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The Aerospace Corporation 2013
Launch Vehicle and Spacecraft Structural Dynamics
State of the Art and Challenges for the Future
Alvar M. Kabe
The Aerospace Corporation
10 April 2013
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Overview
Launch Vehicle and SpacecraftStructural Dynamics
The Load Cycle Process Structural Dynamic Models and
An indispensable test Loads Analyses and Why is not three sigma,
Why root-sum-square combinationscan under predict the truth, and
Why envelope functions are the wayto go
Why many spacecraft are notproperly qualified while subjected tounnecessary risk before launch
Future NeedsCourtesyofNASA
3
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3
They Made it Possible
DAlemberts Principle
(1717 1783)
Newtons Laws (1642 - 1726)
Hamiltons Principle
(1805 1865)
Lagranges Equations
(1763 1813)
d
dtmv
(t)( ) =
f(t)
m
jw
j(t
i)+ f
l(t
i)
l
j
j=1
n
wj (ti ) = 0
TV( )
t1
t2
dt+ WNonconservativet1
t2
dt= 0
d
dt
T w
j
Tw
j
+Vw
j
fdj = fNonconservative j
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Launch Vehicles and Spacecraft Structural Dynamics
During launch and ascent, a launch vehicle, its upperstage, and spacecraft experience severe structural loads
In addition, the launch system undergoes dramatic
configuration changes
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Design and Verification Process is Highly Complex
Numerous organizations involved Launch vehicle, engines, spacecraft, payloads, etc.
Numerous technical disciplines required Structures, dynamics, fluids, propulsion, controls, flight mechanics,
statistics, atmospheric sciences, etc.
Determination of dynamic loads and stresses involves complexmodels, analyses, and tests
Fully integrated launch vehicle/spacecraft system needs to beaddressed
Integrated system cannot be tested prior to flight Significant engineering judgment involved
Schedule and cost play a major role
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Loads Analysis Process
Structural loads are a function of dynamic properties of integratedlaunch/upper stage/spacecraft system
Design changes in one element can result in load changes in allelements
Modeling errors in one element can result in load prediction errors in allelements
Dynamic properties of each element will be a function of structuraldesign of each element
Therefore, design process has to be iterative
No single entity/organization has control over the loads analysisprocess
Hence, negative outcomes can be problematic
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The Load Cycle Process
Air Force Space Command, Space and Missile Systems Center Standard SMC-S-004,
Independent Structural Loads Analysis, 13 June 2008.
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The Load Cycle Process - Models
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Loads Analysis Model - Technical Disciplines
Structures
Structural Dynamics Guidance Analysis
Fluid Mechanics
Flight Software
...
ConfigurationDefinition
Finite Element
Model 1
Drawings
Static, ModeSurvey, other tests
Flight Data
Component 1
Dynamic Model
MassProperties
LTM Definition LTM Generation
Finite Element
Model 2
Drawings
Static, Mode
Survey, other tests
Flight Data
Component 2
Dynamic Model
MassProperties
LTM Definition LTM Generation
Structural
Model n
Drawings
Static, ModeSurvey, other tests
Flight Data
Component n
Dynamic Model
MassProperties
LTM Definition LTM Generation
LoadsAnalysis Model
and LTM
Autopilot
and Engine
Actuator
Definitions
Control System
Model
Aerodynamic
CoefficientsAerodynamic
Loading
PropellantDefinition
HydroelasticModel of
Fluids
SpacecraftModel
ConfigurationDefinition
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Component Mode Synthesis
Technique allows for significant reduction in size and subsequentcoupling of structural dynamic models
From millions of equations in a Finite Element model to a few thousand Modal truncation level depends on frequency content of excitation
Technique invented by Professor Walter Hurty at UCLA in 1965 Craig and Bampton (University of Texas) simplified computation of
constraint modes in 1968, hence Hurty/Craig-Bampton models
Benfield and Hruda (MMC) introduced Component ModeSubstitution in 1971
Couples Hurty/Craig-Bampton spacecraft models to launch vehiclemodels Provides for rigorous coupling of substructure damping properties
Most likely results in complex modes
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Component Mode Synthesis (cont.)
Hurty/Craig-Bampton model of a spacecraft
Benfield/Hruda coupled launch vehicle/spacecraft model
Non-Interface
Interface
xSC(t){ }
N
xSC(t){ }
I
=
SC NC
SC N
0 I
qSC(t){ }
xSC(t){ }
I
MSC =
I
SC
MSC
qI
MSC Iq
MSC II
DSC =
2n SC
0
0 0 II
KSC =
n2
SC
0
0 KSC
II
M** =I
SC
MSC
qI CE LV
I
LVI
T
CE T
MSCIq
I LV
D**
=
2n SC
0
0 2n LV
K**
=
n
2
SC
0
0 n2
LV
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Loads Analysis Model Validation and Verification
Accurate loads analysis models can only be achieved with testverification
Substructure models with over 1,000,000 degrees of freedom typical5-10 million degrees of freedom not unusual
Detail required to model complex hardware makes process costly Significant engineering judgment involved
To date, not a single analytical model of a complex structure has hadacceptable agreement with its mode survey test data prior to
adjustment Significant changes in loads from analytical to test-verified model
Static and other tests, which are performed to qualify structure toanalytically predicted loads, also used to adjust structural analysis models
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Mode Survey Tests, an Absolute Requirement
Performed to measure mode shapes, natural frequencies and damping Rigorous success criteria includes:
All modes within frequency range of interest measured Empirical modes, , satisfy orthogonality requirements of , where
Tests typically performed with: Hundreds of accelerometers Multi-shaker, broadband, uncorrelated random excitation
Measured acceleration and force time histories converted to frequencyresponse functions and modal parameters are extracted
Parameters used to adjust analytical finite element models with goals of
mT M
m = I T I
m
0.1
mT M
a = M
Mij 0.95 i= j
Mij 0.10 i j
n
measured n
analytical
n
measured 0.03
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Mode Survey Data and Model Correlation
Program
Number+of+
Measured+
Modes
Did+Model+
Need+
Adjustment
Program
Number+of+
Measured+
Modes
Did+Model+
Need+
Adjustment
Program
Number+of+
Measured+
Modes
Did+Model+
Need+
Adjustment
Program'1 9 yes Program'15 18 yes Program'32 23 yes
Program'2 20 yes Program'16'(11) 24''3 yes Program'33 12 yes
Program'3 47 yes Program'17 20 yes Program'34 14 yes
Program'4 45 yes Program'18 35 yes Program'35 6 no
Program'5A 35 yes ''''Subsystem'1 18 yes Program'36 13 yes
Program'5B 30 yes ''''Subsystem'2 9 yes Program'37 17 yes
''''Subsystem'1 9 yes Program'19 26 yes Program'38 25 yes
''''Subsystem'2 4 yes ''''Subsystem'1 Program'39 28 yes
''''Subsystem'3 4 yes '''''Conf.'1 5 yes Program'40 14 yes
Program'6 39 yes '''''Conf.'2 6 yes Program'41 8 yes
Program'7 16 yes '''''Conf.'3 6 yes Program'42 42 yes
Program'8 33 yes Program'20 43 yes Program'43 6 yes
Program'9 5 yes Program'21 9 yes Program'44 40 yes
Program'10 23 yes Program'22' 30 yes Program'45 14 yes
Program'11 13 yes Program'23' 6 yes Program'46 42 yes
''''Subsystem'1 6 yes Program'24 15 yes Program'47 28 yes
Program'12 11 yes ''Subsystem'1 8 yes Program'48 17 yes
Program'13 64 yes Program'25 15 yes Program'49 15 yes
Program'14 Program'26 23 yes Program'50 15 yes
''''Conf.'1 28 yes Program'27A 21 yes Program'51 18 yes''''Conf.'2 42 yes Program'27B 20 yes Program'52 55 yes
''''Conf.'3 39 yes Program'27C 6 yes Program'53 17 yes
''''Subsystem'1 8 yes Program'28 12 yes Program'54 16 yes
''''Subsystem'2 5 yes Program'29 11 yes Program'55 7 yes
''''Subsystem'3 10 yes Program'30 25 yes Program'56 7 yes
''''Subsystem'4 18 yes Program'31 16 yes Program'57 6 yes
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The Load Cycle Process Loads Analyses
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Launch Vehicles and Spacecraft Loads
Critical load producing events include:
LiftoffAtmospheric flight (Transonic, Max-q)
Engine ignitions and shutdowns
Jettison events
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The Aerospace Corporation 2013
Courtesy of NASA
Courtesy of ULA
Courtesy of SpaceX
Courtesy of NASA
Courtesy of ULA
Liftoff
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Liftoff Nonlinear transient phenomenon
Vehicle/launch stand interaction Produces critical loads for launch vehicleand spacecraft structure
Forcing function Thrust transients and differentials Ignition overpressure pulse
Ground winds
Gravity Launch stand interaction Dispersions
Modeling considerations Up to several thousand modes to 60 Hz Finite Element Models of substructurestransformed into component mode
models, hydroelastic models of fluids
Residual flexibility Coupled system damping
CourtesyofUS
AirForce
Launch Stand!Interaction
!
Ignition!Overpressure!
Gravity!
Ground!Winds!
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Liftoff Ignition Overpressure Pulse
Computational Fluid Dynamics (CFD) simulation (3.5 sec real time) 30.7 million cells; 120,000 cpu hours or 5000 cpu days
Aerospace Corporation Fluid Mechanics Dept.
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The Aerospace Corporation 2013
Atmospheric Flight Loads
CourtesyofNASA
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Atmospheric Flight
Static-aeroelastic Due to relative wind and non-zero angle ofattack, which varies slowly relative to the
fundamental mode frequency of the system
Turbulence/Gust Non-persistent wind features cause
changes in local angle of attack
Buffet Interaction between separated flow
turbulence, attached boundary layerturbulence, and shock wave oscillations
Autopilot-induced Maneuvering/steering Autopilot noise Mechanical noise (engine gimbal friction)
Other analyses include lack of wind persistenceand dispersions
Aerodynamic!Loading!
Relativ
e!Win
d!
Gust!Turbulence!
Maneuvering!Forces!
Thrust!
Buffet!
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Atmospheric Flight Loads
Computed for various times of flight Ten to twenty Mach numbers (times of flight) Transonic and maximum dynamic pressure (Max-q) times of flight tend to
be most critical
Transonic (maximum buffet) will yield design loads for significantportions of spacecraft
Maximum dynamic pressure will yield critical launch vehicle loadsand some significant spacecraft primary structure loads
Can occur during flight through jet stream and associated highturbulence (gust) loading
Loads analyses incorporate structural dynamic models, aeroelasticeffects, control system, atmosphere, and thrust effects
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Buffet Loads
Buffet event relatively long induration compared to othertransient events
Treated as steady state,ergodic random process
Objective of analysis is tocomputeprobability densityfunctions of response quantities,such as loads, from whichstatistical enclosure values can
be determined Mean, standard deviation Monte Carlo
!!y(t)
!!y(t)
!!y(t)
, !
, !
, !
, ! , !
Flight 1
Flight 2
Flight n
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Buffet Loads Analysis Approaches
Time domain - developed in 1984, first used for buffet analysis in 1988 Frequency domain - first large system implementation in 1988
Broussinos, P., and Kabe, A. M., Multi-Mode Random Response Analysis Procedure,Aerospace Technical Report SSD-TR-90-53, 1990.
I[ ] q(t){ }+ 2n q(t){ }+ n2 q(t){ } = [ ]
T
F(t){ }
Wind Tunnel TestF(t){ }
L2{ } = diag
1
2LTM Hx
*() T
Gf()
Hx ()
T
LTM T
d0
n
L
{ } = L2{ }meanL + kL{ }
Convert to PSDs
andCross PSDs
Retain TimeDomain
DefinitionL(t){ } = LTMT[ ] [ ] q(t){ }+ LTMV[ ] F(t){ }
L
2{ } =1
TL
2(t){ }dt0
T
mean square
values
root mean square (RMS)or
standard deviation
CourtesyofNASA
mean squarevalues
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Why a Mean Response in a Zero-Mean Process?
Of concern are the peak loads; therefore, we are interested in the statisticaldescription of peaks which have a mean
Envelope function of a stationary, narrow-band Gaussian process describedby Rayleigh probability density function
For broadband, multi-mode peak responses, Rayleigh assumption guaranteesreasonable conservatism
maxx
Rayleigh
Normal
Envelope
Max-Peak
!x(t)
!
2"
fx(x) =
1
2!"e
#x2
2"2
$
%&
'
()
f!x ( !x) = !x
!2e
" !x2
2!2
#
$%
&
'(
x(t)
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Turbulence/Gust Loads
Relatively small-scale, short-duration wind features
Loads analyses performed with fully integrated structural dynamic /aeroelastic / control system simulations - two types of excitation:
Synthetic profiles (e.g., 1-cos individual cases) Turbulence profiles extracted from measured winds (Monte Carlo)
Kabe, A. M., Spiekermann, C. E., Kim, M. C., and Lee, S. S., A Refined and Less Conservative
Day-of-Launch Atmospheric Flight Loads Analysis Approach, Journal of Spacecraft and Rockets,
Vol. 37, No. 4, pp. 453-458 (2000).
= +
Turbulence/Gust!
Aerodynamic!Loading! Turbulence!
Gust!
Control!Forces!
Thrust!
Relati
ve!Win
d!
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Lack of Wind Persistence in Gust Loads Analysis
Spiekermann, C. E., Sako, B. H., and Kabe, A. M., Identifying Slowly Varying and Turbulent
Wind Features for Day-of-Launch Flight Loads Analyses, Journal of Spacecraft and Rockets,
Vol. 37, No. 4, pp. 426-433 (2000).
AverageWavelength(ft.)
Turbulent Region
Slowly Varying Region
0 50 100 150
Lack-of-Wind-Persistence Time T (min.)
8000
6000
4000
2000
0
460 T
Rapidly and slowly varying wind components can be separated The boundary is a function of how far into the future (in minutes) one
wishes to predict
40 80 1200
Wind Magnitude (ft/sec)
40 80 1200
Wind Magnitude (ft/sec)
40 80 1200
Wind Magnitude (ft/sec)
10
20
30
40
50
Altitude(Kft)
10
20
30
40
50
Altitude(Kft)
10
20
30
40
50
Altitude(Kft)
60 Min 45 Min 30 Min
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Lack of Wind Persistence and Day of Launch Placards
60 min. before launch
45 min. before launch
30 min. before launch
Kabe, A. M., Spiekermann, C. E., Kim, M. C., and Lee, S. S., A Refined and Less Conservative
Day-of-Launch Atmospheric Flight Loads Analysis Approach, Journal of Spacecraft and Rockets,
Vol. 37, No. 4, pp. 453-458 (2000).
=! +!
=! +!
=! +!
Increasingturbu
lentcomponents
Increasingpersistentcomponents
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Statistical Distributions of Atmospheric Flight Loads
Turbulence/gust loads Gamma, Gumbel
pBN
(x,y)
!
BN(x,y)
1.0
1
3!
x x
yy
Buffet loads Rayleigh
Lack of wind persistence,other dispersions Bivariate Gaussian
pGB
(x;!,")!
GB(x;",#)
!= "1
#= 0.5
!=1
"= 0.5
!=1, "= 0.75
!=1, "=1
!= 1, "= 1
!=
"1
#= 0.5
!=1, "= 0.75
!=1, "= 0.5
x x
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Monte Carlo Combination of Atmospheric Flight Loads
Sako, B. H., Kabe, A. M., Lee, S. S., Statistical Combination of Time-Varying Loads,
AIAA Journal, Vol. 47, No. 10, October 2009.
!8 sec
Turbulence/Gust
Thrust Oscillation
Buffet
Totalij(t) = G
ij(t)+TO
ij(t)+ B
ij(t)
Gij(t)
TO
ij(t)
B
ij(t)
t t
t
t
i = Load Parameter (1-950)
j = Monte Carlo Run (1-3000)
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Monte Carlo Results Compared to Combination Equations
Sako, B. H., Kabe, A. M., Lee, S. S., Statistical Combination of Time-Varying Loads,
AIAA Journal, Vol. 47, No. 10, October 2009.
RSS G, TO,B( ) = G99/90( )2
+ TO99/90( )2
+ B99/90( )22
CLT G,TO,B( ) = G
Mean+TO
Mean+ B
Mean+ G
99/90G
Mean( )
2
+ TO99/90
TOMean
( )2
+ B99/90
BMean
( )22
ENV G,TO,B( ) = 2G +
2
TO+
2B+ G
99/90
2
G( )2
+ TO99/90
2
TO( )2
+ B99/90
2B( )
22
Underprediction
(Load Combination Equation / Monte Carlo) Ratio Histograms
Percen
tof950
Loads
Ratio
Central LimitEnvelope FunctionRoot Sum Square
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So Why Does an RSS Combination Under Predict ?
Frequency (Hz)
PSD
Re
sponses
Greater Likelihood ofPeaks Adding onto PeaksSynthesized Time Histories from Average PSDs
Time (Sec)
Gust
Gust and TO Average PSDs for a Load Parameter
TOFrequencySeparation
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The Aerospace Corporation 2013
Ignition and Shutdown Loads
CourtesyofUS
AirForce
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Ignition and Shutdown Dynamic Responses
Variability in ignition and shut-down transients will producevariability in dynamic responses
Expectedpeak positive andpeak negative responses that will occurduring future mission are of most interest
Thrust transients from past flights are samples from large family ofpossible profiles
Therefore, no guarantee that worst transients for a given responseparameter included in current dataset
If measured thrust transients are random samples of possibleexcitation, then an estimate of theprobability density function ofeach response quantity can be generated
Estimates of probability of non-exceedance (enclosure level) can then beobtained
Accuracy depends on number of samples, hence confidence limits mustbe computed
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Significant variability in thrust transients from engine to engine, andflight to flight
Frequency content is different for translation (sum: ) than rotation(difference: )
Engine Ignition and Shutdown Thrust Transients
f
E1
y
f
y
E2
E1 E2( )
E1+ E2( )
Response Spectra for Two Engines on Same Vehicle
E1+ E2
E1 E2
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Small Sample Enclosure and Confidence Levels
(Tolerance Bounds)
p(x)
x
p(x)
x
One thousand
99% enclosure
values
Ten-sampleSimulations
Twenty-sampleSimulations
90% confidence
One thousand
99% enclosure
values
First ten of
one thousand
simulations
First ten of
one thousand
simulations
90% confidence
Since the number of available thrust transients is usually small, wemust account for the small sample size
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Enclosure and Confidence Levels
Closed form solutions exist fornormaland Rayleigh distributions Numerical simulations can be used for other distributions
Example - eight samples, 99% enclosure with 90% confidence (99/90):1 10 100 1 10 100
1
10
1
10
Sample Size n Sample Size n
50%
90%Enclosure
90%
95%
99%
50%
90%
95% 99%
99%Enclosure
Kabe, A. M., Sako, B. H., Structural Dynamics, to be published.
NormalNormal
y99/90
= y + 3.783
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So, When Does 3 Not Equal Three Sigma ?
NormalOne-sided
Upper Tolerance Bound
NormalTwo-sided
Tolerance Interval
Rayleigh
Tolerance Bound
33330.5
Small Sample Size Tolerance Bound Factors for a Sample Size of 11(for 90 and 50 percent confidence levels)
Normal Rayleigh
0.9987 0.9973 0.9889
0.9987 / 90 k= 4.4032
0.9987 / 50 k= 3.1112
0.9973/ 90 k= 4.4772
0.9973/ 50 k= 3.3208
0.9889 / 90 k= 3.7555
0.9889 / 50 k= 3.0465
k
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The Load Cycle Process Structural Qualification
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Structural Qualification and Acceptance
Adequacy of structure defined via design qualification and hardwareacceptance Designs qualifiedto account for part-to-part variability due to assembly
procedures, build-to-build tolerances, and uncertainties in redundant load
paths
Flight hardware acceptedto screen out poor workmanship or preclude ahazardous condition such as a leak in a pressure vessel
Verification requires consideration of all potential failure modes andall potential load conditions
Potential failure modes include: detrimental deformation, material yield,ultimate failure, structural collapse, buckling, fatigue, delamination
Load conditions include: quasi-static and dynamic launch loads, acousticenvironment, pressure, temperature, gravity, handling loads
AIAA S-110-2005, Space Systems Structures, Structural Components, and Structural AssembliesStandard, July 2005.
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Spacecraft Static Strength and Base Shake Tests
Static Strength Testing Flight/flight-like structure (system and component level) subjected toenvelope of: operational loads (including pressure and thermal), load
cycle computed statistical loads, and design requirements
Strains, deflections, and applied loads measured Known margin included (e.g., limit to ultimate, temperature effects)
Base Shake Testing Flight/flight-like structure (less propellants) subjected to unidirectional,
swept sinusoidal base excitation
Sine environment derived from computed launch vehicle/spacecraftinterface responses and/or measured flight responses
Base and system response accelerations measured and used to limittest article responses
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Issues with Base Shake Tests
Spacecraft modes of vibration on a shake table will be different thanwhen coupled to a launch vehicle
Merged launch vehicle/spacecraft modes form coupled system modes Shake tables do not replicate interface impedance of a launch vehicle
Spacecraft dynamic properties on a shake table are not equivalent tothose on seismic mass, which are used in mode survey tests
Data from two systems shows 20% difference in fundamental modefrequencies when measured on shake table vs. fixed to a seismic mass
Properties on shake table are those of the satellite/shake table system Assumption inherent in loads analysis models (Hurty/Craig-Bampton
model) is that modal coordinates (mode shapes) are relative to a fixed/cantilevered interface
Rigid body motion and interface stiffness are accounted for by constraintmodes
Kabe, A. M., Perl, E., Limitations of Base Shake Analysis and Testing of Flight Configured Spacecraft,
12th European Conference on Space Structures, Materials & Environmental Testing, ESA/ESTEC, March 2012.
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Issues with Base Shake Tests (cont.)
Only translational motions are applied at base, one axis at a time Launch vehicle/spacecraft system, however, vibrates simultaneously in all
degrees of freedom during flight
Some spacecraft modes will be difficult, if not impossible, to excitethrough base excitation
For example, to excite fundamental torsional mode requires offset betweenspacecraft center of gravity and shear center that leads to modal gains thatdo not cancel
Coupled launch vehicle/spacecraft vibration is broadband, notsingle frequency sinusoidal, as in base shake tests
Vibration of spacecraft undergoing sinusoidal base shake will be significantlydifferent than in flight
Flight response involves significant, simultaneous vibration in many modesKabe, A. M., Perl, E., Limitations of Base Shake Analysis and Testing of Flight Configured Spacecraft,
12th European Conference on Space Structures, Materials & Environmental Testing, ESA/ESTEC, March 2012.
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Issues with Base Shake Tests (cont.)
Because of stroke limitations, shake tables are not capable ofinducing steady-state acceleration experienced in flight
Loads must be achieved dynamically, which causes over testing some partsto achieve proper minimum loads elsewhere
Thus, spacecraft must be designed to survive test, in addition to flight
Results in weight penalty and additional cost to spacecraft program
Most internal loads experienced during base shake testing cannot bemeasured and must be established with analytical model
Experience indicates that within the frequency range of base shake testing(up to several hundred Hz), these models are highly uncertain, even when
adjusted to mode survey test data
Kabe, A. M., Perl, E., Limitations of Base Shake Analysis and Testing of Flight Configured Spacecraft,
12th European Conference on Space Structures, Materials & Environmental Testing, ESA/ESTEC, March 2012.
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Issues with Base Shake Tests (cont.)
Spacecraft test article will most likely not be in flight configuration Might not include actual spacecraft launch vehicle adapter Typically, propellants not included in tanks because of safety and
contamination concerns
This leads to a test article with different dynamic properties relativeto a flight-configured spacecraft Makes it very difficult, if not impossible, to induce proper load levels in
many (most) parts of a spacecraft
Invariably requires supplemental static strength tests
Kabe, A. M., Perl, E., Limitations of Base Shake Analysis and Testing of Flight Configured Spacecraft,
12th European Conference on Space Structures, Materials & Environmental Testing, ESA/ESTEC, March 2012.
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Issues with Base Shake Tests (cont.)
For operational launch vehicles, base excitation environments derivedfrom flight data
Measured with limited number of accelerometers near launch vehicle/spacecraft interface
Local responses, including warping of interface area, lead toartificially high environments when rigid interface assumption used Can result in an over test of spacecraft
Assumption that measurements are an input at base of spacecraft istechnically not correct
Measurements are those of the response of a coupled system Spacecraft participates in producing the responses
Kabe, A. M., Perl, E., Limitations of Base Shake Analysis and Testing of Flight Configured Spacecraft,
12th European Conference on Space Structures, Materials & Environmental Testing, ESA/ESTEC, March 2012.
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Impossibility of Inducing Flight-Like Loads in a Base
Shake Test
Coupled launch vehicle/spacecraft responses (e.g. accelerations,loads) to broad band excitation computed
Computed interface accelerations used to establish base excitationper widely used approach
Coupled system responses compared to base excitation results
Tuttle, R. E., Lollock, J. A., Assessment of Base Drive Analysis and Test for Complex Systems,
Spacecraft and Launch Vehicle Dynamic Environments Workshop, The Aerospace Corp., 2012.
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Future Needs
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Future Needs
Increased rigor in statistical analysis, with Monte Carlo analysesoffering a way forward Specify enclosure/confidence levels, not number of standard deviations
Increase complex spacecraft mode survey test limits to 75-100 Hz Currently 50-60 Hz
Improved techniques for adjusting analytical models to better matchmode survey test data
Despite progress, its still trial and error
Reduce/eliminate use of base shake tests of large, flight configuredspacecraft Static and subsystem tests offer proven way forward
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In Closing, a Few Photographs
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Launch Vehicles
Gemini-Titan II 1965
Cou
rtesyofNASA
First Space Shuttle 1981
Cou
rtesyofNASA
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Launch Vehicles
Titan IV
C
ourtesyofULA
Delta IV
Courtesy
ofUS
AirForce
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Launch Vehicles
CourtesyofSpaceX
Falcon 9
C
ourtesyofULA
Atlas V
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Earths City Lights
Courtesy of NASA
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