinge moment, foot-pounds

HPexc excess horsepower, HP - HPexc vaii req'd

6. 1 X moment of inertia about X axis, slug-ft 2

,I¥ moment of inertia about Y axis, slug-fty

iz moment of inertia about Z axis, slug-ft 2

I X z product of inertia, slug-ft 2

iTH incidence of thrust line relative to body center line, degrees

iW incidence of wing relative to body center line, degrees

rolling moment, foot-pounds

length of control stick, feet

T distance from center of gravity to thrust line, feet

M pitching moment, foot-pounds

m mzss, slugs

N normal force, pounds

,r y side load factor

nz vertical load factor

2 lbq free stream dynamic pressure, 1/2 pv , t2

ftR/C rate of climb, feet/minute

i 8GP. takeoff or landing ground run, feet

xiv

- __-,4

SW flat-plan wing area, 450 square feet

St tail area, square feet

SFC specific fuel consumption

SR specific range, nautical miles

T thrust

THP thrust horsepower

tj time to dampen to 1/2 amplitude, seconds

V velocity, feet/second or knots

VLO lift off velocity

W weight, pounds

X distance from cg measured along X stability axis, negative insign if measured to point aft of cg

x w distance from cg to wing moment center along X stability axis

X distance from wing apex to longitudinal aerodynamic center of

ac wing measured aft along keel

X t tail moment arm measured along X axis, feet/second square

x acceleration along X axis, feet/second square

z distance from cg measured along Z stability axis, negative insign if measured to point above cg in feet

acceleration along Z axis

Zt tail moment arm measured along Z axis, feet

Zw distance from cg to wing moment cente,.- along Z stability axis

Xv

angle of attack, degrees

aOL angle of attack at zero lift, degrees

9angle of sideslip, radians or degrees

Esummation symbol

r dihedral angle, degrees

y flight path angle, degrees

6e elevator deflection, degrees or radians

68 control-stick deflection, degrees or ra(ans

E downwash angle, degrees

E angular displacement of X axis and about Y axis, radians

Mrolling-friction coefficient

braking-friction coefficient

P air density, slugs/feet 3

P

SUBSCRIPT SYMBOLS

B body

GR ground run

TD touchdown

TO takeoff

t tail

w wing

xvi

SUMMARY

Preliminary design studies and the final detail design of a FlexibleWing Light Utility Vehicle were carried out, and two test vehicles werefabricated. Structural adequacy was determined by stress analysisand appropriate static and dynamic tests. A limited flight test programwas conducted to determine the handling qualities and performance ofthe vehicle. Initial tests revealed a deficiency in longitudinal controlpower which was corrected by incorporation of an auxiliary elevator atthe aft end of the fuselage. Roll control power was satisfactory, androll control forces were light. A minor modification in the roll controllinkage further lightened these forces.

Positive static longitudinal stability was demonstrated within the rangeof center of gravity positions tested. It was determined, however, thatit would be advisable to move the wing forward 12 inches with respectto the cargo platform in order to improve the landing attitude withcenter of gravity forward.

The tricycle landing gear incorporating an oleo nose strut and fiber glasssprings at the main wheels has performed very well. The fiber glasssprings have required no maintenance.

Engine cooling by means of individual exhaust aspirated stacks provedto be entirely adequate as originally designed and fabricated. Theengine fuel system has operated with complete reliability. Enginestarting by manually turning the propeller has presented no difficultyat any time, starting generally being accomplished in the first quarterturn.

Additional flight testing will be required to fully evaluate the potentialcapabilities of the XV-BA vehicle.

.t1

CONCLUSIONS

From experience gained during the design, tfbrication and test programon the XV-8A aircraft, the following conclusions were made:

1. It is feasible to build a vehicle of the type under considerationthat will carry disposable load equal to its empty weight.

2. The performance of the fiber glass main landing gear springs inconjunction with the nose wheel oleo strut is excellent. Instill air, low-speed ground handling characteristics of theXV-8A are very good.

3. When taxiing at low speed in crosswinds of 5 knots or morewith the wing at high incidence, high lateral control forces arerequired to restrain the wing from canting downwind. Theseforces can be minimized by precanting the wing in the upwinddirection before turning to the crosswind direction.

4. The roll control system, incorporating movable tips at the aftends of the leading edge spars and a wing roll axis substantiallyparallel to the flight path, provides powerful control with lowpilot effort and little or no adverse yaw.

5. The wing should be moved forward 12 inches with respect tothe cargo platforn in order to improve the landing attitude withforward center of gravity.

6. Within the range of center of gravity positions tested, thevehicle is statically stable about all axes.

7. With the auxiliary horizontal control surface, longitudinalcontrol at landing approach speed is adequate. Control forces,however, are lower than desirable in comparison to roll con-trol forces.

8. A damping device as installed on the nose wheel is adequate toprevent shimmy.

2

9. Spanwise battens installed along the outboard portion oi thewing, trailing edge are effective in preventing trailing edgeflutter.

10. Improved longitudinal control results when the wing pitch trimsystem is interconnected with the pilot's control column so asto produce ±I-1/2 degrees of wing incidence chang2 when thecontrol column is moved through its full travel of *19 degrees.Incorporation of such a system would require that the mostadvantageous position of the wing pivot point be chosen, inorder to avoid excessive control force.

3 milli.

RECOMMENDATIONS

The following recommendations are made based on the experiencegained thus far from the XV-8A test program:

1. The wing should be moved forward 12 inches with respect tothe cargo platform, temporary outboard trailing edge battensand auxiliary horizontal tail surface should be replaced withpermanent units, and the aluminum alloy aileron hinges shouldbe replaced by steel parts to improve rigidity and reliability.

2. Tests should be made with progressively decreased freedomof motion of the wing about the roll axis with the objective ofeliminating this motion entirely if satisfactory roll controlcan be obtained by means of the ailerons alone.

3. All tests made thus far have been with the "two-control" sys-tem, rigged so as to produce no interaction between the rolland yaw systems. Further tests should be made to evaluatethe several degrees of interaction that may be rigged into thesystem, as well as the system in which the yaw control isindependently operated by meanb of the rudder pedals. Theobjective of these tests would be to determine the optimumcontrol system configuration for the intended vehicle mission.

4. Additional flight testing should be done to obtain quantitativeevaluation of flying qualities and performance.

4

.. . .. . . . . . , - . . .. . . . . .. . . - - - " ". " - " --. . . . . . "

i i i i i i i i U

DESCRIPTION

Figure 1 shows the vehicle in flight. Figure 2 shows the arrangementof the vehicle describad heirein. It consists basically of a cargo plat-form suspended below a Rogallo-type flexible wing. A pilot's seat andthe necessary flight controls are provided at the forward end of theplatform. An engine, a pusher propeller, and a V-tail are mounted atthe rear of the platform. Provision is made for manually folding thewing and tail surfaces.

Dimensions and Weights

Keel and leading edge length ............. 26 feetLeading edge sweep angle .............. 50 degreesCanopy area (flat) 45 degree sweep;

6 percent scallop ................... 450 square feetVehicle overall length ................. 26 feetVehicle overall height (wing horizontal) .... 14. 54 feetVehicle overall width (wing extended) ...... 33. 4 feetVehicle overall width (wing folded) ....... 8 feetPropeller diameter

(Hartzell two-blade metal) ............ 7 feetMain landing gear tread ................ 9 feetWheel base ....................... 10. 63 feetDesigned gross weight ................ 2300 poundsEmpty weight (actual) ................. 1115 pounds

STRUCTURAL DETAILS

Platform

The basic body structure is a flat deck with a raised platform at theforward end which supports the pilot's seat, nose wheel, control mech-anism, instrument panel and nose fairing (Figure 3). The usable cargoarea, 64 inches wide and 80 inches long, is fitted with twelve standardflush-type cargo tie-down rings. Transverse beams are incorporatedat the forward end of the cargo area, at the main landing gear, and atthe rear where the tail sur'ace loads are carried. Removable doors inthe bottom skin provide access to the fuel tank, piping, control pulleys,and cables. The fiber glass fairing at the forward end of the cockpit isremovable for access to the back of the pilot's instrument panel. Thepilot's seat, a built-in portion of the vehicle structure, is equipped with

5

a standard seat belt and shoulder harness. Space for a back-pack-type

parachute is provided.

Wing Support

A forward A-frame and an aft tripod constructed of aluminum alloytubing attaches the wing to the body structure (Figure 2).

Wing Keel

The wing keel is a tapered, sheet aluminum alloy box type structurewhich attaches to the roll control structure with a hinged fittinginstalled at the keel 46 percent station.

Wing Leading Edge

The wing leading edges are hollow aluminum alloy spars having asymmetrical streamlined cross section. The spars are tapered towardboth ends from a maximum section near the spreader bar attachment.The wing membrane is attached along the trailing edge. The aft 13-1/2percent of the leading edge is hinged to permit ±5-degree motion in achordwise direction. A cable and pulley system is used to control theposition of the hinged portion of the leading edge in flight.

Wing Spreader Bar

The wing spreader bar is a transverse truss work of steel and aluminumtubing attached to the wing keel and leading edges. It resists the inwardand upwatd forces due to membrane tension and transmits lift loads tothe wing support structure. By re-moving two quick-release pins at thejoint between the leading edges and the spreader bar, two men can foldthe spreader bar and bring the wing leading edges inboard.

Wing Membrane

Fabric for the wing membrane is square-weave Dacron cloth coated onboth sides with olive drab polyester resin (Figure 4). The treated fabrichas a tensile strength of not less than 200 pounds Per inch in the warp di-rection and not less than 120 pounds per inch in th - fill direction. Number6 machine screws, spaced 3 inches apart, are usea to attach the wing mem-brane to the aft edge of the leading edge and the keel. The screws passthrough a metal reinforcing strip, bonded and stitched into the hem.

6

A reinforcing cable, sewn in the hem along the trailing edge of the mem-

brane, is adjustable while on the ground for roll trim.

Tail Surfaces

The twin tails mounted at 35 degrees dihedral are shown in Figure 2.The movable surfaces incorporate aluminum alloy spars and fabric-covered ribs. The fixed surfaces are similarly constructed, except thatmetal skins are used to obtain maximum torsional stiffness. Hingedattachment fittings permit folding the tail portions up and inboard fortransportation and storage. As a result of preliminary tests, an auxil-iary horizontal control surface hinged to the aft end of the control plat-form was incorporated in the final design.

Landing Gear

The tricycle-type landing gear (Figure 2 and Figure 5) uses main andnose-wheel tires and wheels of the same size and type (7.00 x 6), in theinterest of keeping required spares to the minimum. The tread of themain landing gear is 9.0 feet; wheelbase is 10.63 feet. Landing loads atthe main wheels are absorbed by cantilever fiber glass springs extendingfrom both sides of the platform structure. Single-disc-type brakes inthe main wheels are hydraulically actuated by a master cylinder locatedin the pilot's cockpit. Landing loads on the nose wheel are absorbed byan oleo-type shock absorber incorporated in the nose landing gear. Thenose wheel may be steered through an angle of 25 degrees either side ofcenter by operating foot pedals In the pilot's cockpit.

PROPULSION SYSTEM

EngineThe engine spectfications are as follo,.vs:

Model No ................... IO-360AType ..................... 6-cylinder, Horizontally

Opposed, Air CooledContinuous Horsepower,

Rated Maximum ............ 195R. P. M., Maximum Continuous . . 2,800Fuel Octane Rating ........... 100/130Fuel Control System ......... Continuous Flow, Injector

7

I -.l - -.- - - --- -- f i ---- --- i

Four flexible rubber mounts are used to attach the engine to a steel tubetruss near the aft end of the platform structure. An exhaust aspiratedcooling system is used. Ten quarts of oil are carried in the enginesump. Oil flow through the air-cooled, engine-mounted oil cooler isregulated by a built-in thermal device.

Propeller

The BHC-C2YF-1A, 7-foot-diameter, all-metal propeller is installedas a pusher. The propeller is operated at fixed pitch by locking theblades at the angle for best all-around performance.

Fuel

A 25-gallon aluminum alloy fuel tank is mounted below the floor of thecargo platform near the center of gravity. (See Figure 6). A float-typefuel quantity gage is used. The tank is removable through a dooron the bottom of the platform. Fuel from the tank flows throughan emergency shutoff valve and fuel strainer to an engine-drivenvane-type pump which maintains a constant pressure of 8 p.s.i.at the inlet to the engine-driven fuel metering pump. The meteringpump supplies fuel to individual-cylinder fuel-injector nozzles atthe pressure required for the particular throttle setting and enginespeed. A hand pump, mounted on the right side of the pilot's seat,supplies about 2-1/2 p. s. i. pressure to the fuel metering pump forengine starting.

Control System

For ease of operation by pilots with minimam training, the AerialUtility Vehicle has been designed as a "two-control" aircraft.Longitudinal control is accomplished by means of the tail surfacespreviously described, actuated by fore and aft motion of a control"3olumn in the pilot's cockpit (Figure 7). Roll control results fromdisplacing the wing about the roll axis. This action is achieved partlyby direct control force applied to the wing and partly by means of servo

- I tabs built into the aft ends of the wing leading edge, both actuated bytmeans of a he ae t the top of the pilot's control colur. In orderto counteract the adverse yaw associated with wing roll, the controlmechanism incorporates means to actuate the two ruddervators differ-entially when roll control is applied. An adjustable stop is providedto limit the amount of up elevator control so as to avoid flying at wing

8

Q.

angles of attack beyond 34 degrees. A pitch trim wheel is providedon the left side of the pilot's cockpit. It regulates the incidence ofthe wing to effect longitudinal trim for any flight speed and center ofgravity position within the design limits. A pair of foot pedals fornose wheel steering and main wheel braking completes the cockpitcontrol installation.

Roll Control

To reduce pilot roll control force requirements, an aerodynamic boostsystem is incorporated. The aft tip portion of each wing leading edgespar is hinged to permit ±5 degrees motion in a chordwise direction.These tips are differentially actuated by the initial movement of thecontrol wheel to produce an effect similar to that of ailerons on a con-ventional wing. The rolling moment due to tip displacement is in adirection to assist the desired wing roll. When the tips reach the limitof their displacement, further motion of the control wheel follows therolling motion of the wing, or augments it, depending on the amount offorce applied.

Yaw Control

A mechanism under the floor of the pilot's cockpit interconnects theaileron control system with the ruddervator system to move the tailsurfaces differentially and thus produce a yawing moment. The linkbetween the aileron system and the ruddervator system can be con-nected in any one of several sets of holes to permit rigging of varyingamounts of yaw control into the system.

Longitudinal Control

A tubular steel control column is hinged to the floor of the pilot'scockpit. An arm on the bottom of the column extends below the floor,where it connects to the control interconnect mechanism describedunder "Control System, Yaw Control. " A system of cables, bellcranks, and push rods moves the tail surfaces up or down in unisonwhen the control column is moved fore and aft.

Longitudinal Trim

The vehicle may be trimmed longitudinally by turning the trim wheelon the left-hand side of the pilot's seat. This wheel rotates a cable

9

drum through a set of reduction gears. The revolving cable drumreels a cable out on one side while taking up the cable on the oppositeside. The cables are attached to the wing keel, one forward of the mainpivot and one aft. Forward rotation at the wheel rim produces a nose-down moment and vise versa. An irreversible clutch in the wheel hublocks the wheel against rotation when torque is applied from the direc-tion of the cable drum, but permits free rotation when torque is appliedat the handwheel.

Roll Trim

A small-diameter steel cable is incorporated in the trailing edge ofeach lobe of the wing canopy. Tightening the cable increases the liftof the wing lobe. The cables are independently adjustable to provideroll trim.

Yaw Trim

No provision is made for yaw trim.

Nose Stee ring Wheel

Foot pedals in the pilot's cockpit are connected by means of cables tosteering arms extending from each side of the nose wheel oleo pistontube. Stops at the pedals limit the steering motion to ±25 degrees.

Brakes

Both rear wheel brakes are operated from a single hydraulic mastercylinder connected to an auxiliary pedal mounted on the right steeringpedal (Figure 5 and Figure 7).

10

~- ..- *I -~--- UI -

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Figure 1. In-Flight View - Flexible Wing Aerial Utility Vehicle

11

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I . ... ... .- -- ~ - --, -

'i Ii , IH i I/i / i i I i f l 4 I + 1 kV D i i , 57 - - i l

'' F IIIi l 11111 / 1 hI ll , i4i SM--ct-:# (^, K, , In",) u s; - - - -- '

, : - -1 .- l /....a-",-. 61 LI YW TI Ace' .0

Iq I I. W oH11 r_ Ii

i I I I IjI

1 ..+ -- _. - -- ; II . . .+ .. . _1 .. . . . . -h i

t. ....

I _ 11I;2ii1 " - i x _I'L'I "

/ l lll 1 I 1/il' +." I re,,,p A-" .eaL i 4-- # (.lilill +- i ILI I

- _-_-_ _ _ __- ---- 4__-------"

AAIr I ) i"L-- A I9

Fi 3.11 Ae a iitI I iF~' , I7~r I L -Ii

A. it 4,.6~

At A%o-LL

III A,:,rAO

'51.cr-1N FD *st. S

1"w- 14.41 U 4 Or~ I

41.0-" 4L0t4 '.

. ! Aga0X1T0144.Q

Noro n. 1l~14 34

1AN I- 1 t t41 i tA41wk*'r' 11) 41v

BOLT W01414 I.5 NIL"(MA"9 All

Fr" orLE7 POVWAr

aT . k. Aq -D,

XNl AS QO r& 7

I wo VF.'jVl 4Jt

U 4 1A L b Q II O I O '0

- 4at CwA50 ~ 7A' a~

40500-AL~~~~~" 7AI A ,54 -416~ 4 AL 1 171.'NJAL

/r ~ I E o1'~b C

15A' T LQ-

-w -

i4.4."0I9"2/7!

" -. " " ..r OIi

, , ,..'( ., -.. .- -,,'- _!!

li 1 Li ,, 1-

"6Ii

.,/ -.. ,,

Figure 4., Wing Assembly and

17

- -I

'4w MT~( AM '' ! j

(~4/.qQ4*j-....

C, .

(~ ~i- g

Wi_~

/

/

~1

//

/7

/I /,

I .

/

.7/

/ I

N. ,

N /

Figure 4. Wing Assembly and Installation

17

/'1~

* ~

An *14'-''' -

.5:: F;:.] i (1 b A wlve

® 4'rrAIL-,- hPACE

,U a c, - - nu - sn e b- Rw e r

A.-Wo/

AM 4 to b g~ a S ~ ~ M L T ~ *.c a~ 4Fl,

-,L5AZqm

o..,'o- -124 CflJ.9k -

A~so-Vqt- ~ -- K

.. ' ,-.V

%,,

tII 4 !l

-I- 4--''.,4.4-.-, 4-"-

-" ("flrre

• ij' 4j'ix.

I . Ii

-- '"4-,--,,,- -.

-3 -- <' - #V ' '. " rt

":4. '.5 *,.1 ffs

:,:, .. ....... .... iFigure 51

__ _ _ I-

---- --

*11

I i "I 2

Ii I

jlI i 1

. . .. ,. -t 1, I

I" I .... ) .... .... .

-/r-*a s. .-.

Ii

* U-AF,- . el5' Main LandflniCGar and Brake Assembly

0.0 or FILLL~r N*CA( (Z.31 Mgt)2c, t Mnosf r

IIe

tt~~r StAL 10!.' WIT.- ,t a.NCz

DCEVCCN'F PVTY -FL'0* .

. 0.

-- -60 E l~( -- - a-- 'o '

~ 40

L W -' 0s . lij

M 177 rtju- W-.A .

,4",*U Z;S ovr m /A& 01.

Vf SLOJf

(1) 4 '1

A,~r (j)

-77I I ~ T Is

= £..IL

~A Iva"~

[ II N !1 -. wi

, -:-l-:... .-

'' z Z I ,,.....

-" -- - --ANW90-O . .I, -- -

I I\ -

......... .,..,

Irt;2W - -

i' , ,' ," L -

.vi- _-iur 6.~ FulSse Isalto

(AM TMC 144f 4 rte ~C~,JIt fdmf 21

.2. Ov,~n,

,,

.

BAMWI

~~~IW .- Wa W )h~

'-11... 1- - - - -

_ . / I'. ..-a#Cw 4r

\VZjb <XA k . - ,s is'-'ll~ll \L- l ' " --

(ha.\ , , r.-~a4 a ~b' '- a . ...... .

Figure B. Fuel System Installation I

17" I I -- I 1 I - - . . ... r~

/1217-2~

I -

WIN,\I "O Ala , ]XC .MUn/ S AVM oL

9"11A ~ -1k CAT11 -- * I

AN L~2 90- ~ APD5 .W f lW 41 wa M1/o //Wku n-IrKV

Nei U0)I (r II'. A ',I ( !V J)IJHMS~~~~~ ~ ~ /44 .6'M~)CItAS.tt E MW

AIi PD546WAW

feWOM~~~ WO4,- AirWN f'fWW

-,4z

IIt

9W C.. C44V14U .1151A,41,li~ .4III

?'TN~ I 4 ,I.4CO17. lt Ukl' (14MFI.A$4.? AJ1.S.4M1,S

OM M kM lUI WO4S vcf

1-111 --l 11 AI! -11--r - c

WI 041 eagi AW, I'doc A 0rff/avirM!

MAWO IIIA' 'itt W.;MdL . 74 4.,s

A64OJI AVIYJ/'VAFX IV.lt( (PUW U ASPI AkU AT41LA' A4CO09 CNFUhkc WW Oglha

/44'W 12 140/ A. 11 44 0 1111 1AA.74

Figure 7. Controls Installation

23

AERODYNAMICS ANALYSIS

PERFORMANCE

Methods of Analysis and References

This section summarizes the performance capabilities and thestability-and-contr0ol characteristics of the Ryan Model 164 Utilityaircraft. Includedarc the lift and drag bases and the methods ofanalyses used. All data are estimated data calculated prior to flighttest of the aircrft.

Drag and Lift

Wing drag and lift were obtained from Ryan and NASA wind-tunnel testdata. Drag of the body, struts, and other untested components wascalculated using standard estimating techniques.

The drag polar for the complete aircraft is presented in Figure BA andthe corresponding lift curve in Figure 8B. Maximum lift-to-drag ratioof the aircraft is estimated to be 3. 93.

Power Required and Available

Powur available was calculated using the engine manufacturer'sspecifications. NASA propeller charts for a two-bladed fixed-pitchpropeller of activity factor 90 were used to determine propellerefficiencies. Figure 8C shows power required and available at sealevel, standard day conditions.

Takeoff

Ground roll and total distance over a 50-foot obstacle are plotted inFigure 8D, as a function of gross weight. Takeoff distances werecalculated using the following equation:

2g (0. 0443) (W) (V1 )

* ~gro)und rol)l1

25

50Air distance over 50-foot obstacle -

tan

where y = arc sin RICV

Rate of Climb

Rates of climb at sea level are plotted versus velocity in Figure 8E forgross weights of 1, 300, 1, 800, and 2, 300 pounds. Maximum rates ofclimb are 690 and 300 feet per minute at gross weights of 1, 800 and2,300 pounds respectively. Rates of climb were determined as follows:

i/C (1li1 available - lip required) (33, 000)W

Range

Specific range is plotted in Figure 8F, as a function of velocity, forsea level, standard day conditions. Maximum specific range is 0. 712and 0. 818 nautical mile per pound of fuel at gross weights of 1, 800and 2, 300 pounds respectively. Specific range in nautical miles perpound of fuel was determined from the following equation:

Velocity __ V knotsFuel Flow SFC x Bill-

Landing

The landing performance of the aircraft has been calculated in twosegnents: (1) distance from 50-foot obstacle to touchdown, and (2)ground roll distanCe. The distance from a 50-foot obstacle will varygreatly With approach sl)(Wd and pilot technique. Landings simulatedOn the analog computer indicated air distances of 300 feet. Optimizingthe approach technique will reduce this value; however, 300 feet willbe assumed to be representative until additional data are obtained fromflight tests. The estimated ground roll distance is plotted versus weightin Figure 8G, showing roll distance from approximately 70 to 120 feet

o a br.a., -n o .cent. o 0 5. -tol distance was calculated with thefollowing equation:

20. 043 W V

S )TD -(q at 0. 7 V

Ca cD- L /(HiAWT

26

I -VJI TiI ---- ' I I --]Iu -i-'"

.5 -- I _ __

MAX 1.28 1.4.46 - BASED ON RYAN AND

NASA W!ND

.4 _.2 TUNNEL TESTS..42 SING 40 FT2BASHI) ON FLATPLAN

LID MAX - 3.93 WN RA

.34 -

.30 - - --

.26 0

oF-..22 4 -__ _ _- -- *. - _ __ __ __

.14 _

.2 .4 .6 .8 1.0 1.2 1.4 0 10 20 30 40 50

KEEL ANGLE OF ATTACK (IEG REES)

A DiuAio ivmit B LIFT - COEFFICIENT VS ANGLE OF ATTACK

PERICENT SEA LEVE -

MAX CONT STD DAYPWlt AVAIL. 100%1600 -

120-WT. L13

(L MAX 2:100 1)X 41 --- - .

100--- --- ] - ~1200---------

-WT, LBI '7%1800

o 60 go.-...

0 ,'r, 19o

400

SEA LEVEl.20 -- 2

8T1JI)DAY TFMP

01L _p II- 0 i20 a() 40 W,') 60 70 1200 1400 1600 1800 20C0 2200 240 )

C HOISEPOWER REQUIIED & AVAILALLE D TAKE-OFF PFIR'FORMANCE

Figure 8. Performance Data (Sheet 1 of 2)

27

. .. ..... ... ~ ~ ............ -- =' -=

SEA LEVEL SEA LEVELSTD DAY

1400 | ! ___I

GROSS WT LBS

1200DO - 1300 1.0 ./ -.

ri 100 -1000 . 01G0OS WT LBS

1000

s 1000 OF

S 400 94 .6 t

200 .5"

U .4120 30 40 Go 60 20 30 40 50 60

TRIUE AIRSHPEEDJ (ICNOTS) VELOCITY (r NOTS)

E HATYI OF CImn vsJ VEiLOCITY F spE-CIFIC RIANGE

SEA LEVEL SEA LEVELSTO DAY SITI) DAYSW - 4bU

00

200 --1-

P -- 0

20 0 '0 0 00 0 30 40 o0 60

Tu..: ll : AWIEI~iT (LB40) VHh ELOIT (LIOTS

E IiA')(1 tNf VS VEOiY I 8CIC IIAG i___i_ I_____

- 1.

06

jJ.0

100 -40 IU 80 -0 20 20 00 20 00 40GROS WEIHT JM)01108 WEGHT 1,M

G AIN GROUN ROL H----PEE

Fiur 8.Promalc aa4lic f2

-4 _2_

Stall Speed

Stall speeds are presented in Figure 8H for sea level, standard dayconditions. Stall speed at the designed gross weight of 2, 300 pounds is34. 3 knots. Stall speeds were calculated as follows (stall velocitiesin knots):

f W cos y (295)

stall [C L raxy SW

STABILITY AND CONTROL

Methods of Analysis and References

The stability and control section includes an analysis of the longitudinaltrim requirements, the estimated stability and control characteristics,and results of an analytical investigation of power-off emergencylandings.

Wing Longitudinal Characteristics

The wing longitudinal characteristics were estimated from Ryan andNASA wind-tunnel tests of Flexible Wing models and from data obtainedfrom the Ryan Flexible Wing test bed. Characteristics of untested com-ponents were estimated using standard estimating techniques. Liftcoefficients for the wing alone are mesented in Figure 9A as a functionof angle of attack. The wing drag polar is shown in Figure 9B.

Tail Longitudinal Characteristics

Studies have indicated that a tail mounted directly behind the propellerand designed to provide adequate longitudinal control in the power-offcondition could easily pitch the aircraft to the stall angle of attack whenpower is on. This is possible because of the large inci'ease in elevatoreffectiveness resulting from the increased dynamic pressure at the tailcreated by ihe propeller slipstream.

A 51-square-foot V-tail has been selected as the longitudinal-controldevice. This cof.uratIon allows the tail surfaces to be mounted clearof the prz;1b,!!,or slipstream. Elevator effectiveness is thus made in-deplnd:it o f wOW," effects.

29

,dfI

1.4 iWING ONLYWING ONLY

INCLUDES DRAG OF STREAMLINEDBASED ON RYAN AND SPREADER BAR (NO TRUSS)

1.2 NASA WINDTUNNEL TESTS BASED ON HYAN AND NASA

BASED ON FLATPLAN WIND TUNNEL TESTS

1.0 WING AREA BASED ON FLATPLAN WING AREA

.. .( ,----------------1.0

.4~~~ 1_ .4 __ -0

0 __.0

0 10 20 30 40 0 0 ,10 .20 .30 .40 .50

KEEL ANGLE OF ATTACK (DEGREES) WING DRAG - COEFFICIENT, CDw

A WING LIFT COEFFICIENT B LIFT CURVE

STICK FORCE BLOWN 13 WIYII BOTH PANELSDEFLECTED FOR PITCH ZERO FUS STA IS 5 INCHES AFT

NO SIDESLIP OF COCKPIT NOSEPROP SLIPSTREAM EFFECTS NOT INCLUDED ZERO ELEVATOR-STICK FORCESEA LEVEL, STD DAY

300 ___ ---- 32 - ___ ___ ___ CO FUS STA

2 00 2 - 103

AMiRPEED (KNOTS) 2,

i.~ 10 __109

40 -- U

w 30

_ _ _ _ _ _16 _

-30 -20 -10 0 10 20 14 is 22 26 30 34

RUDDERVATOR DEFLECTiON (DE"JIEES) WiND ANULk Ot ATTACK, aW (LJ)IE D)

C ESTIMATED ELEVATOR STICK FORCED TRIMED FLtHT

Figure 9. Stability and Control Data (Sheet I of 5)

30

GROSS WEIGHT 2300 LBSEA LEVEL, STD DAY ZERO FUS STA 18 5 INCHES

AFT OF COCKPIT' NOSE.ZERO ELEVATOR STICK FORCE ZERO) ELEVATOR 8TICK FORCE

?EIR, FUS STA IS 5 INCHUS AFT OFCOCKPIT NOSE[Z 32 32

CG FUS STA f70 USSA91 a 103

24109, __ 24

', ,MAX POWER

14 18 22 26 30 34 14 i8 22 26 30 34

WING ANGLE OF ATTACK, aW (DEGHJFES) WING ANGLE OF ATTACK, uW (DEGREES)

E WING INCIDUENCE ANGLE, CIRUIlUOG F WING INCIDENCE ANGLE, MAX POWER

SKA LEVEL, STI) DAY AIRCRAFT TRIMMED WITH FULL POWERTHRUST = DRAG PRIOR TO ENGINE FAILURE.

J,, S1HOWN 15 4 REQUIRED TO MAINTAINORIGINAL TRIM SPEED.

6 3838,__ __ VERTICAL CO LOCATION AT FUS WL 36.6

I 3) 234

GROSS WEIGHT (LB)

2300__ __ _______

R' 30 .... +40

26 '__-30

w 22 +20

0C FUS STA

L? 1__--- .10w .- 109-103

0 .? r7141___ 1 ....... ,. 1 .J L ..... .1-

30 40 s0 60 70 80 30 40 50 60 70 80

TRUE AIRSPEED (KNO TB) w TRUE AIRSPEED (KNOTS)

G TRIM WINW - ANGLE OF ATTACK VS VELOCITY H ELEVATOR DEFLECTION, ENGINE FAILURE

Figure 9. Stability and Control Data (Sheet 2 of 5)

31

V

iw 18. ,..TOUCHIDOWN

ZERO STICK FORICE AT TRIM. a 3 -TOCIWN 4

STICK FIXED. 20 +-' I I tJ W 20: ,3n ~ ~ "-oJ ! J,, I J

GROSS WEIGHT 2300 LSB!

.04 20-& 40

.02 i____ 10__

CG F-l -ITA IJ JCJ

-. 2- ___ UFU T TIME-.- TIME-.

~-10 -

-103 POWER-OFF TOUCHDOWN019 AI'IPIIOACII VELOCITY IRATE OF DESCENT

H 0T-- -20 -q 20 T/ -

-Jr7H 1100

o- -4 - 10 4010170 q

TRUE AIItM'EED tKNOTS) " -10) 0 100 200 300 400.hORIZONTAL DISTANCE (FEET)I LONG TU]INAL 1TATIC MARGIN POWER OFF TOWER-OFF lANDING

--- MODEL 1414 $qTICK FLXED....lYAN MANNEDI 'TE'iT lED AFT (XI (Cli FURS lTA -109)

IN LANa LY TUNNEL UROSS WEJUIIT 2 ,300 LIITAALITY AX RYOCTEM VELOCTA TEVl OTF DCAY

100 ,002 -K2 -Q F T' E

V -Ii.. . .

-.004 O

.061.0

..

s 'o 50 d L i7 II

0 RU 10 mE 30NS 400 00 4 0 200 00 700

WINO ANULE O ATTACKNTiW (REEhi .. TRUE A(C(;1'EED (KNO T )

IN TATE LA L-DIIECTIONA L r2S TATIC TABII,ITY LONSYITUMINA, DYNAMIC LV TAITLITY

Figu, re 9. Stability anid Control Data (Sheet 3 of 5)

12-

4=

STICK FIXED STICK FIXED- - AFT CG (CO FUS STAI09) AFT CO (CG FUS BTA-I09)

GROSS WEIG HT=2300 LB GROSS WEIGHT-2300 LBSEA LEVEL , STD DAY SEA LEVEL, STD DAY

.4 .

I::1 .. 330 40 60 60 70 30 40 50 60 70

TRUE AX ISPEED, (KN3TS) TRUE AIRSPEED, (KNOTS)

. ,w.1-------------- --------

_________ 0 .15

N NN

.2 .1 ---

30 40 50 60 70 30 40 50 60 70

TRUE AIRSPEED (KNOTS) TRUE AIRSPEED (KNOTS)

M LONGITUDINAL DYNAMIC STABILITY N LONGITUDINAL DYNAMIC STABILITY

STICK FIXE)

AFT IW (CI I'US 8TA .109)O ItC)X WEC1IHT-2300 1,i

- - - - - - - - - - - - - ------- SEA LEVEL. 8'11) DAY

ROLL MODE

................ ... .. ..- - STICK FIXELDAFT (XI (CU FUS STA-1OB).... GitusS WEIIIT300 IH

-.- B EA LEVEL, ST) DAY

SIRAIL MODIE

+ El 7Jr73n:0- __-

30 50 1 7TRUE AfISPEED, (KNOTS)

4 rI- _1-i-;-o,--1 20___ ,

DUTCh RHOLL 12 10

U030 40 BU0 60 70 30 40050 00 70

TRUE AIRSPEED (KNOTS) TRUE AIRSPEED (KNOTIS)

10 LATERAL-DIRYCTIONAI. DYNAMIC STABILITY P I.ATEIIAL DYNAMIC STABILITY

Figure 9. Stability and Control Data (Sheet 4 of 5)

33

AFT 0.0G.

OPERATING POINT:

a.V - 40 KNOTSOPERATING POINT b. FULL POWER

4TAKEo. -7PIRAL-DIVEIIL3NCX TIME

.3- TO DOUBLE AMPLITUDE-13 SECId. GROSS WEIGHT 3300M LS

SPIAL STABILMTBOUNDARY

OSCILLATORY STABILITY

BIOUNDARY

CgPER RADIAN (UC OL

0 LATERAL-DIRECTIONAL STABILITY BOUNDARIES

Figure 9. Stability and Control Data (Sheet 5 of 5)

34

The following criteria were considered in selecting the area anddihedral angle of the V-tail:

1. Maximum wing angle of attack with full up-elevator should notexceed the angle for encountering pitch-up or stall.

2. The tail should be capable of increasing angle of attack by atleast 7 degrees with maximum elevator deflection.

3. Tail size cannot exceed space and weight limitations.

A longitudinal 3-degrees-of-freedom analog simulation was performedto evaluate various tail sizes and dihedral angles in order to select avee-tail that would best satisfy the above listed criteria. Evaluationof the analog output resulted in selecting the following fail geometry:

Total tail area ........................ 51 square feetMovable area (including balance area) . ... 34. 8 square feetHorn balance area .................... 4. 72 square feetFixed area .......................... 16. 2 square feetGeometric aspect ratio ............... 1. 6 (1 side)Airfoil ............................. NASA 0009

The equations of motion used in the longitudinal simulation werewritten in the wind-axis system as follows:

.. .2 2 F zL1/-

S-[ C 1 (W1 iw C K )(SWM 1 W 01 (YB 1 q W, 13 _

F B L 1C 6 C1 Z)]nL t C t

JT

- sin ((v - i g cos 'Y V

EX 2 W)2

[ C,)OW KD (", 4 iW) + K (D D D2 ((B'W

3 2 'I1) W D) D LS

01it, t Bt

I- cos (( t iT) - g soin Y V

rnB 'I'll

35

___-_ .. 1'..I-

Z M X W Z Wom"y:- - C +C - -- 4-C +Cm aB +.i0w NW CK A K B aB

O1 C K) Xt Stj qSC K TZc + C rnqw , B ( VLt C K IYY

C C +K ( + i K (( i) + K (YB.iw)3Dw D 1 B W ) 2B W 1) 3 B WW 1 2 3

where Xw and ZW are distances from aircraft center of gravity to wingaerodynamical center, measured parallel and perpendicular to keel,respectively. KDi , ,1.)2 , and 1I7.1) are coefficients to curve-fit equationfor c D1 2 3

Elevator Stick Force

The elevator stick forces were determined as follows:

Work done at the stick = work done at the elevator

F1S C6S If M 6

2 2

where is length of stick in feet, and 6S and 6C are the angularrotations of the stick and elevator, respectively, in radians. Theabove equation reduces to:

6F S IS -;s s

Figure 9C presents the estimated elevator stick forces as a function ofcontrol surface deflection, for several airspeeds.

Longitudinal Trim

The wing-incidence angles requirc-d to trim the aircraft with zerostick-force are presented in Figures 9D through 9F for maximumpower, cruise power, and power-off conditions. Trim angles ofattack are plotted versus velocity in Figure 9G. The trim points wereobtained from a longitudinal-trim digital computer program which

36

yields pitching moment coefficient versus angle of attack for a range ofincidence angles.

The longitudinal trim program utilizes the following equation to solvethe pitching moment coefficients about the aircraft center of gravity.

xw zwC =C +C X C - - +

m eg m 0 W NW C K AW CK flIO

cg 0W 0B1wB K Y eP

Longitudinal parameters pertinent to the trim calculations are listed below:

C = -0. 028

C keel length (26 feet)K

C -0. 0018 per degreeOB

C 1 0. 00021 per degree

(VII

I'l - 0. 048CK

x St tC -C C

in I C K 8(It a t W

t w - W - 07 vW - IW

Elevator Deflection Required To Retrim in Event of Engine Failure

In order to .,ovide clearance beween the propeller and the cargo bed,it was necessary to mount the engin. well above the bed. This highthrust line results in a nose-up moment in the event of sudden loss of

power. The elevator deflection required to retrim to the original trimspeed was determined as follows:

37

n n n u u i n un n u

T ZtC-c (t Se)ImT qs CK In6

where

x tC =Cm C L CK

6 6 K

TI ZT t StCnT qS CK L 5CK (Ae)

0 0

where

'1' ZT

e5 q C 1, X S t(I 1 L S t6

Total elevator deflection to retrini is then equal to:

66 A 6T1RIM FIOAT

Figure 9H1 presents the elevator deflections required to retrini theaircraft, in the event of a complete loss of power, as a function ofairspeed and for three center of gravity iocations.

Stick-Fixed Longitudinal Static Stability

The stick-fixed longitudinal static rn gin was calculated from thefollowing relations:

CMC C a

CL

Values of C were obtained from the digital trim program discussedm

previously. C1 was determined as follows:

38

I . i---nn--- n -e - n~ nI

C =C L C1c L L L

a aW a t (V 1

C = 0. 0422 -4 0. 0020 - 0. 0016 0 0. 0458 per degreeLa'

C is plotted in Figure 91 as a function of xth lift coefficient and of

speed, for three center of gravity locations.

Stick-Free Longitudinal Static Stability

A quantitative estimate of the stick-free longitudinal stability has notbeen made; howevei, the longitudinal static margin is expected to beslightly greater with controls free than with controls fixed. This in-crease in stability is a result of the positive ( which induces the

(-V

elcvator to float downwa id with increased angle of attack. The mag-nitude of the stability increase will, ot course, be reduced by frictionin the control system; howver, the stick-free stability will never beless than that with controls fixed as long :is (I is positive.

Power -Off Emerg ency Landings

Standard landing procedure for the Model 164 will call for partial powerthroughout the approach for the following reasons:

1. Power (icn be used to establish the desired rate of descentduring the approach.

2. Landing flares are accomplished with lower approach speedsdue to the in're,_e(d effective lift -t -d rag ratio resulting fromthe added thrust.

3. Low touchdown rates of descent are obtained.

Conversely, power-off landings require higher approach speeds andrates of descent in order to insure a satisfactory flare. These approachconditions make it difficult to judge altitude and flight path, and forthese reasons, power-off landings should be considered as emei , encyprocedures unless subsequent flight tests prove the maneuver to beless critical than predicted. Power-off landings were investigated

39

i I II - -i I - I i - - - -

during the analog simulation discussed previously. Figure 9J presentsan analog trace of a typical power-off landing. The minimum touchdownrate of descent of 8 feet per second was obtained by initiating flare at aheight of 50 feet. Full elevator deflection of 25 degrees was used.Approach speed was 55 knots. Studies have also shown that power-offlanding velocities may be reduced by following the procedure listedbelow:

1. Establish glide speed as quickly as possible by use of longi-tudinal control.

2. Roll pitch-trim control back while maintaining glide speed bymeans of longitudinal control (stick forward).

3. Apply maximum up-chv~ator to flare.

The elevator deflection available to flare the aircraft is thus increasedby the amount of deflection held during the approach.

Lateral-Directional Static Stability

The lateral-directional parameters C , C and c for the complete

aircraft are presented in Figure OK. These data are based on Ryan andNASA wind-tunnel tests of Flexible Wing xngdels, and on flight test dataobtained with the Ryan Manned Test Bed. The following equations definingvee-tail lateral-directional stability were used to obtain estimates ofC , C and C due to the tail. The derivation of these equationscY , n I'

can be found in reference 8.

'2C - K C sin 1'

(VN

C C Y (1) it fit

t tCy -, -i

fit pt h

where K is a constant for computing slope of lift curve of a vee-tailin yaw (equal to 0. 67 for Model 164 tail).

40

..... . . .. . .. . . .,. .. . . .-...-

Longitudinal Dynamic Stability

Longitudinal dynamic stability was .-alculated by 113M 704 using Stick-fixed, s~mall perturbation equations. The data listed in Table 1 havebeen plotted as a function of velocity and are presenied in this sectionfor evaluation of the dynamic longitudinal characteristics.

TA13L E I_______LONGITUDINAL. DYNAMIC STABILITY DATA

FlightReference Mode Characteristic Conditions

Fig. 9L P1hugoid Time to 1/2 anipl. Aft (,enter ofFigt. 9 L Phugoid Cyc tes to 1/2 ampl. gravity (F. S.Fig. 9M Phugoid D~amped [req. 1 09), grossFig. 9M Phugoid Daminug ratio weight 2300 11).)Fig. 9N short period Tinie to 1/2 ainpl. "ea level,Fig. ON Short period Cycles to 1/2 ampl. std. day

The Smlall pe rturba ~t io ei(quadtions used!c in the a naiys is arie based onlequations dieve lopeii(i i BA El? R~eport. ANF-61 -4 Vo In mu II1, "Dynamiic'sof the Airflra me. "' The following a ssuniptions werie miade:

1. Thu airframle is; a rigid body.

2. '1'het XY p~ila f i it a1p1,111V of Sylllllet ry.

3. IDurinug steady statle flight the( airt ranlic is flying With w i il;'sleVel, all components of velocity other thani ui ar- zn,tile !tit'iranlf! is ill icelcie fligh"t. A O an

4. Only small I pe i thatiP us from (ht le steady ]'light coniditionis aricperittied.

''he( equat ions arc:

Xi 111 I X XW IXoX 6 6 If, (cos -Y)

Xv z IJ j ~ V I -W, (si W )1 Z11 W XV 0 0

41

OM U4 MW4 M W M q6+M 6 ee

The symbol definitions are too lengthy to be reported herein. If re-quired, these definitions may be found in Reference 5. The dynamicstability investigations indicate stable longitudinal characteristicsthroughout the speed range.

Lateral -Directional Dynamic Stability

The methods used to estimate lateral-directional dynamic stability aresimilar to those utilized for longitudinal dynamics, i. e. , small per-turbation equations solved by IBM 704. The stick-fixed lateral-directional perturbation equations, based on equations developed in

Reference 5, are:

Y J1 4Y Y' .1 4 6 - iI cH Y (si-'

r 1 , 0 0

0- Lfi L AL +L 6 L 6 '-i/i[1 P r 6 a a 6 r ' 11& V IX

N N 1 4 N , -i N 6 ( V N -r A- I "

(Y t' Z

See reference 5 for symbol definitions, if required.

Table 2 lists the data obtained from the lateral-directional perturbationprogram. These data are plotted versus velocity in Figure 9, details Nand 0.

TABLE 2LATERAL -DIRE CTIONA L DYNAMIC STABILITY DATA

FlightReference Mode Characteristic Conditions

...... 90 Du rI/t, o 1/2 ampi. Aft cg (F. S.

Fig. 90 Dutch roll l/cycle to 1/2 ampl. 109), GW 2300Fig. 90 Spiral l/time to 1/2 ampl. lb. , sea level,Fig. 9P Roll l/time to 1/2 ampl. std. day

42

A ~- . . . . .I

The above dta indicate stable dynamic stability in the dutch-roll andpure-rollmoaes, and mild instability in the spiral mode. The degreeof spiral divergence, however, is within the acceptable limits formanned aircraft. In addition to the above data, the lateral-directionalstability boundaries have been determined in the takeoff condition.These boundaries are determined from the aircraft's characteristicequation of motion,which has the form

A 4 + B 3 +CX 2 +DX+E =0

The conditions of neutral dynamic stability are:

E = 02 (Spiral boundary)

BCD - AD 2 - B2 E = 0 (lutch roll boundary)

A complete discussion of the method of analysis can be found inReference 7, Chapter 11. Solutions of the above equations were ob-tained by IBM 704 and are plotted in Figure 9Q in the - Cn

plane. Figure 9Q shows that the aircraft will be operating far fromdutch-roll instability, but that mild spiral instability will be present.The time to diverge to double amplitude was calculated to be 13 seconds,which is within the acceptable limits for manned aircraft.

43

43txwtMt

STRESS ANALYSIS

STRUCTURAL DESIGN CRITERIA

Design Weights

Minimum flying weight .............. 1189 poundsBasic flight design gross weight ....... 2300 poundsMaximum flight design gross weight . . . 2300 pounds

Design Speeds

Level flight maximum speed (VH) ...... 56 knotsLimit speed (VL) ................ 84 knotsSpeed for maximum gust intensity (VG).. 50. 5 knots

Ultimate Factor of Safety

The ultimate factor of safety is 1. 5.

Maneuvering Load Factor

The maneuvering load factor is +2. 5. See Figure 10.

Gust Load Factor. See Figure 10.

Vertical nz = +2.25Side ny = 0. 25

Design Landing Load Factor at Center of Gravity

The design landing load factor at the center of gravity is 4. 25.

Design Sink "p ed

The design sink speed is 12. 0 feet per second.

STRUCTURAL LOADS

Wing Keel and Leading Edges

Distribution of air loads along the wing keel and leading edges wasmade from test data set forth in NACA T. N. D-983. See Table 3 forsummary.

44

_ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _

Landing Gear

The landing gear and its attachments were designed to absorb the shockresulting from a descent at 12 feet per second, assuming wing lift to be2/3 the gross weight. See Table 3 for summary.

Tail

Resultant tail loads were determined from Ryan analog simulation data.Chordwise distribution of air loads was made in accordance withCAM-3. Spanwise distribution was in accordance with NACA T. R. 921and NACA W.R. L-212. See Table 3 for summary.

Engine Mount

The engine mount was designed to take the vertical, fore and aft, andside loads resulting from maneuvering, landing, gusts and propellerthrusts: it was also designed to take the moments due to engine torqueand gyroscopic forces. See Table 3 for summary.

Cargo Platform and Body Structure

The cargo platform and body structure were designed to resist the loadsimposed by the landing gear, engine mount structure, wing supportstruts, cargo load, control system and tail surfaces. See Table 3 forsummary.

STRUCTURAL ANALYSIS

The capability of the various structural members to resist the imposedleads was demonstrated by zalculations conforming to accepted engineer-ing practice. Static proof loads representing the most severe designconditions were imposed on actual wing and tail structures to verify thecalculations. Drop tests on the landing gear were made to demonstratethe ability of the shock absorbing system to dissipate landing shockloads. The margins of safety of all important structural membersunder the critical load conditions are summarized in Table 3.

WEIGHT DATA

Weight and Center of Gravity Calculations

The weight and center of gravity calculations are presented in Table 4using reference data shown in Figure 11.

45

Loading ,.nd Burnoff

The XV-8A loading and fuei burnoff results are shown int Figure 12.

Results of Actual Weighing

The results of actual weighing are listed in Table 5. These results are

based on the following:I

Pilot ballast .. .. .. .. .. .. .. .0. 0 poundFuel. .. .. .. .. .. .. .. ... ... 0. 0 poundOil. .. .. .. .. .. .. .. ... ... 15. 0 poundsinstrumentation wiring. .. .. .. 15. 0 poundsDeck angle. .. .. .. .. .. .. .... 0. 0 degree

VERTICAL OUS3T 26 FPSi\.

MANE UVEHMN(

LOAD FACTORI___

0 0 40 60 80 100

G ROW WEIGHT - 2:300 LII

Figure 10. V-n Diagram for Symmetrical Flight

46

0

Q Cq -4 LO 0D w -4 0 J0 C m 0C c C 1t L- ClcoCo com

~Cd

4 r i Z)0 0o*

44 J ,-44 C

0l 4- 00.: o0 v DC CD E -0 DC ) t-D -4 4

O0 0 ~J ~ 000v-4

Q~~ u u Q U U D S-

0i -dMb. nU a0U-3 m M Cd m M

- C %.C

O V.Ci 0 1" 1- r-4 -4 a) b bD bD -1 44 +44

1- a 0D U) (p I-b UC b

P44

uB

*l -4

L-. -Cq -4 m 0 * 0c 4-WV c 1to 0 D 0 t CDlC mC mr m O E- C

Cd (L .~ . .

0) 0

C) C)C : :) 0C )C)C D 0 0InC' d : C)00 0 00000 0 00(: c )i

o mtoC0 0 L- 0 -L-t- 0 t- C) C=

U') 0q toi.4L U)

004 ( 04m 0 Lt3 - C)0t 1- 00(0)t oL- t C') CD -4 00 m 0 In4- Cd

a M 1- IN 1d Lt- 0c~ 0C fl C c 00Lm( 0

0 q o - .4-

4-'Cd 0

17 ) 4..- C -

S 40 0 QS w0j.= ' dcdc ~ nC

u- '-b Z i Z C E-l bCbJ C' Cd41

Cd (D c 'S

Q) -- '

'-4-D 0 m 44 ~ 4

V) ~ ~ U U)a 0 0(

0 0 M) 0u m Q) Cd

' . bDO a

0 C '- -0 0 UC

0 )u 0 ) rCid a)l Cs

-a) (U 0.0.a) 0 0)'. b~ )bb Ma*0~~' 0~c4. ~

C) bl bb -0 w Z, C., M o bl C C

0. 1--S 0 C t r ::5 0 * _- *E_ H~U L)P P

48

M~ ~ -40-4C l " MMMt O0ML

M L-toLML-icet-0 -

C1-O 4 U') co l,-0

C1 -4 to00toto oV ---4C% W0 V

( ~ ~ ~ - -- 4 .

0 0

-4 -4- -4 v-4 N m -- 4 -400c

4.J.

-4 -1 -4 - -

m V0 4 -4 CDC)(DCD0 )

ZH'M - q - >(>C

C1 ; -414 c -1 0

U)

cd~~. b,0 40w4C

~$.0 0

0 0 CI U)rio0

49

TABLE 5ACTUAL WEIGHING RESULTS

Scale NetReading Tare Weight Arm Moment

Reaction (lb.) (lb.) (lb.) (in.) (in. -lb.)

L/H main gear 535 6.0 - -

R/H main gear 520 0.0 - - -

Subtotal (both main) 1,055 6.0 1,049 129.19 135,520Nose gear .02 0. 102 1.75 179Total 1,157 6.0 1,151 117.9 135,699

Weight Adjustments

Weight adjustments to configure to final assembly are listed in Table 6.

TABLE 6WEIGHT ADJUSTMENTS

Gross Weight Fuselage Sta. MomentItem (lb. ) (in. ) (in. )

As weighed 1,151 117.9 135,699Less oil 15 147.6 2,214Less Instl. wiring 15 80. 0 1,200

Total 1,121 118.0 132,285

50

-- I I i_ _ i_ __ _

TO NOSE~

WI, 0 PLAT1FORIM

s'rA

-5.0

STA

2"

WI 144 .11

WING__

FWIN

1031 17.2

WI. 5:.4 . .- --- 'Im'I

WI. :8.3,

Figure 11. Reference Data for Weight and Center ofGravity Calculations

51

106.0 103.0*MAX CGTRAVEL I

107.0 99.0

2300

2100 - -

1700 '.

1.,00 ,,

3-

1300 -

AND CG

*FUSIELAGE STATION 10:1.0 RIAESFNTS 'IllE CURR|ENTLOCATION FOR NOMINA, OPERATING C CNEn OF GRAVITY

Figure 12. Loading and Burl -1f Curve

52

STRUCTURAL TESTS

WING AND PLATFORM STRUCTURE

Test Vehicle

The wing and platform structure test vehicle was a completely riggedvehicle, less engine. The aileron cables were anchored to the bell-crank pivot bolts, to prevent aileron movement, because the aileronbellcranks were not available when the test was conducted. A testfabric was substituted for the production wing fabric.

Test Procedure

A proof load of 3,140 x)unds (2. 5 g) ptr semispan was applied to loaddistributing rods in each wing test fabric, 20. 65 degree., forward and13 degrees outboard, with reference to the plan of the leading edgeand keel and thf, keel et-.terilne, (Figure 13). The wing was at a30-degree angle of incidence to the platform, which was inclined 11degrees nose down, as shown in Figure 14. Loads were applied in 0,20, 40, 60, 80, and 100 percent increments of the proof load,withreturns to 20 percent after each increment over 60 percent to check

for pernmanent set.

The test vehicle was tied to the ground at fuselage stations 56 and 120to react the applied wing loads. The proof loads were applied to thekeel and both leading edges through the wing test fabric, which was at-tached to the leading edge and keel in the same manner as the actualwing fabric. The test fabric was of sufficient width to reproduce theproper designed angles at the leading edges and keel. Proof loads wereapplied by a single hydraulic cylinder and a whiffle tree system attachedto a load distributing rod in each wing section, (Figure 15). The totalapplied load was measured by a load cell.

A 330-pound side load was applied to the apex of the wing with a simul -taneous 2,300-pound (1 g) wing load. The vehicle was tied to the ground,and measurements were taken as in the preceding test.

53

LOA______ J~tOUNImIunAN

Figure13. Whff le ROnfgro

----------- --- ---- --- - 5-

Figure 14. Static Test Setup

Figure 15. Keel Deflection

55

Test Results

The wing and platform structure satisfactorily withstood the appliedstatic proof loads. Some bending of the aft section of the keel wasobserved at higher loads; however, no permanent set was discernible.See Table 7 for measurements taken and the values obtained. Recordswere taken on a recorder.

CONTROL SYSTEM

Test Vehicle

A completely rigged test vehicle was used to test the control sysLem.

Test Procedure

The neutral and extreme positions of the pilot's controls and corre-sponding control surfaces were measured and recorded. Deflectionswere measured with a steel tape.

The steering pedals were loaded simultaneously with 250 pounds each,and deflections were recorded. A 200-pound load was applied to thebrake pedal with the right-hand steering pedal fully depressed (Figures16 and 17).

A load of 53 pounds was applied tangentially to the pitch control wheelin both directions with the keel restrained at the aft pitch cable attach-ment point and against the forward down stop. Deflections of the wheeland keel were recorded (Figure 18 and Table 8).

The ruddervator push-puli rods were anchored at their aft ends with thecontrol column in the neutral position. A 200-pound load (100 poundson each side of the wheel) was applied to the column in both directionsand the deflections were recorded (Figure 17).

The wing was restrained at zero roll by attaching the roll stop cablesto the platform tie-down rings. With the upper aileron bellcranksbottomed, a 1, 126-inch-pound couple was applied to the wheel in bothdirections, and wheel deflections were recorded (Figure 17).

Test Results

The control system withstood the applied test loads. Deflections in allsystems occurred through stretch in cables and/or deflection of thepulley brackets. No bending of any control equipment was observed.

56

TABLE 7WING AND PLATFORM TEST DATA

Vertical Load Only Vertical Load 1 g,Function and Location 2. 5 g Side Load 330 lb.

Forward Pitch Cable

Tension 0 0

Aft Pitch Cable Tension 1,700 lb. 910 lb.

Bending, Keel at Pivot 11, 000 P. s. i. 3, 500 p. s. i.

Axial, Left Hand Fwd"V" Brace 0 2,400 p. s. i.

Axial, Right Hand Fwd 200 p. s. i. 2f 200 p. s. i."V" Brace 200 p. s. i. 2, 200 p. s. i.

Axial, Left HandSpreader Bar 7, 900 P. s. i. 2, 200 p. s. i.

Axial, Keel at Apex 2, 000 p. s. i. 0

Vertical Bending,Left Hand LeadingEdge at Pivot Ball 34, 000 P. s. i. 12, 200 p. s. i.

Lateral Bending,Left Hand LeadingEdge at Pivot Ball 1, 000 p. s. i. 1,000 p. s. i.

Axial, ILeft Hand Aft"V" Brace 1, 000 p. s. i. 0

Axial, Right Hand Aft"V" Brace 2P 000 p. s. i. 1, 800 p. s. i.

Axial, ForwardCenter Strut 5, 000 p. s. i.0

Lateral Bending,Left Hand LeadingEdge at Pivot Point 7,400 p. s. 1. 2, 600 p. s. i.

57

PITCH lIP ) 4 ,(ELEVATOR)

ROLL

11(Il DO)WN

PITCH THIN UP - LE.IVATUOh

TINIT I)W

(htI OI GLAH

STE*. k;H III.4l~ *it# GUAR)

Figure 16. Cockpit Controls Proof Loading Test Configurations

58

CONTRO' COLUMN DEFLECTION

DEFL DEFL

FrWD AFT-J

P - 200 LBF UDDERVATOR PUSH-PULL TUBES FIXED

200 LB FWD, 12 3/4 IN. FWD DEFL200 LB AFT, 9 3/4 IN. AFT DEFL

WHEEL DEFLECTION

M = 1126 IN.-LBAILERON BELLCRANKS BOTTOMED

ROLL STOP CABLES FIXED AT ZERO ROLLFOR CW MOMENT, WHEEL TURNED 27o

CCW CW FOR CCW MOMENT, WHEEL TURNED 25'

RUDDER AND BRAKE PEDAL DEFLECTION

-P BRAKE - 200 LB

-- P RUDDER = 250 LB/PEDAL

BOTH RUDDER PEDALS LOADED SIMULTANEOUSLY1.15 IN. DEFL LH PEDAL

DEFL 1. 30 IN. DEFL RH PEDALBRAKE PEDAL LOADED, RUDDER PEDAL FULL DOWN

NO NOTICEABLE BRAKE PEDAL DEFLECTION

Figure 17. Control Column, Control Wheel,and Brake Pedal Deflection

PITCH TRIM DEFLECTION

PIVOT

C' PITCH ANGLEPITCH DOWN STOP

P - PITCH DOWNPA _ ' P B- PITCH UP

I dP A & PB - 53 LBPB

PITCH UPCABLE STOP

Figure 18. Keel Deflection Under Load

59

TABLE 8KEEL DEFLECTION LOAD RESULTS

Pitch Keel DeflectionAngle Load Point Point A-Point B Wheel(Deg.) Applied Restrained* (in.) (in.) Rotation

15 Pitch dwn. B 2. 5 dwn. 1. 3 up I+ turns15 Pitch up A 0.7 up 1.6 dwn. 1- turns

0 Pitch dwn. Pitch dwn. stop 0. 28 dwn. 0. 28 up + turn30 Pitch up Pitch up stop 0.25 up 0.25 dwn. +turn

*See Figure 18

TAIL ASSEMBLY

Test Structure

The tail assembly test structure was a complete left-hand ruddervatorattached to the vehicle. The aft elevator bellcrank was locked in theneutral position during the test.

Tes. Procedure

The vehicle was inclined to orient the iuddervator in a horizontalattitude. Loads and deflections were measured at the positions shownin Figure. 19. Five-pound shot bags were used to apply the test load.

4 The shot bags were positioned as shown in Figure 20 for each incrementwhile hydraulic jacks were supporting the ruddervator. After each load-lig, the jacs were lowered away from the surface, thus applying theload uniformly. The test structure was loaded in 20, 40, 60, 20, 80, 20,90, 20, and 100 percent increments of the designed limit load.

Test Results

The ruddervator sustained 100 percent limit load for approximately30 seconds, followed by a buckling failure of the lower surface of thefin. See Figures 21 and 22. The deflection and strain data arepresented in Figure 23, details A through F.

60

o - DEFL POINT DIAL INDICATOR

A -STRAIN GAIIiES(TOP & BOTT1OM) % LOAD (INCHES-p.)

NO. 16 STRAIN GAUGE ON AFT

INB'D HINGE FITTING 20 .14240 .151

4 60 .1707 20 .146

S80 .1920 .154

15 14 13 12 16 90 .214

3 2 120 .1.58

8A

9 1010

DIAL INDICATOR

Figure 19. Deflection Point and Strain Gauge Locations, Ruddervator

60

90

9O 'to 80 60 80 90

H0 lot) 21) 80 40 20 20 40 610 100 90

,20 40 C0 20 80 r,0 20 10600

80 20 100 40) 40 -0 CI0 90 '10 20 JO 100

00 -t0 20 80 100) 20 80#80 20 40

610 10 20 80 00 90 so 20 40

10 100 40 o 0o0 20 100 80

80 41 o 010 20 401 80 H00 4 (TY 1)

90 6 0 90 60 2 40 80

L - - - 4 IN. (TY P)

IINMBERS INDICATE 5-1.13 SHOT Y3AGI,()('ATI'()N (CUIMULIATIIVE) 1")I.'O l.(] ACH ,

INCHl-NI.INj tI I)I' SI(IN IMIT I, N(

Figure 20. Shot Bag Configuration, Ruddervator

6; 1

*~Ii~z&

Figure 21. Ruddervator Failure

Figure (LZ. Ruddervator Failure, Shot Bags Removed

6i

!ARGER SYMBOLS INDICATE RETURNS TO 20% (TYP) i.e ____

8 DE-- NO.-1 8DEFLNO. 4

DEFL NO. 2 t DEFL NO.6 ,

1.2 -1.2-

1.0

_I_ I4.8 - --- 4

o Ao I

N 6 --- __ -I-i a.6 . . .. - .. ,

1.0

.2 - - - ----. 2-7 . ..---

P4.

6 IE/ NO 56oDFLNO 8S .2 .2 ,,/ '

20 40 0 20 40 60 80 100

A PEICENT LIMIT LOAl) B PERCENT LIMIT LOAD

8 DEFL NO. 7 8 DEFL NO. 10 o

-3

. .. .... ... .. :3..... ,,- -....... . -.. --,_----..-..- , - - . .,,,.,, .. . ;

2.4 013.0 - - -

2.0 2.6 -

~1,6 2.0

1.2 jO

.4 .5-

0 0 4 0 80 100 0 20 40 60 80 100

c PERICENT LIMIT LOAD) D PERCENT LIMIT LOAD

Figure 23. Ruddervator Deflection Data (Sheet 1 of 2)

63

8 DEFL NO. 12 8 DEFL NO. 14

DEFL NO. 13 DEFL NO. 15

8 DEFL NO. 16

12- 3---

10 ___- I

I I

1-' I i-2 _

0 20 40 60 80 100 0 20 40 60 80 100

E PERCENT LIMIT LOAD F /tCENT LIMIT LOAD

Figure 23. Ruddervator Defle ion Data (Sheet 2 of 2)

DROP TESTS /

Test Vehicle /"

The entire vehicle, less wing and support sti lcft' e, M/as dropped Underconditions set forth in the. following paragi",ph. The centcr of gVavity

of the test vehicle was located at fuselage station 103. A 435-pounddummy engine was used,and the test vehicle drop weight was/'1, 635pounds.

Test Procedure

The test vehicle was dropped from various heights and attitudes ontosloped and level surfaces to simulate the following landing conditions:

3-point landing2p '1oint landing

Tail-low landingNose-low landingSide-drift landing

64

W r

The following drop configurations were used. Vehicle attitude refersto distance from wheels to ramps. A minus ramp angle indicatesspring-back load.

Configuration 1 - Vehicle level, Main Landing Gear (M. L. G. ) ramp-26. 7 degrees, Nose Gear (N. G. ) ramp -37. 6degrees. See Figure 24, detail A.

Configuration 2 - Vehicle and ramps level. See Figure 24, detail B.

Configuration 3 - Vehicle level, M. L. G. ramp -15 degrees, N. G. ramplevel. See Figure 24, detail C.

Configuration 4 - Vehicle 7 degrees nose low, ramps level. See Figure24, detail D.

Configuration 5 - Vehicle level, left-hand M. L. G. ramp 38. 9 degrees,right-hand M. L. G. ramp 31 degrees, N. G. ramp level.See Figure 24, detail E.

The ramps were greased to simulate main gear wheel "run-out" whichoccurs during actual landings.

Test Results

The vehicle satisfactorily withstood all the drop conditions with nodiscernible damage. All accelerations were recorded, and an oscillo-graph recorded all strain gauge outputs. Data from recordings arecontained in Tables 9 through 13.

LANDING GEAR STATIC TESTS

Test Article

The test article consisted of a spring-axle assembly (Figure 25, detailA.)

Test Procedure

The spring-axle test assembly was mounted in a rigid jig In the samemanner in which it attaches to the airplane. See Figure 25, detail B.A hydraulic cylinder was used to apply the load in increments of designlimit, , 50, pounds, u,,-1l faiiure occurred. The axle deflections werevisually nea. ured with a graduated scale and the data recorder. Land-ing-gear loads were measured with strain gaul,:es, which were locatedas shown in Figure 26.

65

/

Detail A Detail B

Detail G

Figure 24. Drop Test Configuration (Sheet 1 of 2)

66

Y -1 .......L .._.__._._.

Detail D

Detail E

F'iIure 24. Drop Tlest Configuration (Sheet 2Z of 2)

67

-y

,)-4

r-

68

t t ti

cu)

684

~4S

Cd0

C).

;-4 C-4 c'Go m-4

vi cA; uSZeqd

0 . bO 0O b aO b0-4 U)

-4

o0 ~CD CD

00

-E d- d 0c0 4

0)0

,.44 CU 4 -4 C

Q) 0)

ci m

0C Cd 05-4 44

4-J *r- 040 .-

0U 0bi2 0 j) .2 42

r. -4 90 - ;04-

69

bfl b O l 0 0 0)0- ~4 CQ t-

4-J C* w q4 U0

4 NP M~ *M 0 : 3

cl Cr uC cZi Ol; Lr l

Z~ Lo o

z 'tC bfl *Du 0 0 CD

Z oq 0 C~j LOl0 04 L0 00 Cl C lt V

bn ho bb bbN CD f C) Cd

0 l

C) 02 cob

-4 -4

02 I2 02 0DCD)

co Uc) in C:) 06; 6 CD CU ct

~t i C) -H- -H- C' 44 '-4 C%j tW) 0

0) id a )

41- - 4 .

Cd~a a) )-4 Cd io IL-4 -0 bO C. D0

a)Q .a ~0 0 0Cd CD :j -' C;~- 0 060C)dt C) 4 d C.)d ,. PQ P E ' l

0 Cd > ~ -0 Z) 00

'~~ 0 0-d b. J ) a) V) b.0 . %

Cd Cd C

:: 0) bI cd mJ Q)Z z f 0~~*~P4 P4 < .

70

~~C)

00 4 Q iC

0 o 0 0 0CO M~b Ln bn 0 VL) 0

u~ at~00~ * . obb

~ CD ~-V)

6dc

Hl 0 0

E,~~fl bl; cl d Cj ll 0 0

0

w) 0 )C

z 0 0 -0cr- C*1 c

Cd

41 (3 ) Db

0 ;-

'.4 .) a ~0d 0.

C) Ua, ) ba)4

Cd CU 0U1

71

V) a

rLca ho blo bDbfU bfl bfl C) C)

V; vi c6 ci 6

z *CI (A2 C~ 1

0 4 b.ObLuO.Lf a bb b o bD C ) C

0 0.

E- *D U b1) bi)

PH 0 0m

0 0l 04C2 0

0

-- 44cl 0Cd aoO d 0) b

u Ud U r.U) dd C. -4bI

~~' a)0 -Wd ~ ~ O

- U ~ C' -0UU C- WJ

0 4 0d 0 -C

0 bl 0 c(1) wZi oZ b

72

ob.0

t2 4 >: '~ -

z Ie m .

M c 0 0 bil bD 0 0 00. co

cv4 w~I~

Cd

00

0 04 C

-4 0- 14 Cz 0

*- CF2

c ~ ~ - -4 Aj -d41 Cd (1 bCdC

-4-

CU CU 4

0

> -4

73

Test Results

The axle failed just after 150 percent of designed limit load wasreached. Figure 25, details C and D show the nature oi the failuread also indicate secondary damage incurred when the spring impactedtl. broken end of the axle onto the concrete under the jig. The straingauge and deflection data appear in Figures 27 and 28.

12 IN.

. 8-9 TYPE IIE-141-350

8-23 TYPE C12.- 142B-350

i Figure 26. Landing Gear Test, Strain Gauge Locations

INCREASING ADDEC EASING LOAD

8-23 -9 8-7

100-/ .- -- - :~~0 _ /_ _,-"

S I U O I

3 r -

STUTRLTS. LN I GER1OE 6

50 -H -141-350 I IN.

- T 2x o3

4x 0 6x1o3

8x0o 1ox1oa

12X1O, 14xbo

U STRAIN (MIC lOINCI|EI4/INCII)j

Figure 27. Landing Gear Test, Load Vs. Strain

74

A O

______ NG OA

-B-23- -9s . -7

23

22 - - - - - - - - - - -

21 - - - - - - - -

20 - - - - -

15-- -

13 ---

a

3---

0 10 20 30 40 A0 00 70 80 90 100 110 120 130 140 150

LOAD (PERCENT OF LIMIT)

Figure 28. Landing Gear Test, Load Vs. Deflection

7;5

FLIGHT-TEST EVALUATION

INSTRUMENTATION

The two aircraft were instrumented as set forth in Ryan Report No.62B021. Measurements were taken for proof of design from a stand-point of loads, stresses, performance, stability and control infoizna-tion which will be used in the overall evaluation of the aircraft. Thefirst aircraft completed was instrumented for structural static tests.When these tests were completed, it was instrumented for flight tests.

Flight Test

Both aircraft were instrumented for flight test to determine controlsystem displacement and forces, accelerations, airspeed, altitude, andg forces in the X, Y, and Z axes. Displacements were measured withpotentiometers, and cable tensions were measured with load links.Accelerometers were used to measure acceleration, g forces, and vibra-tions. Low-pressure transducers were used for altitude and airspeedmeasurements. Engine operation, r.p.m., and temperature were alsomeasured and recorded. Records were taken on a 26-channel oscillo-scope during the taxi and initial flight tests to determine accelerationforces, control surface position, and control forces. As the test pro-gressed, additional measurements were taken to determine stress levelsin the structural members of the aircraft. Two standard 12-volt storagebatteries connected in series were used as a power source for in-strumentation. The recording instrumentation and associated equip-ment, including signals conditioning boxes, were installed on removableplywood pallets secured to the aircraft cargo bed with the cargo tie-down rings.

GROUND FUNCTIONAL TESTS

General

The two instrumented aircraft were given complete ground functionalte.. s to insure satisfactory operation of the power plant and controlsystem. The tests performed were:

1. Power plant operation and control.2. Fuel, oil, and cooling system checks.

76

3. Flight control and trim system characteristics.

4. Controls proof test.

5. Braking system check.

6. Aircraft weight and balance.

Power Plant Operation and Control

The power plant and its controls were checked to determine engineperformance, control settings and adjustments. Engine performanceand operation of controls were satisfactory; as necessary, the mixturecontrol was set, the idle cutoff was adjusted, and the magneto operationwas performed. The propeller pitch setting was determined and set toobtain 2,700 r.p.m. with full throttle. Induced engine vibration was notexcessive.

Fuel, Oil,and Cooling System Checks

A complete check of the fuel, oil,and cooling system was performedto determine the fuel tank capacity, wobble pump pressure, fuel pres-sure at engine pump outlet and at injectors, oil pressure, and accuracyof the fuel quantity and fuel pressure gauges. The fuel tank capacityis 25 gallons, with the fuel quantity gauge being calibrated accordingly.Instrumentation measurements were taken for cylinder head tempera-ture, barrel temperature, oil pressure and oil temperature at a sus-tained r. p. m. of idle, 1,000, 1,200, 1, 500, 1, 800, 2,000 and 2,200.All measurements were satisfactory, with none of the temperature andpressure ratings being exceeded.

Flight Control and Trim System Characteristics

The following items were inspected and/or checked for proper installa-tion prior to subjecting the vehicle to tests of any kind:

1. All fasteners for security; i. e., pulley brackets, cable unions,control mounting, etc.

2. All movable joints and hinges for excessive play. loosebushings, undue friction, etc.

77

- , .- -- ~

3. Alignment of attached components; i. e., cable pulleys, con-trol surfaces, foot pedal/nose wheel, control wheel/aileronwing, etc.

All discrepancies noted were corrected prior to performing tests.

Controls Proof Test

Prior to application of proof loads, the extreme ranges of the controlsystem were measured with a steel tape. These measurements weretaken of both the pilot's controls and associated control surfaces, withelevator control measurements being taken at the center and bothextremes of pitch trim. The following loads were applied to the pilot'scontrols with the control surfaces mechanically locked in the neutralposition and hydraulic cylinders used to apply the loads:

1. A 53-pound load was applied tangentially to the top of the pitchtrim control wheel in both directions.

2. Two 100--pound loads were simultaneously applied at the rim,normal to the plane of the main control wheel, and diametri-cally opposite horizontally. Loads were applied in both direc-tions. Load point deflections were measured at 40, 60, 80and 100 percent of the applied load.

3. Two 80-pound tangential loads were applied simultaneouslyto diametrically opposite points on the rim of the main controlwheel as a couple.

4. A 300-pound load was applied to the right-hand rudder pedalwith the nose wheel locked in the neutral position.

5. A 300-pound load was applied to the brake pedal with the left-hand rudder pedal locked. Maximum deflections of the brakepedal and right-hand rudder pedal were measured.

The control system withstoed the applied test loads satisfactorily.

Braking System Check

The complete braking system was checked for brake pedal operation,for excessive pedal movement, and for leaks. Operation of the brakingsystem was satisfactory.

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Aircraft Weight and Balance

Aircraft weight and balance was determined by using three platformscales. The as-weighed condition included the empty aircraft withfull oil and instrumentation wiring only. The aircraft was weighed inthe following conditions prior to the first taxi operations:

Platform Angle N. W. Oleo Strut Wing Incidence

6. 0 deg. Fully extended 0. 0 deg.0. 0 deg. Snubbed 0. 0 deg.0. 0 deg. Snubbed 26. 0 deg.

20. 25 deg. Fully extended 0. 0 deg.

The results of this initial series of weighings were as follows andconstituted the basis for weight and balance calculations throughoutthe remainder of the program:

Gross Weight ............... .1,151 poundsHorizontal Arm ............. .117. 9 inchesVertical Arm ................ .54. 5 inches

The recommended center of gravity limits for the XV-8A in the fullyloaded condition are from fuselage station 97 to 109 inches.

It was found during the aircraft weighing operation that variation inwing incidence angle had no effect on horizontal center of gravitylocation.

TAXI TESTS

Low Speed

The low-speed ground handling characteristics are very good. Theresponse to nose wheel steering inputs was instantaneous and positive,and required only light pedal forces. Minimum radius turns of 20 feetcan be accomplished. No significant sw'ty, lean or heeling was evidentduring relatively fast turns. The vehic i is highly maneuverable evenwhen operating on unprepared surface, Braking action is satisfactory,although fairly high pedal force is required. There is little or no pedalfree play. Abrupt brake application or hard steady application did notinduce any landing gear chatter, oscillation or excessive deflections.

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_______ I.

No significant brake heating was apparent during any of the test opera-tions. The brakes are adequate to hold the aircraft static againstpower application up to 2,710 r. p. m. 90 to 95 percent of takeoff power.At higher r. p, m. the wheels remain locked and the aircraft will skidforward. No adverse or unnatu,-al control procedures or techniqueswere encountered as a result of having a single toe brake pedal con-figuration on the right nose wheel steering pedal. The behavior of thefiber glass main landing gear struts was excellent. Approximately 900r. p. m. ip required to initiate a taxi roll from a standing start.

High Speed

Lateral canting of the wing downwind during a crosswind taxi requireshigh lateral control force inputs on the part of the pilot to hold the winglevel. It was noted thatwhen taxiing at speeds in excess of 15 milesper hour, these forces tend to lessen, which can be attributed to theservo effect of the leading edge tip ailerons. Elevator control effec-tiveness was not in evidence at taxi speeds up to approximately 33 milesper hour at a wing incidence angle of 20 degrees, and only a slight de-gree of effectiveness was observed at 28 degrees of wing incidence.The elevator control forces were very light. The aircraft became lighton the gear at approximately 33 miles per hour calibrated airspeed, andsome lateral rocking was present. Nose wheel shimmy was absenteven with the aircraft light on the main gear. Directional response tolateral control inputs was evaluated at wing incidence angles of 20 and24 degrees with the nosewheel steering pedals unrestrained. Negligibleresponse to the control inputs was noted at speeds up to 30 miles perhour calibrated airspeed. Divergent oscillation of the nose wheel wasobserved as a resulting phenomenon to these inputs if the pedals wereleft free. Even at low taxi speedsa light tap on the pedal would inducea divergent nose wheel oscillation with the pedals free. This conditionwas eliminated by the addition of a shimmy damper to the nose gear.No lateral-directional coupling was evident when the landing gear was

' ,ifirmly on the runway. Some brake fading was apparent when brakingto a stop from approximately 33 miles per hour calibrated airspeed.

Crosswind Taxi Operations

Taxi operations were evaluated in moderately gusty crosswinds rangingfrom 8 to 15 knots. S-turns were executed at a speed of approximately10 miles calibrated airspeed. The control force required to holdthe wing from canting downwind was excessive. As these tests pro-gressed, the wing incidence angle was increased from zero degree in

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_ _ _ __ __ _

2-degree increments. The lateral control forces increased steadily tothe point where at 16 degrees wing incidence, it became very difficultto move the wing in the upward direction. Minimum lateral controlforce is required at 0-degree wing incidence angle; however, a con-siderable amount of wing lobe flutter is present at this condition whichmay not be conducive to a long service life of the fabric. It was alsonoted that while taxiing at a wing angle of approximately 20 degrees,there was no tendency for the aircraft to heel excessively even with thewing canted over to the stop.

HANDLING QUALITIES

Nose Wheel Lift-Offs (Basic Configuration)

High-speed taxi runs were made at wing incidence angles of 20 to 24degrees in an attempt to achieve nose wheel lift-off at a mid center ofgravity location at speeds up to 43 miles per hour calibrated airspeed.In all cases the main gear broke ground before or, at best, simultaneouslywith the nose gear. Lateral control forces were light, and no responseto any longitudinal control inputs was observed. Additional tests wereconducted at wing incidence angles of 20 to 24 degrees to attain nosewheel lift-off at the aft center of gravity location. The speeds attainedranged from 20 to 43 miles per hour calibrated airspeed, in 5-mile-per-hour increments. Nose wheel lift-off could not be accomplishedon any of the aforementioned runs.

Aircraft Lift-Offs (Basic Configuration)

The first lift-off was made at a wing incidence angle of 24 degrees atthe aft center of gravity with 2200 r. p. m. at a calibrated airspeed ofapproximately 46 miles per hour. The aircraft became airborne throughpower effects rather than elevator effectiveness, and no attitude changewas apparent following lift-off. The lateral control was satisfactory;forces were light and within the capability of one-hand control. Lon-gitudinal control was ineffective. The landing was accomplished bypower retardation rather than aft stick movement. Additional lift-offswere conducted at wing incidence angles of 24. 5, 25. 5, 26, 26. 5 and27 degrees at the aft center of gravity location to determine handlingqualities. All lift-offs occurred with essentially a constant body attitude.Lateral control was acceptable, ut the control forces ,4-.... ^A. e t-o

high when compared to a conventional airplane. Longitudinal controlwas unacceptable even for an emergency condition, in that large control

81

inputs produced only very small pitch changes after a considerabletime lag. The only difference noted, when changing the incidence anglefrom 24. 5 to 26. 5 degrees, was a reduction in lift-off speed by 2 milesper hour, from 46 to 44 miles per hour calibrated airspeed. At the wingincidence angles of 26. 5 to 27 degrees, control was even more marginal,especially in pitch. Lift-off occurred at approximately 43 miles perhour calibrated airspeed in a manner similar to that described above.

A high rate of descent to touchdown produced a pitch-up about the maingear sufficient to scrape the elevator trailing edges on the runway whenin a full-down position. One additional run was made at the mid centerof gravity location which exhibited completely unacceptable character-istics. The aircraft attempted to fly off with the nose gear skipping onand off the runway. Longitudinal control inputs were completely ineffec-tive in alleviating this condition. All remaining operations were con-ducted at the aft center of gravity limit. Lift-offs were also conductedat incidence angles of 20 and 22 degrees at the aft center of gravitylocation. Lift-offs occurred between 50 and 60 miles per hour calibra-ted airspeed. Elevator effectiveness appeared better than at the higherincidence angles because of the higher lift-off speeds, but was stillunacceptable. Landings were extremely touchy by reason of the lackof longitudinal control and the higher landing speeds, with the landinga questionable item until the main landing gear was firmly on therunway.

High-Speed Taxi (Auxiliary Elevator)

Additional high-speed taxi runs were conducted after the addition of anauxiliary wooden elevator mounted at the aft end of the bed between theoriginal elevators. Total elevator area was thus increased by 1, 750square inches by this additional surface. These runs were made atairspeeds over the range from 33 to 52 miles per hour calibrated air-speed at incidence angles of 16 and 18 degrees, during which the pilotpulsed the elevator. No vibrations were induced and no adversecharacteristics were noted. One taxi run was accomplished at 46 milesper hour calibrated airspeed to observe the effect of trailing edgebattens on the wing tip flutter. No flutter was observed during thistaxi run.

Nose Wheel Lift-Offs (Auxiliary Elevator)

Nose wheel lift-off was attempted at wing incidence angles of 20 and 22

82

--1 o-om1 - A.

degrees at speeds up to 53 miles per hour calibrated airspeed. Thenose wheel could not ] 0 lifted off at this aft center of gravity position.In order to augment the elevator control power, a flexible cloth sealwas added between the aft end of the bed and the auxiliary elevator, andthe engine ejector hole in the bed was skinned over. Test results were

the same as for the auxiliary elevator alone during the one operationconducted in this configuration.

Aircraft Lift-offs (Auxiliary Elevator)

Several lift-offs and flights were accomplished at 24. 5 degrees wingincidence angle at airspeeds between 50 and 55 miles per hour calibrated

airspeed. Although the nose wheel lift-off was not attained, a positiveattitude change capability was achieved. Elevator control power wasstill insufficient to achieve an initial main gear touchdown at landingspeed, 49 miles per hour calibrated airspeed. Large longitudinalcontrol displacements were required to produce attitude and subsequentaltitude changes; however, the associated time lag was not objectionable.Platform attitude was nose-up during these operations. The auxiliaryelevator was lightened from 31 to 21 pounds by the drilling out oflightening holes and skinning the surface with fabric. Taxi and takeoffoperations indicated no appreciable effect due to lightening of theelevator. No pitch-trim changes were evident at lift-off as enginer. p. m. was progressively increased from 2, 300 to maximum r. p. mi.for takeoff. A slight amount of left wheel application is required atlift-off to counteract torque at maximum r. p. in. Shorter ground rollsand steeper climb angles resulted from increase in r. p. m. for lift-off.Slight control wheel buffet was observed which was attributed to trailingedge flutter near the wing tips. Lateral battens, 61 inches in length,were added to the left-hand and right-hand outboard wing panels parallelto the trailing edges to eliminate wing fabric flutter. No flutter or con-trol wheel vibration was not'd during the resulting lift-off and runwayflight. Additional aircraft lift-offs were made at incidence angles of26. 5 and 27. 5 degrees at lift-off speed of 46 and 42 miles per hourcalibrated airspeed respectively. Longitudinal control was unacceptable,iL 26. 5 degrees and nonexistent at 27.5 degrees ;ni'i(flne angle. Thed I rcraft was unstabie during te landin r at 27. 5 deg;recs incidence angle.

PERFORMANCi'

Some preliminary level-flight performance data at wing incidence angleof 24, 5 degrees are presented in Figure 29. These data are preliminaryin nature and are uncorrected in view of the limited number of data

83

points obtained. Indications are that these performance parameterswill be equal to or will exceed predicted values. These level-flight datapoints were obtained with an H-23 helicopter as a chase-pace vehicle.

STABIL'ITY AND CON ROL

The first traffic pattern flight was made with maximum r. p. m. (2, 850)for takeoff. Lift-off occurred at approximately 50 miles per hourcalibrated airspeed, and the ground run was estimated to be 350 feetinto a 6-knot, 40-degree crosswind. A climb was made to 180 feet inaltitude, and power was retarded to 2, 400 r. p. m., which resulted in astabilized airspeed of 54 miles per hour calibrated airspeed at 100feet. One traffic pattern flight was accomplished during which shallowturns were made. The lateral and longitudinal control forces werelight; although the lateral forces were too high for a good harmony ofcontrol, they were sufficiently light to permit one-hand control. A slightright-hand wing heaviness became apparent during the flight. Post-flight inspection revealed that the right-hand wing bolt rope had failed.Fluttering of the wing fabric was present over the outer three feet ofeach wing panel. No difficulty was encountered with the crosswind, anda good landing was made. A fixed tab having a 15-inch span and 2. 5-inch chord was added to the auxiliary elevator at a 15-degree depressionangle. Six traffic pattern flights were accomplished. Some adverseyaw was encountered during S-turns at 56 miles per hour calibratedairspeed. In general, the aircraft handles satisfactorily. The maximumlongitudinal forces encountered at an incidence angle of 24.5 degreeswere 15 to 20 pounds push and pull over the speed range from 43 to 68miles per hour calibrated airspeed. Left-wing heaviness appeared toincrease with speed,as did the lateral control forces required to correctthe situation. Considerable trailing edge wing flutter was encounteredin the outer three feet of wing span which was fed back through the tipailerons to the control wheel. Descent during landing was completelyarrested by use of up elevator, requiring no power change. The rateof deflection was excessive, a momentary pause in level flight beingrequired prior to reestablishment of the touchdown rate of descent.One pattern flight was made with the trailing edge battens installed atan airspeed of 55 miles per hour calibrated airspeed. The batten in-stallation appeared to eliminate the wing trailing edge flutter.

Aircraft Lift-Offs (Wing Pitch Control System)

The auxiliary elevator was removed and the programmed wing pitch

84

LEVEL FLIGHTPRESSURE ALTITUDE A 550 FT

SYMBOL iwFTO OAT GROSS WEIGHT

M 2 4 1/2- 13 95' F 1900 LB

2700 - ___ ___

2600 .__

240

2200

;i 2000 -

1800

1600 I I I I

* ~~~1400 ...... I30 40 50 so TO so

CAB (MPH)

Figure 429. Performance Data, Level Flight

85

-. --L I ,zz"*

control system was incorporated for evaluation. This system pro-gramme& a ±1. 5-degree wing incidence change from trim upon full(19 degrees) forward or aft stick deflection. Tests showed that theresultant control forces were high. The ability to accomplish altitudechanges was essentially the same as with the auxiliary elevator con-figuration except that the auxiliary elevator provided a more positiveattitude resnonse. One other operation was conducted using the wingpitch control system together with the sealed auxiliary elevator. Thesame schedule of taxi runs was conducted as was the case for the wingpitch control system alone (an incidence angle variation of from 16 to24.5 degrees over the airspeed range of from 46 to 52 miles per hourcalibrated airspeed). Longitudinal stick forces were still too high (40to 50 pounds). This configuration was more responsive than with thewing pitch control system alone; however, the control forces were 4objectionably high.

Trim Change

An attempt was made to evaluate wing izicidence change (Ai ) as ameans of changing body attitude. A decrease of 90 degrees of pitchcontrol wheel (1/40 Ai ) resulted in an immediate descent to touchdown.The significance of this result was somewhat masked by the closeproximity of the aircraft to the ground. During subsequent tests thetrim wheel was rotated forward and aft to a maximum of 180 degreesfor a short period of time. No apparent trim change was evidenced duein part to the relatively small wing incidence angle change and the shortperiod of time that the new trim position was held.

Lateral Control

The incorporation of re-sized lateral control bellcranks did produce adecrease in lateral control wheel force which made one-hand controlacceptable. There was no apparent decrease in lateral control effec-tivenessand the sensitivity appeared to remain unchanged. Additionaltests were conducted following the incorporation of the auxiliary eleva-tor. Aircraft response to lateral control wheel force was good. Thecontrol forces and wheel displacement were satisfactory during shallowbanks and turns, and the aileron stops were not contacted during theselimited maneuvers, indicating that wing tilt was following aileron deflec-tion sufficiently to preclude full aileron travel in order to accomplishthe maneuvers executed. A light was incorporated in the cockpit toglow when full aileron travel had been reached. The light was not seento glow.

86

................ . . : .. . _ _ _ _ _.., - '- . - . ... . . . .... ..

Bolt Rope Effect

Two lift-offs were made at a wing incidence angle of 24. 5 degrees atthe aft center of gravity location with the trailing edge bolt rope dis-connected. No difference in body attitude was noted,and the oaly dis-cernible difference in characteristics, which could be attributed to boltrope removal, was a slight control column buffet.

Simulated Malfunctions

Aileron cable failure was simulated by first disconnecting one aileroncable and then both aileron cables with the wing incidence angle set at24. 5 degrees. Three taxi runs 9d short lift-offs were accomplishedwith one aileron cable disconnected. This condition produced an in-crease in lateral control forces, and the wing had to be tilted oppositeto the side having the inoperative aileron. The handling qualities weresuch that the vehicle could probably be recovered and landed should acable failure occur. One taxi run was made with both ailerons discon-nected. Lateral control forces were high,and approximately ±30 degreesof lateral control wheel input produced no response. This condition ismarginal and would probably cause damage to the aircraft during alanding.

OPERATIONAL TECHNIQUES

General

Techniques for performing the various operational functions with theFlexible Wing Aerial Utility Vehicle follow (it is not intended that thesetechniques will enable untrained or inexperienced personnel to operatethe vehicle):

Starting Engine

After checking the vehicle for readiness, start the engine as follows:

1. Set the fuel mixture control to the FULL RICH position.

2. Advance the throttle control to the 1/8- to 1/4-inch position.

87

- ~- - - - ----- -, -- -.-.

3. Rotate the propeller several turns by hand.

WARNING

Always stand clear while turning the propeller.When the engine is hot, or when the ignitionswitch is in the BOTH, LEFT, or RIGHT posi-tion, the propeller may kick or the engine maystart, causing injury to personnel.

4. Pump the wobble pump until fuel drips from the fuel injectoroverflow lines on the engine.

5. Turn the ignition switch to the BOTH (magnetos) position.

6. Signal the ground crew to crank the engine with the propeller.

7. If the engine fails to start, after several attempts, turn theignition switch to OFF, place the throttle in the full open posi-tion, and turn the engine over several turns with the propellerto clear the cylinders.

8. Repeat steps 2, 4, 5, and 6.

9. After the engine is started, adjust the throttle to idling speed(900 to 1,000 r, p. m.).

Engine Ground Operation

After the engine has started, assure that oil operating pressure iswithin limits, and warm up the engine by operating the engine at 900 to1,000 r. p. m. for at least 1 minute; follow this warmup by one at 1,200r.p. m.

Engine Ground Test

After the engine oil temperature has reached 75 degrees Fahrenheit(23.8 degrees centigrade), increase engine speed to 1,700 r. p. m., andperform the following steps:

1. Move magneto switch and tachometer switch to "R" positionand note the engine r. p. m.

88

2. Move magneto switch and tachometer switch to BOTH, then tothe "L" position and note the engine r. p. m. Maximum r. p. m.drop in the "L" and "R" positions should not exceed 125 r. p. m.,nor should there be greater than 50 r. p. m. difference betweenthe r. p. m. at "L" or "R" positions.

3. Return magneto switch to BOTH position.

4. Slowly and smoothly increase engine r. p. m. to maximum.Check for r. p. m. approaching maximum, engine runningsmooth, and oil pressure between 30 and 60 p. s. i.

5. Return throttle control to idle position.

Taxi

Check steering and brake controls before taxiing the vehicle. Taxispeeds should be relatively slow, especially where crosswinds aregreater than 5 to 8 knots. In order to reduce "heeling" tendencies incrosswinds, apply lateral control into the wind. This may best beaccomplished by leading the anticipated crosswind with wheel cortrolsuch that the wing is tilted into the wind before the crosswind is actuallyencountered. Failure to follow this procedure may cause the wing to bepicked up by the wind with subsequent excessive heeling of the entirevehicle. Little more than idling r. p. m. should be required to startthe vehicle moving and maintain a satisfactory taxi speed.

Takeoff

The Flexible Wing Aerial Utility Vehicle is designed to take off fromsemiprepared or rough fields in short distances. To take off, proceed

,as follows:

1. Set the brakes.

2. Check the fuel mixture control at RICH and increase enginer. p. rn. to maxiinurn.

3. Release the foot brakes and commence takeoff, using nosewheel steering to maintain a straight ground run. Aircraftwill fly itself off, with little or no aircraft rotation involved.

4. Use the pitch trim handwheel to trim to a 45-knot climb airspeed.

89

5. Reduce power and fuel mixture to that consistent with thedesired climb speed. Trim speed as required.

6. In takeoffs in crosswind, use lateral control to lower wind-ward wing leading edge into the wind to prevent heeling anddrift; maintain heading by means of nose wheel steering.

Climb

Following takeoff, power may be held at the maximum of 2, 800 r. p. m.for maximum climb performance. With a wing trim setting of 24. 5degrees incidence angle and an indicated airspeed of 45 to 50 milesper hour indicated airspeed, the vehicle is in good balance as very littleaft stick force is required. The climb flight path is relatively steepunder light to moderate gross weight conditions (up to 1, 900 pounds).

Descent to Landing

Recommended speed during approach to landing is 45 to 47 miles perhour indicated airspeed, with the minimum of 42 miles per hour indica-ted airspeed for 1, 900 pounds gross weight. To provide a flight pathwhich is not too steep for a comfortable flare to landing touchdown,approximately 2, 100 r. p. m. should be maintained during the glide.

Landing

Landing is best accomplished by accurately ,adijzl Auilg the flare maneuverto arrive at the bottom of the flare one or tv"' V ,t above the runwaysurface and reducing power slowly to idle. if the flare maneuver isexecuted too high, with the power setting for glide (2, 100 r. p. m. ), theairspeed will fall below that required for minimum longitudinal control(approximately 37 miles per hour indicated airspeed), and the sink ratewill be too great for a satisfactory landing. In such a case, power shouldbe EASED ON to control the descent rate to touchdown.

After Landing

Following touchdon,the vehicle may be declerate-d in a very shortdistance by applying brakes until wheel skid is felt. There is no tend-ency for the vehicle to diverge in direction if the nose wheel is centeredprior to its contact with the ground. Once all wheels are firmly on thesurface, excellent control for roll-out and taxi exists.

90

_ _ _ _ _ _ _ _ _ _ __. = II - U

Engine Shutdown

Shut down the engine after taxiing to the parking area as follows:

1. Place the throttle control in IDLE position.

2. Place the mixture control in IDLE cutoff position.

3. Turn the ignition switch to OFF.

91

MAINTENANCE REQUIREMENTS

GROUND SUPPORT EQUIPMENT

The equipment originally provided for ground support during the testprogram consisted of a tow bar, two wheel chocks, one 8-foot stepladder, and two 10-foot aluminum poles for support of the leading edgesduring the folding operation. During the test program it was found thatconsiderable difficulty was encountered when attempting to insert thequick-release pins in the spreader bar joints after extending the wingfrom the folded position. This operation has been greatly facilitated byproviding the quick-release pins with long tapered extensions whichserve as drift pins to bring the parts into proper alignment as the pinsare pushed into place. When towing the vehicle with the wings folded,it is necessary to pad the sheet .nietal leading edges where they contactthe tubular steel roll control structure. Removable pads have beenprovided which can be attached or removed from the roll control struc-ture by means of quick-release pins. These pads are now consideredto be a necessary item of ground support equipment.

SPECIAL TOOLS REQUIRED

It was found that a special wrench was needed to tighten the ignitioncable attachments at the lower spark plugs. This was made by slottingone side of a standard 7/8-inch-deep socket wrench. All other assemblyand maintenance operations on the vehicle were accomplished withstandard tools.

17

92

-_YAW". _.. ...

BIBLIOGRAPHY

REFERENCES

1. Campbell, J. P., "Summary of Methods for CalculatingDynamic Lateral Stability and Response and forEstimating Lateral Stability Derivatives," N. A. C. A.Report 1098, 1952.

2. Gray, V. I., "Representative Operating Charts of PropellersTested in the N. A. C. A. 20-Foot Propeller ResearchTunnel," N. A. C. A. W. R. L-286, 1943.

3. Hoerner, S. F., "Fluid Dynamic Drag," 1958.

4. Liddell, R., "Wind-Tunnel Investigation of Rounded Horns

and of Guards on a Horizontal Tail Surface," N. A. C. A.ARR No. L4J16, 1944.

5. Northrop Aircraft Inc., "Dynamics of the Airframe," BUAERReport AE-61-4, Volumes H and IV, 1952.

6. Pass, H., "Analysis of Wind-Tunnel Data on DirectionalStability and Control," N. A. C. A. TN 775, 1940.

7. Perkins, C. and Hage, R., "Airplane Performance, Stabilityand Control," John Wiley & Sons, Inc., 1960.

8. Purser, P. E., "Experimental Verification of a SimplifiedVee-Tail Theory and Analysis of Available Data onComplete Models with Vee-Tails," N. A. C. A. ACRNo. L5A043, 1945.

9. Ribner, H., "Formulas for Propellers in Yaw and Charts ofthe Side Force Derivative, " N. A. C. A. Report 819, 1945.

10. Ryan Report No. 63B021, "Flexible Wing Aerial Light UtilityVehicle," Volumes i through V, Ryan Aeronautical Company,San Diego, California, May 1963.

93

11. Sacks, A., "Aerodynamic Forces, Moments and StabilityDerivatives for Slender Bodies of General Cross Section,"N.A.C.A. TN 3283, 1954.

12. Sears, R., "Wind-Tunnel Data on the Aerodynamic Char-acteristics of Airplane Control Surfaces," N. A. C. A.ACR No. 3L08, 1943.

13. Sivells, J., "Tests in the 19-Foot Pressure Tunnel of aRectangular N. A. C. A. 0012 Horizontal Tail Plane withPlain and Horn-Balanced Elevators," N. A. C. A. ARRNo. 4, 1942.

HANDBOOKS

Pilot's Handbook for the Flexible Wing Aerial Utility VehicleXV-8A (Ryan Model 164), March 1964.

MANUFACTURER' S SPECIFICATION

Continental Motors Corp., Detail Specification for ContinentalAircraft Engine Model 10-360-A, April 25, 1962.

94

DISTR IBUTION

US Army Materiel Command 9US Army Mobility Command 5uS Army Aviation Materiel Command 5US Strike Command 2Chief of R&D, DA 3US Army Transportation Research Command 64USATRECOM Liaison Officer, US Army R&D Group(Europe) 1

US Army Natick Laboratories 1US Army Engineer R&D Laboratories 3US Army Limited War Laboratory 2Army Research Office-Durham 1US Army Test and Evaluation Command 2US Army Combat Developments Command, Fort Belvoir 1US Army Combat Developments Command

Aviation Agency 2US Army Combat Developments Command

Armor Agency 1US Army Combat Developments Command

Transportation AgencyUS Army Combat Developments CommandQuartermaster Agency 2

US Army Combat Developments CommandExperimentation Center 1

US Army War College IUS Army Command and General Staff College IUS Army Aviation School 2Deputy Chief of Staff for Logistics, DA IUS Army Infantry Center 4US Army Aviation Test Board 2US Army Airborne, Electronics and Special

Warfare Board 4US Army Aviation Test Activity zArmy Liaison Officer, Special Air Warfare Center 1Air Force Flight Test Center, Edwards AFB 1Air University Library, Maxwell AFB 1Commandant of the Marine Corps 2Marine Corps Liaison Officer,

US Army Transportation School INASA-LRC, Langley Station 6

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P_ On

Manned Spacecraft Center, NASA 2NASA Representative, Scientific and Technical

Information Facility 2Research Analysis Corporation INational Aviation Facilities Experimental Center 3Canadian Liaison Officer, US Army Transportation School 1Defense Documentation Center 20Ryan Aeronautical Company 5ARFO (LA) 2ARFO (ME) 2RDFU-T zRDFU-V 2ARPA (For Project AGILE) 21 Ith Air Assault Division 1Yuma Test Station 1Hq USATDS 2US Army Board for Aviation Accident Research 1

96

! .

UnclassifiedSecurity ClassificationSecurity________________DOCUMENT CONTROL DATA -R&D

(Security cleai.ifo tion of title, body of abatIrct end ltden. nootton m.f be enlered whe,, the -r ie report In cl.1, fljd)

I ORIGINATIN G ACTIl'IY (Corpor.te .uthor) 2a REPORT SECLIRII t A- ATa/ yan Aeronautical Company - UnclassifiedSan Diego, California Z ,R oR

N/A

3REPORT TITLE

XV-8A Flexible W ing Aerial Utility Vehicle

4 DESCRIPTIVE NOTES (Type of report e.d Iricitote d1t..)

Final Report August 1964 to January 1965

5 AUTHOR(S) (L. " irame. ftr t --see, Ini.l)

Landgraf, F. and Girard, P.F.

6. RErFORT DATE 71 TOTAL NO OF RGEFS 7 IO oi IFS

February 1965 96 s

Be. CONTRACT OR GRANT NO. In ORISINATOR'S REPORT NUMBER(S)

DA 44-177-AMC-874(T)b. PROJECT NO. USATRECUM 64-74

Task ID1214UIA14172SII I THER REPORT NI(S) (Any other n--her tI may be ... 0i.d

IIlar report)

d 65B003

10 AVAILABILITY 'LIMITATION NOTICES

All distribution of this report is controlled. Qualified DDC users shall requestthrough Secretary of Oefense, ATTN: ARIA-AGILL, Washington, D. C. 20310Release or announomantrhc kc p ublicrizd.__

Ii SUPPL EMENTARY NOTES I2 SPONSORING MILITARY ACTIVITY

Sponsored by: US Army Transportation Research CommandAdvanced Research Projects Agency Fort Eustis, Virginia

I3 ABSTRACI

This final report is based on interim report 64308 of contractor. Contractnumber was UA 44-177-AMC-874(T), AKIPA Order 294-62 Amendment No. 3, unclassified.This report discusses procedures and accomplishments of the design, fabrication

and the t,3st program of the XV-8A vehicle and

1. Investigates the construction of a Plexible Wing Light Utility Vehicle,simple to operate, and capable of transporting a 1,O00-pound payload for a dis-tance of 100 miles at a speed of 50 miles per hour.

2. Covers the design, building and testing of two vehicles.

3. Concludes that the vehicles as described are feasible, and recommendsfurther tests.

DD JAN64 1473 Unclassified

97 Security Classification

IJ

Unclassified

1N LINK A LINK B LINK CKLY WORDS H~ - OL L W, - .HOL -t_

INSIMC TIS

I.ORIINATINt. ACTIVITY: Fuler n1 ame- ild adrn o VAILAIIILI-IY LIMITATION NOTIVICEh EnT.r any lIm-ot hecohralo. :o.,,tri-il. )iiIt- Dtpatiiittis I lM I I s ot furher ilIsS,5(itaio 0 of the report. oth.- than to-or

the aIi~ rohri ai~~ii .irtctit utor ion llitseiib li ( itt iasiliiIon. USIInfr standard sltlenenla

R..1E1PORT SECURITY CL.ASSIFICATION: Fnter the to er- such .. : tli eustr (V(4110 ltlS(fti

.iII .:,irIly Ilii"r.ii hut ZO o the es,-rf. Indicate whether (1 reuii from ciit 017(bai2opeso'tt'tR-trs.I'd Dien a Ini tit." Mink li iii tit be ini' acr-f

Lil ithl iipril -ictiril y r.-gul.lIi'ni$ (2) torn i4'Iigo itinti and iisentitio ofi isi11(

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i/i-it.

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If a, imeaningfl title cannot be selected Without via isi (ig- rept It FdirviI thfruiitD Otequlfduer

ti-I, show title cl.i-sfiest ion ill iill Cafii iii parenlhest hmirqusihrui

inmediatelv illiiwitiv tie (dIe. _____

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re-port, e s. , interint, lirotse sti. so tnmary12. annjual, or fin[la. itidil UDC users siallI reqluest Ithrough

(live tile iricloaye ]ali-a whenl a specitic rh-oirtinig tn-ri .. IIiovreel. IIf the report hits been furnished to the Office of TechnicalS. AUThIOR(S): I-nter the ltnoe(5) ofl author(s) an shown till Services, Department of Commerce, for male to the public, indi-

4 or in the rep on. Koeti last nrame, fiest name, middltIiia tI at- this factIgo anene tile price, if k tombIf miilit ary, alshow rank sod branch it siervice. The name oifSthe princpal Luint (r is an atisolt oe nil Iinrilii reoiiretin,1 11, SUPPLEMENTARY NOTE&: Uae fI'm additional 'ipi anac

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oti the report, use dale of publicatio. the departmtental proct office or laboratory spov..;yring (pa-ita4 (or) t he research and development. Include address.

7s. TOTAL NUMBER OF PAGES: The total page count,shbould follow normal pagination proiceldreli. i.e., enter Ile. I I1 ABSIS.RAC.LEtit I iigsii-t giniii a Irii-f and) tfit iat

niOltier Of pages containing infornijon. .i ..i.... Liry of [In. .1--oiont iiiis ~ if the, -1rert.o tlougt.ilt 115(1dn itt-ut, .'l4.hi-tlir in tile bobl oif til- ti-clii.:it ri-

76. NIJMI3ER OF RICLIRENCES Entter Iii totlt otunkter if Ip .If 'Idil- ilt 'pa io. - '-iitrd. 11 cllouLiioll !Oilq tc-

refert'nce sc ci in the report. thult te Ltt ai-b'..

8.a. CO NTR ACT OR GRA NT NUlMBER: If 4qpiiit o, cot er It1 i hyigh ILiti l.at IThe .- lot itc nf. ti-son' h Itthe Alpi -blP nlumbler Of tile conotract or grant under which t.(t 'if til lassiriti. la-li Iiaravit itf The abstractl shltlthe riepoti Wts Wittei --tntf with anl inditicatiin -4C111(-wilitory secirity cliilinaon

811, lIc, & 8d. PROJ ECT NUMB ER: Enter I he appropni ati I liihe - ill tini iItle (LirliogtiIir'proat-nt'- cis ( TS). (S).

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subprnject nuntc systesi ouwbera. Task number, ec. Th-rie is OTi lnilltli-in oin the IL-niitli of the abIstracti. How-

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cial re-port number bly which the docuon-t will be icI-o- Ii-I I- KIEY %kstI. tr: Key -- bti oiri t ec-hnically oeaniol le inandi cionttoled biy the otiifltail iti ativity. This OOmlit (ut- orrt phirio-s 0-1 , ti.iiaiitI-rz &I (tlt tit ni my lie usedt as

be uniique to tis. repvolt. Iex n-iries~ foril loii ie( repo-rt. Key wtis nilust he

9bf. OTHE 10REPO-RT NUMB ERf S: If the report has b~een ou-l .t w,.-n v Iaoif a i- i, qrit)reit. iun-

aslgneil any oth-r reilort numhers (v:tIir ill the n-ritaLtir '(i.Li m r Oiitnii Model iio tiiitail,- lile. -ih-

or h1- Ii&. sponttoor), alIso cnter this numbier(%). ("(rv h)r..,.-- -I ~imm-, V(Lioli.riiri tociat iii, 01111 bi isedt 1goi-Y brE ill will be. b~-ei by ill inoliiLtioni Mtt'ii~

-inl~t Iti l%(4g~ll'lt : liiiks, rules. anti Vc11iiIti- 11;

Unc lassified98- . Security Ckssificii~o -

98

DEFENSE ADVANCED RESEARCH PROJECTS AGENCY3701 NORTH FAIRFAX DRIVE

ARLINGTON, VA 22203-1714

November 20, 2001

Ms. Kelly AkersDefense Technical Information Center8725 John J. Kingman RoadSuite 0944Ft. Belvoir, VA 22060-6218

Dear Ms. Akers:

This is to advise you that the following documents have been reviewed and/or declassified andreleased under the Freedom of Information Act.

- • Document Number: AD 803668Unclassified Title: Sailwing Wind Tunnel Test ProgramReport Date: September 30, 1966

-'. Document Number: AD 461202Unclassified Title: XV-8A Flexible Wing Aerial Utility VehicleReport Date: February 1, 1965

* Document Number: AD 460405Unclassified Title: XV-8A Flexible Wing Aerial Utility VehicleReport Date: February 1, 1965

*- Document Number: AD 431128Unclassified Title: Operational Demonstration and Evaluation of the Flexible Wing PrecisionDrop Glider in ThailandReport Date: March-July 1963

" Document Number: AD 594 137LUnclassified Title: Communist China and Clandestine Nuclear Weapons-Input SubstudiesA-J, SRI ReportReport Date: October 1.970

" Document Number: AD B 176711Unclassified Title: Overlay and Grating Line Shape Metrology Using Optical ScatterometryReport Date: August 31, 1993

If you have any questions, please contact Mr. Fred Koether, our Declassification Specialist, at(703) 696-0176.

_Z(62,/ 4K2Pc z ..-.. ',/ ,Sincerely,

-- Sancy a er6 DirectorLC

C_ &7".Security and Intelligence Directorate