N A S A C O N T R A C T O R

R E P O R T

COTTO-JI

N A S A C R - 2 4 8 8

ANALYTICAL AND EXPERIMENTAL STUDY

OF THE EFFECTS OF WING-BODY

AERODYNAMIC INTERACTION ON

SPACE SHUTTLE SUBSONIC FLUTTER

Richard R. Chipman and Frank J. Ranch

Prepared by

GRUMMAN AEROSPACE CORPORATION

Bethpage, N.Y. 11714

for Langley Research Center

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION • WASHINGTON, D. C. • J A N U A R Y 1975

https://ntrs.nasa.gov/search.jsp?R=19750006743 2020-05-27T03:35:34+00:00Z

1. Report No. O/IQQ 2' Government Accession No.

4. Title and subtitieANALYTICAL AND EXPERIMENTAL STUDY OFTHE EFFECTS OF WING BODY AERODYNAMIC INTERACTION ON SPACE SHUTTLE SUBSONIC FLUTTER

7. Author R ichard R> chipman and Frank J. Rauch

9. Performing Organization Name and Address

Grumman Aerospace Corporat ionBethpage, NY 11714

12. Sponsoring Agency Name and Address

National Aeronaut ics and Space Admin is t ra t ionWashington, D.C. 20546

3. Recipient's Catalog No.

5. Report Date

JANUARY 19756. Performing Organization Code

8. Performing Organization Report No.

10. Work Unit No.

506-17-32-0311. Contract or Grant No.

NAS1-10635-1813. Type of Report and Period Covered

Contractor Report14. Sponsoring Agency Code

15. Supplementary Notes

Final report.

16. Abstract

To determine the effect on flutter of the aerodynamic inter-action between the Space Shuttle bodies and wing, l/80th-scalesemispan models of the orbiter wing, the complete Shuttle andintermediate component combinations were tested in the NASALangley Research Center 26-inch Transonic Blowdown Wind Tunnel.Using the doublet-lattice method combined with slender body theoryto calculate unsteady aerodynamic forces, subsonic flutter speedswere computed for comparison.

Using calculated complete-vehicle modes, flutter-speed trendswere computed for the full-scale vehicle at an altitude of 15,200meters and a Mach number of 0.6. Consistent with findings of themodel studies, analysis showed the Shuttle to have the same flut-flutter speed as an isolated cantilevered wing.

17. Key Words (Suggested by Author(sl)

Space Shuttle FlutterCalculated Subsonic Flutter SpeedsCalculated Full-scale Flutter Spds1/80-scale Semispan Model

19. Security Qassif. (of this report)

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For sale by the National Technical I nformation Service, Springfield, Virginia 22151

ANALYTICAL AND EXPERIMENTAL STUDY OF THE EFFECTS.OF WING-BODY AERODYNAMIC INTERACTION ON

SPACE SHUTTLE SUBSONIC FLUTTER

By Richard R. Chipman and Frank J. Rauch

SUMMARY

To determine the effect on flutter of the aerodynamic interaction "be-tween the Space Shuttle bodies and wing, l/80th-scale semispan models of theorbiter wing, the complete Shuttle and intermediate component combinationswere tested in the NASA Langley Research Center 26-inch Transonic Slowdown WindTunnel. Using the doublet-lattice method combined with slender body theory tocalculate unsteady aerodynamic forces, subsonic flutter speeds were computedfor comparison.

Aerodynamic interaction was found by test and analysis to raise theflutter speed in some configurations while lowering it in others, such thatthe flutter speed of the complete Shuttle at M = 0.7 was the same as that of acantilevered isolated wing. Although at Mach numbers not greater than 0.75 pre-dicted speeds correlated to within 6% of those measured, rapid deteriorationof the agreement occurred at higher subsonic Mach numbers, especially on themore complicated configurations. Steady-state pressure distributions were alsomeasured and computed, revealing similar trends.

Using calculated complete-vehicle modes,, flutter-speed trends were com-puted for the full-scale vehicle at an altitude of 15,200 meters and a Machnumber of 0.6. Consistent with findings of the model studies, analysis showedthe Shuttle to have the same flutter speed as an isolated cantilevered wing.

INTRODUCTION

The flutter characteristics of an aircraft component can be affectedby aerodynamic interaction between it and other proximate components. Classicexamples, when both components are lifting surfaces, are T-tail flutter(Ref. l) and wing-tail flutter (Ref. 2). To predict subsonic aerodynamicforces arising in such configurations, the kernel-function method of Watkins(Ref. 3) and the doublet-lattice method of Rodden (Ref. k) and of Stark (Ref. 5}

were developed and extended to handle multiple planar and nohplanar sur-faces. The success obtained in applying these theories to wing-tail configura-tions is recorded by Sensburg (Ref. 6), Mykytow (Ref. ?), and Triplett(Ref. 8). Similarly, Chipman used the doublet-lattice flutter analysismethod to correlate successfully with wind tunnel data for pairs of closelyspaced wings such as might be found on the fly-back-booster Space Shuttle con-figurations (Ref. 9).

The aerodynamic forces arising from the interaction between bodies andlifting surfaces can also be significant. In the steady-state case, methodsfor computing these forces were developed by Woodward (Ref. 10) and Hess (Ref.11). Rodden and Giesing (Ref. 12) have combined slender-body theory with thedoublet-lattice method to solve the unsteady problem in subsonic flow.

The current Space Shuttle concept features four large, flexible bodies:the orbiter fuselage, the external tank (ET), and the two solid-rocket boosters(SRB's). The proximity of these bodies to the orbiter wing admits the possi-bility of a change in vehicle flutter boundary as a result of aerodynamicinteraction. Consequently, a two-phase study was initiated to obtain analyticaland experimental confirmation of this contention. Reference 13 reports thefirst phase of this work: A preliminary design - the Grumman G III - of the •parallel-burn, Space Shuttle concept was analyzed using Rodden's method (Ref.12) to determine the anticipated effect of wing-body aerodynamic interaction onflutter. This report describes the second phase: Comparisons of experimentallyand analytically determined wing-body, interaction flutter effects are made sub-sonically for a model of a recent design to assess the applicability of Rodden'smethod to the Space Shuttle.

In the present study, l/80th-scale semispan models of the Shuttle weretested for flutter in the Mach number range 0.6 to l.it. Additionally, steady-state pressures were measured on the wing. Results were correlated with flutterspeeds and pressures predicted using Rodden's doublet-lattice/interaction-panel method. Various combinations of vehicle components'- wing alone; wingand orbiter fuselage; wing, fuselage and external tank; and complete Shuttle -were studied.

Because the analyses of Phase I indicated that forces arising from bodyflexibility did not appreciably alter the Space Shuttle flutter speed, themodel bodies in the present investigation were designed to be rigid. Exceptfor two configurations in which flexible SRB attachments were used, only thewing was flexible.

A final check of interaction flutter effects on the current Shuttlewas made by calculating flutter speeds using calculated full-scale vehiclemodes which included flexibility on all the bodies.

LIST OF SYMBOLS

b Reference semichord = 0.0572 mr

c Local chord

CT Lift coefficientL

C Life curve slopeLa

C Moment curve slopea

C Pressure coefficient

AC Coefficient of the net pressure difference acrossa surface or "body

K, Net pressure coefficient distribution on the slender

body elements

AC I Net pressure coefficient distribution on the inter-' i action panels

NW

[Dww]

Net pressure coefficient distribution on the wingpanels

Downwash on the interaction panel collocation pointscaused by unit pressures on the interaction panels

Downwash on the interaction panel collocation pointscaused by unit pressures on the wing panels

Downwash on the wing collocation points caused byunit pressures on the interaction panels

Downwash on the wing collocation points caused byunit pressures on the wing panels

ET External tank

f Frequency

Downwash on the interaction panel collocation pointsdue to unit pressures on the slender body elements

Downwash on the wing collocation points due to unitpressures on the slender body elements

LIST OF SYMBOLS (continued)

h Modal deflection

k Reduced frequency = bu/U

M Mach number

m Mass of exposed wing = 0.0525 kg

q Dynamic pressure

R Slender body radius

s Exposed semispan

SRB Solid rocket booster

U^ - Free stream velocity

V Flutter speed

v Volume of frustum of a cone encompassing the exposedW wing = 0.00133 m3

V-g-f Velocity-damping-frequency

W I Prescribed downwash at the interaction panel colloca-•I tion points

W I i Prescribed downwash at the wing collocation points

x Streamwise coordinate

Ax Distance that the center pressure lies aft of the localleading edge

y Spanwise coordinate

z Vertical coordinate!

a Angle of attack or incidence

CL, Body incidenceE>

a Wing incidence

a Ratio of air density in tunnel to that at sea level

LIST OF SYMBOLS (continued)

Mass density ratio =PSLVW

SL Density of air at sea level

w Circular frequency

o> Reference frequency = UjO Hz

( )' Derivative in x'direction

( )" Second derivative in x direction

TEST APPARATUS

Models

Semispan models of the current Shuttle design shown in Figure 1 werebuilt for flutter testing. They consisted of rigid fairing replicas of theor"biter half-fuselage, the half external-tank (ET), and the solid rocket"booster (SRB) in proper proximity to a flexible orbiter wing. As can be seenin Figure 2, the wing is a double delta with an inner-panel leading-edge sweepof 79° i an outer-panel sweep of 5°, and an aspect ratio of 2.1. The crank inthe leading edge occurs at a station 25% of the distance from the root to thetip. The wing model consisted of aluminum spanwise-tapered core plates withcutouts, designed to match the torsion-to-bending frequency ratio of the full-scale wing.

Figure 3 shows the wing in various stages of assembly. On the core,balsa wood was affixed and shaved to a 12% maximum thickness-airfoil shaperepresentative of the Shuttle. Wing camber and twist were not modelled; themodel section was made symmetrical. To eliminate structural coupling, the rigidgeometrically scaled fairings representing the fuselage, ET, and SRB com-ponents were fastened rigidly to a splitter plate located at the mid-vehicleplane. This assembly is shown in Figure h. A second splitter plate was madewhich could be located at the wing root when the body fairings were removed tostudy flutter of the isolated wing end-plated at its root.

Two sets of flexible connectors were fabricated, each of which could besubstituted for the rigid fasteners linking the SRB to the ET. With theseflexures, SRB pitch and heave modes could be attained. One of the sets was

designed such that the frequencies of these modes in model scale were equal tothose of 8KB pitch and vertical bending of the nominal Shuttle design. Withthe second set, the frequencies were chosen to be closer to wing torsion sothat the likelihood of interaction would be greater. Table I summarizes thefrequencies of these and the primary wing modes.

A rigid wing with 38 pressure taps was constructed to measure steady-state pressures. Provisions were made on the mounting plate to pitch eachconfiguration either 0 of -3°. With pressure data from each of these twoangles of attack, the primary effects of wing thickness could be removedby subtraction of the two pressure distributions.

Instrumentation on Models

Each flexible wing was instrumented with strain gages which, throughtheir response to bending and torsional motions, were used to detect the onsetof flutter and to measure flutter frequency. Additional gages were mountedon the SRB flexures to detect SRB excitation.

The pressure taps of the rigid pressure wing were connected to trans-ducers that measured the difference between the manifold pressure and thepressure on the wing surface.

Wind Tunnel

Flutter and pressure tests were conducted in the NASA Langley ResearchCenter 26-inch Transonic Slowdown Wind Tunnel for a Mach number range of 0.6to l.U. During the flutter runs, an oscillograph was used to record thefollowing items:

• Output of model strain gage and magnetic coil circuits

• Total test section pressure (P )

• Settling chamber temperature (T )

• Static test section pressure (P )s

• Reference trace used in determining P , P and T trace deflection& o' s o• Movie camera correlation trace.

A high speed movie camera was used to obtain a visual record of model behaviorduring each flutter run.

For each pressure run, the following information was placed on mag-netic tape:

• Output of pressure gage circuits that record wing pressure data

• Total test section pressure (P )

• Static test section pressure (P0)o• Temperature (TO)

• Manifold pressure (?>,)•

TEST PROCEDURES AND DATA REDUCTION

Vibration Survey

Prior to wind-tunnel testing, each wing model was subjected to avibration survey. Modal shapes, frequencies and damping coefficients were.determined for one model, and the subsequent models were checked for likenessof frequencies and damping coefficients. Noricontacting excitation and mea-surement systems were utilized to assure distortion-free mode shapes andfrequencies. The output of a noncontacting inductance-type pickup was con-ditioned, and displayed on a vacuum tube voltmeter and oscillograph to extractmode shapes and damping data. During the survey, the models were rigidlyclamped at the root fto simulate the tunnel installation.

Shown in Figure 5 are the locations of the concentrated mass pointsand the points at wriich the modal data were measured. Table II gives thecoordinates of these points and the calculated masses. Figure 6 presentsthe measured mode shape data in tabular and graphic form.;

Similarly, for each set of flexible attachments, modal data were mea-sured on the SRB for the two lowest frequency modes. These modes are pre-sented in Figure 7. Because of the inherent flexibility of the SRB model,the modes include a significant amount of bending as well as the intendedpitching and heaving motions. The frequencies of these and the measured wingmodes are summarized in Table I.

Flutter Tests

Before each tunnel run, the models to be used were visually examinedfor signs of damage due to previous runs. In addition, after being installedin the tunnel, the wing model was excited, and oscilloscograph records of thestrain gage outputs were monitored to check frequencies and dampings of thefirst four modes.

On the basis of the results of previous runs, a desired tunnel opera-ting "path" was selected for the run. This path was followed until eitherthe tunnel air supply was exhausted or.flutter was detected visually. Severalruns following different paths were made on each configuration to determinea boundary of flutter speed vs Mach number.

As shown in Table III, flutter boundaries were determined for six con-figurations: the isolated wing endplated at its root, the orbiter (wing andfuselage), the orbiter and external tank, the full-up. Shuttle (orbiter, tankand SRB) with rigid interstage attachments, the full-up Shuttle with thenominal flexible attachments and the full-up Shuttle with a second set offlexible attachments.

Flutter points were initially determined from the oscillograph recordsand visual observations during the run, and then either confirmed or adjustedby subsequent examination of the movie data. The flutter point was taken asthe point during the run where the model circuit traces indicated a frequencyregularity and/or significant increase in the model vibrational amplitude.Tunnel data converted to the parameters Mach number, air density ratio, anddynamic pressures are presented in Table IV.

For each flutter point (or significant stable point), the Mach numberand dynamic pressure are presented in Figure 8. In addition to the datapresented in Table IV, several non-flutter points during each run were studiedto determine the tunnel path taken. These paths are plotted in Figure 8 andhelp define the suggested flutter boundaries indicated.

Pressure Tests

After the rigid pressure wing was installed in the wind tunnel to-gether with appropriate body models, the-pressure lines were connected totransducers and calibration was performed.

During each run, the tunnel was programmed to quickly attain a pre-chosen combination of Mach number and dynamic pressure. So that the flowwould have time to stabilize, this point was maintained for 3 seconds beforeany data were collected. Pressure was sampled five times at each orificeand then a constant Mach number path was followed to a second, higher dynamicpressure at which the pause and sampling were repeated. The data were storedon magnetic tape and later reduced by digital computer.

For each of the three configurations'noted in Table III, pressuremeasurements were made at M = 0.6, 0.8, 1.0 and 1.2, once for the vehicle atan angle of. attack equal to 0° and once for the vehicle at a -3° angle ofattack. The data were automatically plotted as chordwise distributions of Cat both the upper and lower surfaces at each of the four spanwise stations,y/s = 0.1, O.U, 0.7 and 0.9. Only'the subsonic data were further reduced andstudied for this report.

After inspecting the data and-eliminating points where gages wereeither clogged, damaged, or otherwise unreliable, it was determined that dataon the outer portion of the upper wing surface were largely unusable butthat the lower surface data were acceptable at most locations.

To determine a pressure difference across the wing due to the angle ofattack, one would normally subtract the lower surface pressure from theupper surface pressure measured at the angle of attack, and from this incre-ment subtract a similar increment measured at zero incidence. This computationcan be expressed:

ACP (a)= (C

P, u(a)-CP,*

(a)>- (CP,u (0) ' °p,* <°»

where AC (a) is the pressure difference at angle of attack,

C is the pressure coefficient on the upper surface,

and Cp,£ is the pressure coefficient on the lower surface.

Since the outer upper-surface pressures were unreliable, an alternate proce-dure was used for y/s =0.7 and 0.9 and checked at y/s = O.k. If a flow obeysthe assumptions of linear potential theory,.the effect that a small angle ofwing incidence has on pressure is equal and opposite on opposite sides of athin airfoil - even in the presence of geometric asymmetries. Thus, one coulddetermine the total pressure difference at angle of attack by the calculation:

ACp (a) = (Cp)U(a)-Cp5u.(o))-(Cp>£ (a) - C^ (o»

= -2 (Cp,l (o)--Cp,£ (°)}-

The validity of this procedure can be appraised by inspecting Figure 9,which' shows that the angle of attack produces very similar measured pressurechanges on the upper and lower surfaces at the inboard chord where y/s = O.U.Since the inner chord (y/s = O.l) is close to the reflection plane and thebodies, viscous effects would be expected to be'larger here than on the outerchords. Consequently, the pressure difference was computed directly for thischord and use was not made of the above approximation. The resultant pres-sure distributions are shown in Figures 22 through 25 and discussed in a sub-sequent section.

THEORY

Aerodynamic Idealization

Using the doublet lattice method of Ref. 12, aerodynamic influencecoefficients were calculated at various subsonic Mach numbers for the configu-rations studied. As depicted in Figure 10, the aerodynamic idealization con-sisted of panels of doublets modeling the wing and the idealized surfaces ofthe bodies in the vicinity of the wing - where interaction effects would beexpected to be largest - and of axial doublets representing the bodiesthemselves.

The strengths of the axial doublets were determined separately, usingslender body theory:

RAC = 2^Ro >]

2bR h \ + 2Tr--i J R 'h + R h1

2 o (1)

while the strengths of the body surface doublets were determined jointly withthe wing doublets. .To introduce coupling between the two solutions, theboundary conditions.to be satisfied by panel doublets were modified bythe downwash created by the axial doublets:

.ACp,w

AC

D D _w,w w,I

DI ,vDI , I

-1ww •w,B

ACp,B

(2)

Ignoring the orbiter fin for the purposes of this study, only the wing, theorbiter fuselage, the external tank.and the solid-rocket booster were modeled.

Vehicle Idealization Studies

Wing. - To determine a suitable aerodynamic grid for the analyses,pressure distributions'were calculated for several idealizations of the Shuttlewing double-delta platform as shown in Figure 11. Pressures, spanwise liftdistributions, and generalized aerodynamic forces were calculated for a zero-frequency pitch mode corresponding to a wing incidence of one radian at M := 0.7'Additionally, a control surface rotation mode was studied.

10

As can be seen in Figure 12, grids 1 and 2 give unsatisfactorypressure calculations on the highly swept inner wing panels . Whereas erraticpressures arise from these idealizations, grid 3 results in smooth distribu-tions .

Chordwise and spanwise variations in the number of boxes in grid 3 wereperformed as indicated in Figure 11. A major geometrical feature of the wingis leading crank located at 25$ of the exposed semispan. As shown in Figure13,.the panel distribution of eight chordwise, three spanwise inboard of theleading edge crank, and seven spanwise outboard of this crank (grid 3.H) givessmooth converged pressures for both wing pitch and control surface rotation;thus, it is used to analyze the full-scale Shuttle with complete vehicle modes.When control-surface modes need not be represented, fewer lattice boxes are re-quired. Hence, grid 3-5» which has only six chordwise panels and gives thesame pressure distribution for a pitch mode as does grid 3.^, is used as theaerodynamic idealization for the wing. Smooth oscillatory pressures calculatedusing this grid are shown in Figure lit for pitch about the wing apex. Noticethe growth of the imaginary part of the pressure as frequency is increased.

Bodies.. - The representation of the orbiter fuselage, the external tank,and the solid-rocket booster consisted of axial elements, strengths of whichwere calculated by slender body theory, and of panels of doublets applied" toidealized surface of each body to account for aerodynamic interaction betweenthe wing and bodies. In Ref. 13, idealization studies were made to determinea distribution of slender body elements and interference panels that would giveconverged generalized aerodynamic forces for the flutter study of a Shuttleconfiguration. Essentially the same modeling was used in this study.

To appraise effect of the presence of the orbiter fuselage on liftdistribution of the wing, steady state lift calculations were made for wing inpitch with various root conditions and compared to the lift obtained withpitched orbiter fuselage in place. In all, five conditions were studied:

(1) Reflection plane located at the wing root, endplating it asin the tested wing alone configuration

(2) Reflection plane located at the vehicle centerplane; the root isfree

(3) Reflection plane located at the vehicle centerplane; gap betweenthe root and the centerplane is filled with a rigid plate at zeroincidence

(H) Root condition same as in case 3 but rigid plate is at the sameangle of incidence as the wing

(5) Pitched fuselage is modelled by slender body elements andinterference panels.

Figure 15 shows resulting lift distributions. In case (2), the liftcorrectly falls to zero at the root because of the gap. Case (3) has a higherwing lift than (2) because the plate supports lift but has lower lift than (l)because the reflection plane in case (l) is more effective than the plate. When

11

the plate is pitched, it gives rise to even higher lift as shown in case (U).When in case (5) the fuselage is modelled, the lift on the wing is logicallymore than in case (3) and less than in case (l) because the side wall of thefuselage acts as an abbreviated reflection plane.

Because the bodies on the Shuttle have blunt or truncated aft ends andthe orbiter has the rather small ratio of length to diameter of 5-^* theadequacy of slender-body theory is questionable. To resolve this question,calculations of the lift and moment of an orbiter fuselage were made and com-pared with available unpublished low-speed wind tunnel test data. Al-though this fuselage is not the same design as the current Shuttle, it has al-most -the same length to diameter ratio (5-6) and the same truncated aft end.In this analytical idealization, this latter feature was represented by assum-ing the body radius to remain constant at the end rather than returning tozero. A comparison of theoretical and experimental aerodynamic coefficientsis given in Table V. The correlation is fairly good: The lift coefficientsagree within 7% and the moment coefficients within 15%.

Pressure Calculations

To compare with the distributions measured on the various configura-tions at a constant angle of incidence to the flow, steady-state pressureswere calculated on the wing for these cases. In this simple case, the slender-body pressures of equation 1 reduce to:

ACp,B (3)

Flutter Model Analyses

Using modes measured in the vibration survey, flutter solutions-weredetermined for the models tested. For configurations 2 through k of Table III,the rigidity of the bodies causes the slender body pressure to vanish so thatequation 2 becomes:

ACp,w

AC

D D _w,w w,I

DT D_I,w 1,1

-1ww

12

In configurations 5 and 6, the SRB:is allowed to heave and pitch and,hence, gives rise to slender-body pressures as calculated by equation 1. Sincethe lower frequency mode in each case has very little curvature, the h" term iszero for this mode. • • -.. . .

In configuration 1, the-base case, there are no bodies to give riseto .aerodynamic interaction. In this case"equation h is-further simplified to:

AC Dw,w

-1 Ww

(5)

Full-Scale Flutter Analyses

Calculated symmetric normal modes for the Space Shuttle were obtainedfrom Ref. lh. Table VI lists the frequencies of the lowest 30 modes togetherwith a brief description of each. Only 20 of these modes, as designated in.this table, were used in a flutter analysis of the'full-scale vehicle. Themodes ignored were those that were aerodynamically inactive, .such as longi-tudinal modes. Since the modes entail body motions of a general type, allthe terms in equations 1-and 2 are needed in calculating the unsteady pres-sures and, hence, the associated generalized aerodynamic forces required inthe flutter analysis.

MODEL FLUTTER RESULTS

Analytical Results

Shown in Figures 16 and IT are representative analytical plots of damp-ing and frequency as a function of airspeed for the .isolated wing and theShuttle. From V-g-to plots such as these, flutter-speeds and frequencies weredetermined for the various configurations at several values of air density overa range of subsonic Mach numbers. The flutter mechanism is the coupling of thefundamental wing bending mode with the first wing torsion mode. Although fivemodes were used in the flutter calculations for all the various configurations,sample calculations on the isolated wing made including only these two modesgave rise to the same flutter speeds.

For the isolated wing and for the orbiter, Figure 18 shows the result-ing trends in flutter speed. Both .configurations have basically the sametrends: Flutter speed varies little with M until high subsonic Mach numbersare reached, where the variation experienced is strongly dependent on' airdensity. In correlation with test data, consequently, the density ratio mustbe matched by analysis. Flutter frequency exhibits a similar trend but is notas dependent on air density.

13

Flutter Boundaries at Test Conditions

Shown in Figure 19 (a) are the flutter speed and flutter frequencyindices measured on the isolated wing. Also shown are the analytical resultsobtained at the test conditions. At Mach numbers below 0.9, the analyticalflutter speed prediction is at most 6% conservative, while the predictedflutter-frequencies are 5% low at worst. Hence, the correlation in this basecase is excellent.

Since the minimum measured flutter speed is only k 1/2% lower than themeasured flutter speed at M = 0.65, the wing can be said to have almost notransonic dip. Because the recovery in flutter speed is sharp, the level ofthe supersonic flutter speed was not ascertained.

The flutter speed and flutter frequency indices determined for theorbiter configuration are presented in Figure 19(b). Both experiment and anal-ysis show that subsonic flutter speed is higher than that of the wing alone.The accuracy of the predicted flutter speeds is still within 6% and that of thefrequencies is within Q% . Although the transonic dip of 1% is slightly morethan occurs on the wing alone, it is still very shallow.

As can be seen in Figure 19(c), agreement between test and analysisbegins to deteriorate on the orbiter/tank configuration. Up to a Mach numberof 0.8, the flutter speeds correlate within J%, but at M = 0.9 the analysisis 11$ conservative.

On the full-up Shuttle configuration with rigid SRB attachments, theMach number at which correlation, deteriorates is seen in Figure 19(d) to shiftlower. At M = 0.75, the analytical flutter speed is only 6 1/2% low; but atM =' 0.8 and 0.85 the discrepancies are 9% and 13%, respectively.

Figures 19(e) and 19(f) present the flutter boundaries for the full-upShuttle with flexible SRB attachments. At any given Mach number and densityratio, the flutter speeds calculated by analysis were the same for these twoconfigurations as for the Shuttle with rigid SRB attachments. Hence, whenadjusted to the test conditions, the analytical flutter speed boundaries forthese three configurations are roughly the same. The test boundaries, however,are lower when flexible attachments are used.

Comparison of Flutter Boundaries

To compare realistically the flutter boundaries of the various config-urations, it is necessary first to adjust the test data to a constant airdensity, since, in the blowdown wind tunnel, different operating points in gen-eral correspond to different densities. This scaling was made by multiplying

the test flutter speed "by the ratio of the analytical flutter speed at thedesired density to that at the test point density. To minimize the amount ofscaling to be made, the "boundaries were adjusted to a typical tunnel densityof three times that at sea level. Figure 20 shows the resulting test flutterboundaries.

At moderate subsonic Mach numbers (M .<_ 0.75),- the orbiter flutter speedis 1% higher than that of the wing, the orbiter/tank flutter speed is 3%higher than that of the wing and the full-up Shuttle is only 1% higher. Noappreciable transonic dip occurs on any configuration. Although all config-urations show a sharp flutter speed increase as Mach one is approached, thisrecovery occurs at increasingly lower Mach numbers as the tank and the SRBare added to the vehicle. This trend causes the order of the flutter speedsto change at the high subsonic Mach numbers so that this order becomes: wingalone lowest, orbiter higher, orbiter/tank higher still, and complete Shuttlehighest. This shift in the recovery can probably be attributed to higherlocal Mach numbers on the wing in the vicinity of the tank and SRB when thesecomponents are added to the orbiter.

The analytical flutter boundaries at the same air density of threetimes that at sea level are shown in Figure 21. At moderate subsonic Machnumbers, the trend is the same as revealed by experiment. As can be seen inTable VII, the predicted percentage differences in the flutter speeds of thevarious configurations are quite close to those measured at these Mach numbers.Comparison of Figure 20 and 21, however, shows that the transonic trends pre-dicted do not agree with experiment: The theoretical recovery is less and moregradual than experienced in the tests and the presence of the tank and the SRBdelay rather than speed the recovery. These failings of the analysis are notunexpected, since the theory assumes that Mach number is constant throughoutthe flow field.

The adjusted flutter boundaries for the configurations with flexibleSRB attachments are presented in Figure 22. These flexibilities cause 2% andh% decreases in the measured subsonic flutter speed relative to the rigidcase, whereas the analytical flutter speed is unaffected. As seen in Table I,the nominal SRB modes are close in frequency to the lowest wing mode, whichanalysis shows couples with the second wing mode to produce the flutter in-stability; therefore, an effect by the SRB modes on the flutter speed is notunreasonable. For the tuned flexure, the frequencies of the SRB modes areclose to the flutter frequency and the even greater effect on the flutterspeed should be'expected. The cause of the theory's failure to predict thiseffect is unknown. However, studies of the Shuttle by Ericsson and Reding inRef. 15 have shown that flow separation occurs at the shoulder of the SRBnose. It is conceivable that, when the SRB oscillates, at frequencies closeto the wing flutter frequency, shed vortices impinge upon the wing causingsignificant deviations from the unsteady pressures expected from potentialtheory. It is also possible that, since the axial doublet strengths are notdetermined simultaneously with the panel doublet strengths, the present methodis not adequate for calculating interaction effects when the body is allowedto oscillate.

15'

PRESSURE STUDY RESULTS

Base Case

For the isolated model wing end-plated at its root, Figure 23 showsgood agreement between the calculated and measured steady-state pressuredistribution, ACp, when the wing is pitched 3°. Here, and in more detail inFigure 2h, even the spillover lift from the outer.panel onto the inner panelof the double-delta wing platform can be seen to be predicted.

Correlation on Wing-Body Configurations

For the thr.ee configurations for which steady-state pressures weremeasured, Figures 2U and 25 give a comparison of .the calculated and measured.distributions for M = 0.6. Comparison of the wing and the orbiter/ET trendsfrom both test and analysis shows that replacing the root-chord reflectionplane by the orbiter and tank lowers the pressure at the inboard leading .edgeof the wing; this effect diminishes outboard and aft.. Agreement betweentest and analysis is good, with analysis overestimating the inboard pressuredrop slightly. ' Similarly,- the presence of the SRB lowers pressure further .at the inboard leading edge and- the amount of this decrease is again predictedfairly reliably. At the inboard trailing edge,, both theory and test1 showthat the SRB causes a pressure increase. Outboard the SRB causes the pressureto drop, but analysis underestimates this decrease. Overall the correlationbetween test and analysis is good; every qualitative feature of the measureddistributions is .matched.

Figures 26 and 27 show, a similar comparison of pressure at M = 0.8. .Generally, the analytical trends are the same-as for M = 0.6. The measureddata, however, does not agree well with these trends. Outboard, the SRBdecreases the pressure more than predicted. Inboard, although the leadingedge pressures agree with theory, pressures on the. remainder of the chordare higher for the- more complex configurations than for the isolated wing -the reverse of the predicted trend. This lack of agreement is consistentwith the degradation of the flutter correlation on the complex configurationsat high subsonic Mach numbers and can probably be attributed to transonic flowphenomena not represented by the theory. There is also the possibility thatflow field is being complicated by flow separation and shed vortices.

Configurational Trends

Comparison of pressure distributions for the various configurationswith the flutter results seemingly leads to contradictions: Figure 28 shows

16

that the spanwise loadings for the isolated wing and the orbiter are almostequal as are the centers of pressure. It might be concluded from this thatthe flutter speeds for these two configurations also should be the same;nevertheless, both test and analysis have shown the orbiter flutter speedto be higher. Also, since in the presence of the external tank and SRB's theloadings decrease and centers of pressure move aft, the flutter speed shouldincrease; however, experimental and analytical flutter results show a flutterspeed decrease.

Explanation for this confusion is that these steady-state pressureswere measured and calculated for configurations where the entire vehicle wasat an incidence angle to the flow. When steady-state pressures were calcu-lated with only the wing at incidence to the flow and the various bodies notinclined to the flow, trends that do support the flutter results are obtained.The resultant spanwise loadings and centers of pressure from such calculationsare shown in Figure 29. The following conclusions are drawn:

• Comparison of the wing alone and orbiter configurations shows thatreplacing the end-plate at the wing root by the rigid, unpitchedfuselage extending to the vehicle 'centerline causes the loadingto naturally decrease and the center of pressure to move slightlyaft. These are stabilizing trends and a higher flutter speedshould be expected on the orbiter.

• The addition of the external tank causes the loading to increasealmost up to the original level of the end-plated wing. The flut-ter speed should be lower than that of the orbiter and slightlyhigher than that of the wing.

• The addition of the SRB (which one can envision as acting somewhatlike a ground plane) raises the loading to a higher level thanthat of the end-plated wing. This effect should lower the flutterspeed again and quite possibly, if the slight aft change in centerof pressure does not entirely negate it, drop the flutter speedbelow that of the wing.

These predictions are entirely consistent with the findings of the flutteranalyses.

FULL SCALE FLUTTER RESULTS

Using the 20 symmetric normal modes listed in Table VI and the aero-dynamics grid 3-^ shown in Figure 11, flutter speeds we're calculated for theisolated wing, the orbiter, and the Shuttle at a Mach number of 0.6 and analtitude of 15,200 m. The full scale vehicle geometry is shown in Figure 1.The resultant V-g-f plots are shown in Figures 30, 31 and 32.

IT

While the flutter speeds of the isolated wing and Shuttle are practi-cally identical, the flutter speed of the orbiter is 3% higher. This is thesame trend as was found on the models. On the full Shuttle, two marginalinstabilities occur at speeds less than 685 knots but as little ab 1% struc-tural damping .stabilizes these roots at all speeds less than 685 knots.

Additional analyses were performed eliminating selected normal modes.As shown in Figure 33, a flutter speed that is only 2% different from thatobtained using 20 modes can be calculated using only five: wing 1st bending,inboard elevon rotation, fuselage pitch (on interstage fittings), fuselagevertical translation (on interstage fittings), and wing 1st torsion.

CONCLUSIONS

• For general combinations of large rigid bodies in proximity to lifting sur-faces, the doublet-lattice/interaciton-panel method successfully predictsthe effects of aerodynamic interaction on the flutter speed of the surfaceat moderate subsonic Mach numbers (M < 0.75).

• More investigation is needed to determine if the effect of flexible or flex-ibly supported bodies on the oscillatory pressures and, hence, on theflutter speed of a proximate lifting surface can be successfully predictedby the present analysis or a modification of it.

• Steady-state pressures on a lifting surface due to the interaction between• it and nearby bodies can be adequately calculated by the present method atM <.T5.

• The subsonic flutter speed (M <-75) of the wing in the presence of all thebodies on the Shuttle is practically the same as that of the wind end-platedat its root. The subsonic flutter speeds of intermediate configurations,such as the orbiter, are higher than that of the wing.

REFERENCES

1. Stahle, C.V.: Transonic Effects on T-tail Flutter. The Martin CompanyReport RM-2H, 1959.

2. Topp, L.J.;'Rowe, W.S.; and Shattuck, A.W.: Aeroelastic Considerationsin the Design of Variable Sweep Airplanes. ICAS Paper No. 66-12,Sept. 1966.

18

3. Watkins, C.E.; Runyan, H.L.; and Cunningham, H.J.: A SystematicKernel Function Procedure for Determining Aerodynamic Forces on Oscil-lating or Steady Finite Wings at Subsonic Speeds. NASA TR R-U8, 1959-

k. Albano, E. ; and Rodden, W.P.: A Doublet Lattice Method for CalculatingLift Distributions on Oscillating Surfaces in Subsonic Flow. AIAAJournal, Vol. 7, No. 2, Feb. 1969, pp. 279-285.

5. Stark, V. 3. E. : Aerodynamic Forces on a Combination of a Wing and aFin Oscillating in Subsonic Flow. SAAB TN 5U, Feb. 196U.

6. Sensberg, 0.; and Laschka, B.: Flutter Induced by Aerodynamic Inter-ference between Wing and Tail. Journal of Aircraft, Vol. 7, No. k,July-Aug. 1970, pp. 319-32U.

7. Mykytow, W.J.; Noll, T.E.; Huttsell, L.J. ; and Shirk, M.H.: Investi-gations Concerning the Coupled Wing-Fuselage-Tail Flutter Phenomenon.Journal of Aircraft, Vol. 9, No. 1, Jan. 1972, pp. kQ-^h .

8. Triplett, W.E. ; Burkhart, T. H. ; and Birchfield, E. B.: A Comparisonof Methods for the Analysis of Wing-Tail Interaction Flutter. AIAAPaper 70-80, Jan. 1970.

9. Chipman. R. R.; Rauch, F. J. ; and Hess, R. W. : Flutter of Pairs ofAero dynamic ally Interfering Delta Wings. Journal of Aircraft, Vol. 5,No. 6, Dec. 1973, pp. 728- 731*.

10. Woodward, F. A. : Analysis and Design .of Wing-Body Configurations atSubsonic and Supersonic Speeds, Journal of Aircraft, Vol. 5, No. 6,Dec. 1968, pp. 528-53H.

11. Hess, J. L. and Smith, A. M. 0.: Calculation of Non-Lifting PotentialFlow About Arbitrary Three Dimensional Bodies. Douglas Aircraft Com-pany Report No. E.S. U0622, 1962.

12. Rodden, W.P. ; Giesing, J. P.; and Kalman, T.P.: New Developmentsand Applications of the Subsonic Doublet-Lattice Method for NonplanarConfigurations, AGARD Symposium on Unsteady Aerodynamics for Aero-elastic Analyses of Interfering Surfaces, Paper No. U, presented inTonsberg, Oslofjorden, Norway, November 1970, AGARD- CP-80- 71-

13. Chipman, R.R. ; and Shyprykevich, P.: Analysis of Wing-Body Inter-action Flutter for a Preliminary Space Shuttle Design. NASA CR-2U29,

ih. Chesler, J. ; Mitchell, L. ; and Clark, ¥. : Preliminary Shuttle FlutterAnalysis and Vibration Analyses, Vol. II. Grumman Aerospace Corp. ReportLD RS-5, Oct. 1973.

15. Ericsson, L.E.; and Reding, J.P. : Unsteady Aerodynamic Analysis ofSpace Shuttle Vehicles. Lockheed Missiles and Space Company ReportLMSC-D352320, Aug. 1973.

19

TABLE I.- 1/80TH-SCALE SHUTTLE MODEL

FREQUENCIES

MODE FREQUENCY, Hz

WING 1st BENDING

WING 1st TORSION

WING 2nd BENDING

WING 2nd TORSION

WING 3rd BENDING

SRB PITCHNOMINAL FLEXURETUNED FLEXURE

SRB HEAVENOMINAL FLEXURETUNED FLEXURE

169.1

449.5

548.0

912.1

1045.9

124.0222.0

242.9438.0

20

TABLE II.- CALCULATED MASS & NODE POINT LOCATION

VIBNODE

31322933302134

221235

23

136

36

24

14

75

37

25

15

8

1

38

26

16

9

2

39

271940

282017

10

3

18

11

4

MASSPOINT

12'3'4'

5*6*7*8*

910*11*12*1314*

1516*17*18*

1920*2122*23 •24*

25* "

26*27 .28*29 '30*3132*

3334*35*

3637*3839*4041*

4243*44*

4546*47*48*4950*51*52*53

54*

55

56*57*58

59*

60

61*

X

METERS

0.25050.25050.28610.28610.33020.33020.3302

0.3543

0.35430.35430.3543

0.37810.37810.37810.37810.37810.37810.39660.39660.39660.39660.3966 ,0.39660.3966

0.39660.41500.41500.4150

0.41500.41500.41500.41500.41500.41500.4308

0.43080.43080.4308

0.4308

0.43080.43080.43080,43080.44160.44160.4416,0.44160.45910.45910.45910.4591

0.4416

0.4416

0.44160.4416

0.44160.45430.4543

0.4543

0.4543

0.4543

INCHES

9.86

9.86

11.2611.2613.0013.0013.00

13.95

13.9513.9513.9514.8914.8914.8914.8914.8914.8915.6115.61

•15.6115.61

. 15.6115.61 •

1 15.61 .,

'15.6116.3416.34'16.34 '•

; 16.34'16.34

: 16.3416.3416.3416.34

16.96

1696

16.96 .

16.96

16.96

16.9616.96

16.9616.9617.3917.39

17.3917.3918.0818.0818.0818.08

17.3917.3917.3917.3917.3917.8917.89

17.8917.8917.89

Y

METERS

0.03970.04600.03970.0524

0.03970.05240.0651

0.03970.0524

0.06510.07780.03970.0524

0.06510.07780.09050.1032

0.03970.0524

.-0:0651

0.07780.0905 .0.10320.1159

0.12860.03970.0524

0.06510.07780.09050.10320.11590.12860.1413

0.0397' -0.0524 •

0.0651

0.0778

0.09050.1032

0.11590.12860.1314

0.03970.0524

0.06510.07780.03970.0524

0.06510.0778

0.09050.10320.11590.12860.14130.09050.1032

0.11590.12860.1413

INCHES

1.56

1.81

1.56

2.06

1.56

2.06

2.56

1.56

2.06

2.56

3.06

1.56

2.06

2.56

3.06

3.56

4.06

1.56

2.06

2.56 :3.06

3.56

4.06

4.56

5.06

•1.56 •2.06

2.56

3.06

3.56

4.06

. 4.56

5.06

'5.561.56 ;

2:06

2.56

.3.06 '

3.56

4.06

4.56

5.06

5.56

1.56

2.06

2.56

3.06

1.56

2.06

2.56

3.06

3.56

4.06

4.56

5.06

5.56

3.56

4.06

4.56

5.06

5.56

MASS

KILOGRAMSx103

1.0480.5912.6731.4773.1722.2241.2011.243

0.6800.7380.664

2.4572.183

1.9481.2450.6200.6911.952

2.019- 1.261

1^092

0.709.0.4160.403

6.356 : ;1.931 -

1 .867

1 .026

0.9750.8550.5560.6500.3810.347

1.716

1.7180.569

6.8200.747

0.542

0.6960.333

0.5710.3110.3840.274

0.4650.2460.6510.2470.394

0.1750.1790.2000.1470.2270.2930.231

0.279

0.3670.415

SLUGSx103

0.07180.04050.18320.10120.21740.15240.0823

0.0852

0.04660.05060.04550.1684

0.14960.13350.08530.04250.0473

0.13370.13840.08640.07480.04860.0285

0.02760.0244

0.13230.12790.07030.06680.05860.03810.04450.02610.0238

0.1176

0.11770.0390

0.05610.0512

0.03720.04770.0228

0.03920.02130.02630.01870.0319

0.01690.03840.01690.02700.01200.0122

0.01370.0101

0.01560.02010.0159

0.01910.02510.0285

•VIBRATION SURVEY MEASUREMENT POINTS

21

TABLE III.- MODEL CONFIGURATIONS STUDIED

CONFIGURATION

1.

2.

3.

4.

5.

6.

ISOLATED WING

ORBITER

ORBITER &TANK

SHUTTLE (RIGID CONNECTION)

SHUTTLE (NOMINAL FLEXURE)

SHUTTLE (TUNED FLEXURE)

•INVESTIGATIONSPERFORMED

P AND F

F

PANDF

P AND F

F

F

•P-PRESSURE DISTRIBUTION (STEADY STATE)

F-FLUTTER BOUNDARY DETERMINATION

22

TABLE IV. - FLUTTER TEST DATA

(a) Wing

R U N

1313141414I S15is16I f ,161ft17171717IS1H1829793031323?33333434353535363ft

R U M

37373S383839401.040414141474?

4?

43

4344

45

45

45464647474848494951505151

O O I N T

001002001002003001002

2.1001002003004001002003004001002

3001002001001001001002

3001

2001

23 -

ooi002

P O I N T

0010020010020030010011.1

002.1

001002on i00?003001002001001002003001002001002001002001002001002001002

••103 FL "OREL V A C H9 F . H A V I O R

1M» FI 0.312I V

I VVVI X

I XIV\ 'fI V

1"

I'-'I V1"

1"iv]M

1W

IV

2"2V7V2-"2V2"?M

' 7 V2M?.M

?'•'2M2W

20'

FSS 0.796S 0.830

FI 0.932FSS 0.946

LD 0.666FI 0.663FSS 0.679

FI 1.034FI 1.026

FSS 1.016FD 0 .98HLD 0.870FI 0.863

FSS 0.873FSS 0.865FSS 0.700FSS 0.682

F I 0 . 7 0 US 1.357S 1.346S 1.244S 1.113

PI 0 .67BFSS 0.93*FSS 0.950

F I 0.944FSS 0.832PSS 0 .32 aF I 0 .736

FSS 0.738FSS 0.735

FI 0.6852v FSS 0.705

V O D F L M O D E L V A C HB E H A V I O R

2V* FI 0 .857?v* F[, 0.349?w» LD 1 1.0692V* LD 1.0492V* FI 1.0302v* S 0.5932v* LD 0.7672M* FSS 0.7627V* FSS 0.7462V* FI 0.9562M» FSS 0 .9272V» FSS 0.9007W* FI 1.0467V* FI 1.0397V* FI 1.016?v« FI 1.1072V* S 1.1047.M* S 0 .6242f* S 0.7542W* , FI 0.7482y* FSS 0.7372f-'» FI O.S137M* FSS 0.8212f* FI 0.9337«» FSS 0.9242V* FI 0.992>M» FSS 0.9792V* S 1.1932M* S 1.1712V» S 1.2562'V» S 1.262?«» S 1.3702V» S 1.353

S STABLE

D Y N A ' - ' I C( M / V 2 )

S6961.95735.5'<685.

111154.123572.37521.97052.

104987.146817.17Q503.186556.175802.

7D132.3S174.

1C0619.104749.109040.106S99.97442 .

179919.202349 .208201.171242.118289.11 6920.139466.109039.108011.

96774 .111159.116968.105996.115162.123183.

(b)

D Y N A M I C( N / V 2 I

119324.129388.134920.145533.165913.94554.96903.

114096.126296.98381.

123602.126151.168611.1804B3.180713.183123.190101.96665.9 2 6 5 H .

106064.123351.88946.

106925.95810.

120118.127486.161864.177C16.207943 .187575.211671.174738.217863.

PRFS.< r > S l )

12. Mi1 3 . R 7

7.9216.1117.9!12.7014.0715.2221.2"24 .7327.0625. 4«10.1612.7814. 5P

1 5 - l f iis .s t i15.4914.1226. n«29.3330.182 4 . P, 217.14

16.9520 .2115.8015.6514.0216.1116. .9 515.3616.6917.85

T R U E V E( " / S E C !

258.2253.1265.3289.57.92.5218.1216.6221.2317.7314.4308.7300.9230.4276.9275.0272 .9220.6214.72 2 3 . 6394.038B.3366.3331.2221.6299.6301.3302.7269.7271.5240.2241.7239.8221 .6227 .4

L O C I T YI F T / S E C )

247 .2830.5870.5949.8959.6715.7710.9725.7

1042.31031.51012,9

9 B 7 . 2920.1908.7902.5895.5723.9704.4733 .3

1292.81274.11201.91086.7

727 .2983.1988.7993.4884.9890.8788.3793.2786.8727 .3746.2

S I GMA

2 .1272.4381.2672.1642.3563.0043 . 3 7 33.5012.3732.6163.1*53.1681.4541.S752.1692.2943.6543 .7B33.1781.8902.1392.5312.5463.9292. 1242.5051.9402.4222.1423.1413.2653.0063.8243.865

MASS D E N .R A T I O

15.2213. 2b25.5614.9613.7410. 7b *

9.609.25

13.6411.50lU.li10.222 2 . 2 617 .2V14. 9<:14.12

b.868.56

10. 1917.13m.7912.7912.7;

0.2415.2412. 9<:16.6913.37Ib . l i10.31y .92

10.778.478.33

F R E O .( H Z )

i'U.2b3.

2 7 U .271.315.311.315.306.330.350.315.<:4b.260.260.263.312.310.304.

260.Hi'243.

^75.264.305.300.300.315.314.

F L O T T E USPEEO I i N D E X

0.3V20.4110.3100.4430.467O..J930.4140.4.JO

0.5090.5490.5740.5570.3520.3940.4^10.4300 .4J90.4^40 .4 i50.5640.5930.6J60.5500.4570.4540.4960.4390.4360.4130.4430.4540.4320.4510.466

Wing and Fuselage

P R E S .( P S I 1

17.29IS,. 7519.5621.0924.0513.7014.Q416.5418.3C14.2617.9118.2824.4426.1626.1926.5427 .5514.0113.4315.3717. fl?12.8915.5013.8917.4118.4823.4625.6630. 1427.1930.6825.3331.58

LD LOW DAMPING

T R U E V EP - ' / S E C )

271.1268.2328.73 2 2 . 7317.01 9 0 . 7246 .62*1.5230 .3302.2293 .22S2 .4318.8313.7301.4331.9328.0193.5244.0241.1236.6260.9262.3296. 3290.7309.7303.5356.1348.4373.9365.0399.6390.0

FSSFD

L O C I T Y( F T / S E C )

689.63S0.1

107B.41058.71040.1

625 .6609.2792.5755.7991.7962. a926.5

1046.01029.3

9 B K . 81089.01076.3

634.9B00.6791.3776.5856.1860.7972 .4953.9

1016.0995.9

1166.41143.11226.81197.51311.31279.5

STEADY

S I G M A

2.6432.9332.0372 .2602.6934 . 2 4 22.5993.1903.8841.7572.3452.5812.7062.9923.2462.7122.8824.2122.5382.9753.5932.1312.5351.7792.3182.1692.8662 . 2 7 72.7952. 1892.5921.7B52 .337

STATE

MASS U i i N .R A T I O

12.2311.0415.8914.2012.02

7.6312.4610. 15

B. 34is. 4313.8112.55i l . 9 b10.82

9.9711.9411.23

7.6912.76

•10.809.01

15.<:o12.77Id. 2013.9714.9311.30I t .Z ' t11.5914.7912.4VIB. 1413.85

FLUTTER

F R E Q .( H Z )

26t>.272.279 .290.316.

300.2 9 U .303.250.^aa .270.300.310.310.313.

295.3lu.2 b u .270.270.270.2 b O .310.

F L U T T E RS P E E D I N D E X

0.4590.4780.4o80.5070.5410.4080.4130.4490.4720.4170.4070.4720.3450.5b40.5050.5690.5790.4130.4040.4330.466O.J960.4340.4110 .< toOO . H 7 40.5340.5590.6060.5750.6110.5550.620

DIVERGENT FLUTTERFl INTERMITTENT FLUTTER

23

TABLE IV.-CONTINUED

(c) Wing, Fuselage and Tank

5253535454

54555555565656565757585H5H5H58595959595959606060616263646464646465656565

U N

V91919202020202121222222?324242525 '25272727282828282 H2S28?.8

=01, -IT

00100100200100?

2.10011.1

002001

1.1002

• 2.1 -001

1.1.1

0010022.12 .2.1

0011.11.21.31.4

0011.11.2

001001001

.1001002

2.12.2

001002

2.12.2

P O I N T

0011.1

0020011.1

002?.l

0010020010020030010010020010020030;)1002or 30010020030044.1

0055.1

006

MODEL MODELB E H A V I O R

2.''* - S2M» FSS2'<» FSS3M» |_D

?-v» FSS2,M« FSS '2M» FSS2M», FSS?.".« ' FSS2M« -FSS2"« FSS2M* FSS?>'.* FSS2M» FSS2«» FSS2M* ' FI2M* S2M* ' FSS2M« FSS2M» FSS

• 2'-l* FSS2M» FSS?M» FSS2-M* FSS2M» - FSS2M» S2M» • s2.'-* FSS2M» FI2M» ' ' S2M» S2M* FSS -2M» -FSS2M» FSS2M* FSS2M* S2M» FSS2M» FI2M* ' S2M* FSS2M* FSS

M O D E L M O D E LB E H A V I O R

- 1«I MI'MI M1M

IM1 "-1

I MIM

- IM• IM

I M1"I M1-MIM

IM

1"

I MIMI M

IM1:/

IMIKIMI MI MI M

F IFSS

SFIFIF IF!LD

FSSLO

' FSS•FSS

SSSSsS

LDFI '

FSSSs

LDFI

FSSFSSFSS

• LD

v-ACH

0.587O . B 4 5O . S 3 20.7150.7050 . 6 8 e0 .7750.7840.7720.8990.9060.9C150.8940.9440.9320.9900.9730.9130.8560.9211.1001.1051 .0871.0771.0731.0561.1311.1071.1211 .2381.3460.8850.7440.7230.7020.6320.6801.0020.9780.9380.874

(d)

I 'ACH

0.6710.6650.6730 .8470.83S0 .8420.8030.7720.7540.9470.9090.9100 .5751.3301.3251 .0931.1121.1010.9770.9580.9100 . 7 6 V0.4670.8960.9090.7930.7950.7360.679

D Y \ ' A " r c|.\V"2)

93414.

113215.117244.10052B.115459.109010.1095B2 .11448B.11S550.11382?.121001 .13071H.141739.155693.157042.1S0115.1H1185 .157004.132701.151126.186142.188494.188381.183482.188506.185107.204015,189588.198355.217723.215941.119489.108780.109193.118945.99129.

116380.163780.168271.152000.125955.

or?ES .("S! 1

13.5416-. 4116.9914.5."16.7415.CC15. P.P.16.5917.1815.5017.54IS. 0520.542 2 . 5 722 .7626.1126. ?622.761 9 . ? 323.3526. 9S27.3227.3127 .3227 .3226. = 329.5727.4828.7531.5631.3017.3215.7715.8317.2414.3716.8723.7424.3922.03IBi 26

Wing, Fuselage,

D Y N A M I C( N / M 2 )

106548.106718.108835.111838.126919.136933.124400.99556.

109565.105388.134074.142454.92755.

170303.208968.93413.

171939.197209.148236.174558.164237.

32570.17190.30524.

126922.134249.138309.108626.37254.

P R E S .( P S I I

15.4415.4715.7716.21l f i . 4 019. P518.0314.431 5 . S ?,15.2719.43

2 0 . 5 513.4424. 6P.30.29-13.5424 .9223.5921.4925.3023.8111.97

2 .4911.6713.4019.4620 .0515.7412.64

T R U E V E L O C I T Y( v / S E C ) ( F T / S E C )

1H6.7 612.8272.0 3 9 2 . 3266.7 875.1232.0 761.4224 .4 736.5219.1 716.3249 .1 K 1 7 . 5250 .7 B 2 2 . 5247.0 810.4284. 1 932 .2285.0 938 .52 B 6 . 0 938.42S1.5 923.6292 .8 960.9289.6 950.1304.0 997.4296.9 973 .32 B O ; 2 919.3263.4 664.42 S 2 . 5 926 .9334.4 1097.2335.1 1099.7330.7 1084.9326i9 1072.6325 .2 1067.2316.1 1037.1-334.6 1097.9329 .9 1082.5333.5 1094-.7364.5 1195.9383.7 1258.9279.3 916.4241.1 - 791.2234.6 769.92 2 7 . 2 74S-.6206 .0 675.8219.9 721.6305.0 1000.8299.1 981.5286.7 .940.7263/0 ' 879.3

S I GMA

4. 3692.4972.6oiJ3.0463.7363.7052 . 6 7 92 . 9 7 23. 1702.3002.4132 .6072 .917

'2 .9613.0553.1803.3563 .2623.1193 .2942.7152 .7372 i B 1 02 . 8 7 72 .9073 .0232 . 9 7 22 .6422 .907

'2 - .67S2i 3932.4983.0513.2353.7583.8123.9252.8713.0683.0162.861

MASS D E N .R A T I O

7.4112.9712.0410.62

6 . 606.7<-

11.23

10. 9J10.2114.061 3.4212. *<L

11.1010.9310.6010.16

V . 6 3V.9 i

10.369.83

11.9211.5311.5211.2511.1410.7110. B911.39•11.1412.1113.5312.9610.6110.01

8.628.498.25

11.2610.5510.7311.32

FRE-J.(HZ )

273.265.300.310.312.295.295.300.270 .2 7 3 .285.295.315.30 J.262.

<:S6.186.286.320.312.31^.310.316.

320.320.320.

275.296.295.310.

312.310.

300.260.

F L U T T E RSPEEO I K O t X

0.4050.4470.4550.4210.431

0.4.;90.4400.449

0.4570.448U . 4 o 20 . 4 o O0 . 3 U O0.5240.3260.5640.5550.3260.4640.5330.3730 . 5 7 70 . 5 7 70 . 3 7 70 . 5 7 70 .5720.6UO0.3780.5920.6200.6170.4590.43'B0.4390 .45B0.4160.4530.5380.5450.5180.471

Tank, and Rigid SRM

T R U E V E L O C I T Y( . " . /SECI ( F T / S E C )

213.3 699.8211.1 692.8214.3 703.1261 .5 857;92 5 8 . 7 -849.0259.5 851.6249 .1 817.32 4 7 . 0 810.5239.0 7 8 4 . 2 '297 .2 '975.1283.9 931.4284.0 93T.S184. 9 - 5 0 6 . 7386.4 1268.03R1.8 1252.8337.6 1107.7336.1 1102.9330.3 1083.6303.1 994.5296.1 971.5281 .4 923.2247 .1 810.7149.9 - 4 9 1 . 9284. 6 933.7?93.6 963.5352.1 8 2 7 . 2?57 .2 843.8231.8 760.6-214.0 7 0 2 . 1 -

S I GMA

3.8213.905

• 3.8662.6683.0923.3163 i 2 7 02.6613.1291.946

• 2 . 7 1 42.8814 .4251.8602.3381.3372.4822.9492.6323.2483.3642.2061.2471.6222.4013.4463.4113.2983.10B

MASS D £ N .RAT 10

0 . 4 76 .296.37

12.1310.47

V . 7 69.90

12.1710.3516.6411.9311.24

7.3217.41 '13.8524.2213.0510.9812.30

9.979.57

14.6825.90IV. 9713.469.409.499.82

10.42

F K E U .( H Z )

320.326.

290.314.

' 310.310.294.31U.2 6 2 .317.3 0 M .

306.332.335.

260.310.312.315.310.i!4.

F L U T T t KSPEED I i M D t X

U . 4 3 40.4340.4380 .444

0.4730.492

.0.4680.4190.44J0.4J10.466O . b O l0.4040.5480.6U70.4060.5510.3900.3110.5550.5360.3020.1740.3770.473O.4b70.4940.4J8

' 0 .3V2

TABLE IV. - CONCLUDED

(e) Wing, Fuselage, Tank and SRM (Flexure I)

DC) I NT •MODEL MODEL •••ACHBEHAVIOR

AAA7A9A9A970707071717?7?727373737474

7576777879797979808080

O01001

.1001002

11.1

002.1

001.1.2

0010011.1

002001002001001001001

.1

.2001002

.1001002

2M* FSS2,M* FSS2-'-'» S2M* FSS?<•'•5M5M5MAMA.MAMA''.A'-'&••'•AMAMAMAMAMAMAMAMA'MAMAMAMAMAMAM

FSSFI

FSSFSSFSSFSSr I

FSSFSSFI

FSSFSSS

FSSSSS

FSSF I

FSSFSSFSSFSSFSSFSS

0.6350.8740.8450.7910.7290.7760 . ?. 1 S0.7790.8950.3910.960C.9520.9460.9760.9510.9161.0730.9191 .1211.2171.3210.917O.S590.3410.8470.8240.6600.6620.657

CYMAVICI.N/M2)

1T2440.125769.104252.111031.112225.135526.128469.124381.155919.164993.1 4 9 3 5 fi .161750.166140.18300S.172217.158893.201990.136234.209194.214990.213B39.126972.95719.101472.11018H.120471.93679.95302.99380.

PRES.(PS! 1

14. PSlfl.2315.1115.1016. 7A15.291°.621 S . 0 324.1923.9221. A523.4524.0826.5324.9623.0329.2819.753P.3231.1631.0018.4013.8714.7115.9717. 4A13.5813.8114.40

TRUE VELOCITY(M/SEC)

208.9278.3273.8255.9235.0?49.0260.2249.1232.7231.2300.0295.6294.3300.5293.9234.4326.12B3.5339.2358.7383.8288.2276.3269.6271.5263.3214.0214.6212.0

(FT/SECI

535.5913.339.8.6339.577].2316.9H53.9817.4927.5922.5984.2973.1965.5986.1964.3933.21070.0930.4

1113.11177.11259.4945.5906.5884.6390.9863.9702.2704.2695.6

SIG.MA

3.82V2.6482.2672.7683.3132.7773.0943.2693.4003.4042.7063.0JO3. 1303.3053.2533.2043.0V82.7642.9652.7252. 3682.4942.0452.2772.4382.8353.3363.3753.607

MASS Osrc.RATIO

d.4612.2314.2011.709.7711.00

10.469.909.50V.51

11.961J.7V10.349. BOV.95

10.1 110.4511.7110.92ll.ua13.6812.9815. B314.22

13.21111.429.709.596.VS

FRt^.ti-U)

JlV.20B.

300.310.300.293.305.330.330.320.330.340.340.310.310.

295.

307.270.280.27s.295.308.318.310.

FLUTTERSHEEL) i.-iL)

0.4250.4710.4290.4430.4450.4jl0.4760.4680.5430.3400.5130.3340.5410.5680.5510.5jO0.5V70.4VO0.6080.6160.6140.4730.4110.4230.4410.4610.4060.4100.41V

(f} Wing, Fuselage, Tank and SRM (Flexure II)

U'-l 00 I NT MODEL MODEL MACH

BEHAVIOR

81

81818282828383838484858ft87878888

89

8990.9f)90

91929?9394

nnil .1

002.1

001OH2

.1

.2001

.1. 001001001

.1oni. i

001. 1

001. 1

001002001001

2001001

A"AMA'-'AMAMAiV

A'-'A'-'6M6'-'A"A''1

AM6MA'-1AMA'-'AMAMAM6"A'-'AM6V

A'-:AMAM

FIFSSFSSFI

FSSFSSFI

F.SSFSSFI

FSSFSSFSSr j

FSSFI

FSSr i

FSSFI

FSSFSSFSSS

FSSS

FSS

0.675C.A640.6510.7870.7860.7t<20 . 7 2 a0.7360.7310.95V0.936C.90o0.8530.35bO.S460.8250.5060.7030.7021 .0060.9860.9200.9511 .:isv0.99V1 .2120.923

TYNA^'IC(N/M2 1

91354.9050.".97453.H9479.9S638.1026S3.B7532.94451.100547.146029.146373.106747.110811.100297.1J7357.93975.93569.37367.96376.165962.183312.149524.139327.217. H 10.

174H7-.

215703.119143.

po£5 .

(PS I 1

13.2413.1314.1212.9714 ,3C'1 4 . « ••'

,12.7013.6914.5721.1721.2215.4715. .'614.5415.5513. A214.29l?.ftA13.972 4 . '. 626.5721.5720.1930.R525.3531.2717.27

TRUE V(•••/SEC)

219.6215.9214.4252.1251.0? 4 .-; . 7237.6236.5233.2302.62V4.12 a a . 9269. B273.7268.7265.2257.9225.7226. ;'i '.; 9 • l3 0 3 . 12B1.7294.7332.2•JOV.H359.52B9.2

ELOCITY S I Gi'.A(FT/SEC) •

7207oe703B?7823816779776765992964947bi)5B9rt8S187034674-J741

1:)14994924967

10SO100V1179949

.6

.6

.6

.1

.7• 0.7.1.1.3.9.9.4.2.3.3.1.5.6. l.6.4

.0

.1

.3

.rt

.0

3.0903.16V3.4572.2972.5532.70-2.5322.7543.0172.6022.7612.0662.4a2

2 . 1 b 32.4252.1792 . 4 1 b2.7VO3. 0772 .OJ43.2543.0732.6173.1453.0122.7212.323

MASS OEMRATIO

10.43

l'J.229.3614. 1012.6o11. VO12.7V11. Vo10.7312.44

11.7315.5213.0414. Bi13.3514.86lj. 3V11.37L'j. 321 1 . 4 JV . V510.54

1 2 . 3 71 J . 3 u10.7511. VO1 J. 94

FKEu.(HZ 1

305.310.310.2 b J .2V'J.2Vi .300 .3UO .31J.320.320.^<;5.20J.

270.280.2bU.ioO.JO J.3uJ.330.335.3J5.32U.

330.

300.

FLUTTcKSPEEu IJMUL.

0.401U . 4 J 0

o.<Ub0.iV70.4170.4 60 . j V 30.4060.4210.5GB0.5JBJ .4 j>4

U.442

0.4*10.4JS0.4J70.417G. 393J.4120.5-41.0.56V0.5140.4V60.013

0.5560.0170.458

25

TABLE V.- AERODYNAMIC COEFFICIENTS FOR ANORBITER FUSELAGE

EXPERIMENT

ANALYSIS

CLa(DEG~1)

0.0041

0.00384

CMa(DEG-1)

0.0015

0.00173

26

TABLE VI.- SHUTTLE SYMMETRIC FREE-FREE NORMAL MODES FOR NOMINAL ELEVONS

MODENO.

'1

2

3

4

5

6

7

8

9

10

11

12

13

14

15

16

17

18

19

20

21

22

23

24

25

26

27

28

29

30

31

32

33

34

35

ORDER INFLUTTERANALYSIS

1

2

3

4

5

6

7

8

9

10

11

12

13

14

15

16

17

18

19

20

FREQ,Hz

2.25

2.77

3.07

3.91

4.04

4.33

5.01

5.17

5.51

6.17

6.96

7.36

8.13

8.88

10.36

10.63

10.89

11.23

11.98

12.59

12.97

14.11

14.30

15.14

15.30

16.83

17.06

17.93

17.99

18.62

19.69

20.01

20.69

22.40

23.81

GENWEIGHT,LB

82,088

376,078

231,358

319,346

100,841

60,322

12,607

944

34,416

340,368

273,739

35,857

41,131

100,011

624

1,318

11,567

1,238

1,425

10,464

13,611

2,932

25,771

363

4,014

36,957

14,801

6,382

13,470

9,136

20,870

377

1,163

14,638

6,547

GEN ,MASS,Kg

37,235

170,589

104,944

144,855

45,741

27,3625,719

428

15,611

154,391

124,168

16,265

18,657

45,365283

598

5,247562

646

4,746

6,174

1,330

11,690

165

1,821

16,764

6,714

2,895

6,110

4,1449,467

171

528

6,640

2,970

PREDOMINANT MODAL DESCRIPTION

FUSELAGE 1st VERTICAL BENDINGSRB YAW

SRB PITCH

SHUTTLE LONGITUDNIAL (FUSELAGE VS OXYGEN)

SRB 1st LATERAL BENDING

ET 1st VERTICAL BENDING

PAYLOAD. PITCH

WING 1st BENDING

ET LONGITUDINAL (OXYGEN VS HYDROGEN)

ET/SRB LONGTIUDINAL (SRB VS HYDROGEN)

SRB 1st VERTICAL BENDING

PAYLOAD 1st VERTICAL BENDING

FUSELAGE 2nd VERTICAL BENDING

ET 2nd VERTICAL BENDING

INBOARD ELEVON ROTATION

FUSELAGE PITCH (ON INTERSTAGE FITTINGS)

SRB 2nd LATERAL BENDING

FUSELAGE VERTICAL (ON INTERSTAGE FITTINGS)

WING 1st TORSION; OUTBOARD ELEVON ROTATION

AFT CREW COMPARTMENT TRUNNION

PAYLOAD 1st LONGITUDINAL

FIN

ET 1st LOCAL OXYGEN TANK VERTICAL BENDING

WING 2nd BENDING; OUTBOARD ELEVON TORSION

FORWARD CREW COMPARTMENT TRUNNION

SRB 2nd VERTICAL BENDING

FUSELAGE 1st LONGITUDINAL

FUSELAGE 3rd VERTICAL BENDING

ET 2nd LOCAL OXYGEN TANK VERTICAL BENDING

SRB 3rd VERTICAL BENDING

SRB 3rd LATERAL BENDING

INBOARD ELEVON TORSIONWING 2nd TORSION; OUTBOARD ELEVON ROTATION

ET 3rd VERTICAL BENDING

FUSELAGE 2nd LONGITUDINAL

27

TABLE VII.- EFFECT OF AERODYNAMIC INTER-ACTION ON FLUTTER SPEED OF A1/80th-SCALE SHUTTLE MODEL

CONFIGURATION

ORBITER •

ORBITER/ET

SHUTTLE(RIGID COUPLING)

SHUTTLE(NOMINAL FLEXURES)

SHUTTLE(TUNED FLEXURES)

M

0.7

0.9

0.7

0.9

0.7

0.9

0.7

0.9

0.7

0.9

VF/VF OF WING ALONEANALYSIS

1.05

1.02

1.02

0.97

1.00

0.95

1.00

0.95

1.00

0.95

TEST

1.07

1.01

1.03

1.02

1.01

1.06

0.99

1.03

0.97

0.97

28

MODEL

0.594m

1

FULL SCALE

23.8m

Figure 1. — Space Shuttle Design Studied

29

WING APEX AT X = 0.174m FROM ORBITER NOSE

0.303m

Figure 2.— Orbiter Wing Model

30

32* 0 42 • >6*x'

• 3 3 • 4 3-34

N* VIBRATION SURVEY POINTSPOINT 11 MEASURED AT Y = 0.0883mPOINT 17 MEASURED AT Y = 0.1120m

Figure 5.- Mass Point and Vibration Measurement Locations

32

(a) FREQUENCY = 169.1 Hz

FIRST BENDING MODE

DAMPING COEFFICIENT = .008

VIBNODE*

NORMALIZEDDEFLECTION

VIBNODE

NORMALIZEDDEFLECTION

VIBNODE

NORMALIZEDDEFLECTION

1234567891011121314

0.72

0.82

0.87

1.0C

0.50

0.26

0.36

0.44

0.49

0.57

0.62

0.08

0.14

0.18

15

161718

19

20

21

22

23

24

25

26

27

28

0.22

0.26

0.290.33

0.18

0.23

0.01

0.03

0.05

6.06

0.08

0..10

0.11

0.12

293031

3233

34

35

36

37

38

39

40

0.01

0.01

0.01

0.01

0.02

(a) 1st Mode*See First Column of Table

Figure 6.- Mode Shape Data, 1/80th-Scale Wing

33

(b) FREQUENCY =449.5 Hz

FIRST TORSION MODE

DAMPING COEFFICIENT = .012

VIBNODE*

NORMALIZEDDEFLECTION

VIBNODE

NORMALIZEDDEFLECTION

VIB: NODE

NORMALIZEDDEFLECTION

1

2

3

4

5

67

89

1011

121314

-0.565

0.022

0.435

1.000

-0.913

-0.892

-0.717-0.348

0.109

0.435

0.825

-0.500

-0.500

-0.390

15

16

17

18

19

20

21

22

23

24

25

26

27

28

-0.152

0.109

0.3260.630

0.260

0.565

-0.109

-0.195-0.195-0.130-0.0440.109

0.175

0.410

29

30

31

32

33

34

35

36

37

38

39

40

-0.022

-0.044

0

0

-0.022

-0.022

-0.022

-0.022

-0.022

0.0660.109

0.240

(b) 2nd Mode

Figure 6.- Continued

*See First Column of Table II

(c) FREQUENCY = 548.0 Hz

SECOND BENDING MODE

DAMPING COEFFICIENT = .019

VIBNODE*

NORMALIZEDDEFLECTION

VIBNODE

NORMALIZEDDEFLECTION

VIBNODE

NORMALIZEDDEFLECTION

1

23

4567891011121314

0.590.790.901.00.07-0.28-0.17-0.14-0.14-0.17-0.24-0.21-0.24-0.31

1516171819202122232425262728

-0.38-0.48-0.62-0.79-0.59-0.83-0.03-0.07-0.10-0.17-0.24-0.34-0.48-0.69

293031323334353637383940

0

-0.03

-0.07

-0.10

-0.21

-0.41

(c) 3rd Mode

Figure 6.— Continued

*See First Column of Table II

35

(d) FREQUENCY = 912.1 Hz

SECOND TORSION MODE

DAMPING COEFFICIENT = .02

VIBNODE*

NORMALIZEDDEFLECTION

VIBNODE

NORMALIZEDDEFLECTION

VIBNODE

NORMALIZEDDEFLECTION

1

2

3

4

5

6

7

8

91011121314

-1.00-0.54-0.120.72-0.590.30-0.16-0.230.030.280.760.500.360.18

1516171819202122232425262728

0.090.090.110.14-0.086

-0.260.160.250.230.140.03-0.108

-0.22-0.55

293031323334353637383940

0

0.07

0

0

0.02

0.04

0.06

0.02

-0.05

-0.17

-0.31

-0.74

(d) 4th Mode

Figure 6.— Continued

•See First Column of Table II

36

(e) FREQUENCY = 1045.9 Hz

THIRD BENDING MODE

DAMPING COEFFICIENT = .021

VIBNODE*

NORMALIZEDDEFLECTION

VIBNODE

NORMALIZEDDEFLECTION

VIBNODE

NORMALIZEDDEFLECTION

1

2

3

4

5

6

7

8

g101112

13

14

-0.260.130.48

1.00

-0.20

0.32

-0.09-0.29-0.32

-0.31

-0.27

0.48

0.31

0.08

15

16

17

18

19

20

21

22

23

24

25

26

27

28

-0.12

-0.19-0.23-0.27-0.060.040.150.23

0.18

0.090.03

0.04

0.11

0.35

29

30

31

32

33

34

35

36

37

38

39

40

0.03

0.07

0.03

0.03

0.03

0.04

0.04

0.03

0.04

0.13

0.23

0.65

(e) 5th Mode

Figure 6.— Concluded

*See First Column of Table II

37

f = 124.0 Hz f = 242.9 Hz

NODE: 1

NODE NO.

MEASURED DEFLECTjONS FOR MODE 1

MEASURED DEFLECTIONS FOR MODE 2

1

.78

.41

2

.55

.43

3

.32

.46

4

.09

.53

5

-.18

.75

6

-.40.

1.00

(a) Nominal Flexures

f = 222.0 HZ f = 438.0 Hz

NODE NO.

MEASURED DEFLECTIONS FOR MODE 1

MEASURED DEFLECTIONS FOR MODE 2

1

1.5

1.0

2

.89

.50'

3

.39

.40,

4

.07

.40

5

-.34

.60

6

-.80

.90

(b) Tuned Flexures

Figure 7.- Mode Shape Data, 1/80th-Scale Rigid SRB on FlexibleAttachments

38

3.0 x 10°

q N/m--

2.0

I

1.0

0.00

• FLUTTER

O INTERMITTENT FLUTTER

O STABLE

— — SUGGESTED FLUTTER BOUNDARY

TUNNEL

T&J JI

PATHI

J

t 4'/r '• i

:\ ?i i,i

? 9i >i \

\\

.6 0.8 1.0 1.2 1.

M

(a) Isolated Wing

Figure 8.— Variation of Measured Dynamic Pressure with Mach Number

39

3.0 x 10a

2.0

,2

1.0

0.00

• FLUTTER

O INTERMITTENT FLUTTER

O STABLE

— — SUGGESTED FLUTTER BOUNDARY

TUNNEL PATH

^y \ \

*^M ^T

9\

9^ ' '\ t\\\

r ?iii

.6 0.8 1.0 1.2 1.

M

(b) Orbiter

Figure 8.- Continued

3.0 x 10°

2.0

1.0

0.0,0

.

• FLUTTER

9 INTERMITTENT FLUTTER

O STABLE

— — SUGGESTED FLUTTER BOUNDARY

TUNNEL

- >Cf W 1 '

1

PATH

*&*

4^f \% :\ * i

., «^rK

l '

.6 0.8 1.0 1.2 1.'

M

(c) Orbiter/External Tank

Figure 8.- Continued

3.0 x 103

i N/m2

20

1.0

0.00

• FLUTTER

O INTERRMITTENT FLUTTER

O STABLE

SUGGESTED FLUTTER BOUNDARY

TUNNEL P

v

J

ATH

K/ I

r "^" J^ |/ \~s'9 S* v '

/

Q'• yr**^ \

^^^^ t^ \

^^

~ v\\11

1111

,\1

?

\

I

6 0.8 1.0 1.2 1.

M

(d) Shuttle, Rigid SRB Attachments

Figure 8.— Continued

1*2

3.0 x 105

I N/m2

2.0

>

1.0

0.00

• FLUTTER

O INTERMITTENT FLUTTER

O STABLE

^— ^— SUGGESTED FLUTTER BOUNDARY

TUNNEL

.TV

PATH

1

.if/t

H^ •"•To

g-*O '

/

9 9i i

6 0.8 1.0 1.2 1.

M

(e) Shuttle, Nominal SRB Attachments

Figure 8.— Continued

3.0 x 10"

|N/m2

2.0

>

1 nI .U

0 0

• FLUTTER

O INTERMITTENT FLUTTER

O STABLE

— SUGGESTED FLUTTER BOUNDARY

TUNNEL

_ + ttH

PATH

/

/'T^T"!

' i( ' .\

/i

cj)ii

0.6 0.8 1.0 1.2 1.

M

(f) Shuttle, Tuned SRB Attachments

Figure 8.— Concluded

.2

ACp 0

-.1

O

D

ORBITERTANKY/s = 0.4

M = 0 . 6

O TOP SURFACE

D BOTTOM SURFACE

ACp = Cp (a) -Cp (0)

a = -3°

ACp 0

-.1 aa

o

FULL-UP SHUTTLE

-.2

ACp 0

-.1D

D

WING ALONE

-.2

0.2 0.4 0.6 0.8 1.0

Figure 9.— Measured Pressures on Wing Surfaces Due to Incidence

ORBITER FUSELAGE

.SLENDER BODYELEMENTS

SOLID ROCKET BOOSTER(SRB)

WING

EXTERNAL TANK(ET)

Figure 10.- Aerodynamic Idealization of Space Shuttle

1*6

GRID 3

\\ \ \\\\ \x\ \ \\\\ \

VAXN\\\ T

\VARIATIONS IN GRID 3

GRIDNO.

NO.OF

BOXES

CHORDWISE

SPANWISE(INNER PANEL)

SPANWISE(OUTERPANEL)

3.0

8

5

7

3.1

7

5

7

3.2

10

5

7

3.3

10

6

7

3.4

8

3

7

3.5

6

3

7

3.6

8

6

7

3.7

8

5

5

3.8

8

5

9

Figure 11.— Aerodynamic Grids Studied on Orbiter Wing

DA • • •

Oa =M

GRID1

GRID 2

GRID3l(3.0)

1 RADIAN

0.7

10

ACn

. 6 . 7 3

X-FROM WING APEX/X OF WING ROOT

1.0

Figure 12.— Effect of Misaligned Aerodynamic Grid on Lift and Pressure Distributions on the Wing.

3DQ3

oa:

z3LLO

XX111

zS

occLL

I

X

3903 9NIQV311VOOT

Oai

O

QC LUH_lZLLO"JUQ

£a.I.

CO

£§>

a.O a.

O

REAL PART IMAGINARY PART

k = 0.07

k = 0.12

k = 0.26

k = 0.61

Figure 14. - Oscillatory Pressures on Isolated Wing at M = 0.6 for Pitch About the Apex

50

(2) —(3) ——(4) ....

4 I-(5) O

REFLECTION PLANE AT WING ROOTGAP BETWEEN ROOT & CENTER LINEUNPITCHED PLATE IN GAPPITCHED PLATE IN GAPFUSELAGE -MODELLED

97

M = 0.7a= 1 RADIAN

\L-.25 0.0 0.750.25 0.50

Y/s

Figure 15. - Effect of Wing Root Condition on Spanwise Lift Distribution.

1.0

51

i >

Figure 16. — Damping and Frequency as a Function of Airspeed for the Isolated 1/80th-Scale WingEndplated at its Root

52

Figure 17.— Damping and Frequency as a Function of Airspeed for the 1/80th-Scale Wing in the Presenceof the Other Shuttle Components (Fuselage, External Tank and Solid Rocket Booster)

•53.

.48

.44

.40

.36

.32

.48

.44

.40

.36

.32

.5

ISOLATED WING

.6 .7 .8M

4.0

= 3.0

o = 2.0

o= 1.0

.45

.5 .6 .7 .8 .9 1.0M

.9 1.0

Figure 18.— Calculated Flutter Speed Indices for the Isolated 1/80th-Scale Wing and theWing in the Presence of the Orbiter Fuselage as Functions of Mach Number andAir Density

20

10

.7

.6

.5

.4

.6

ANALYSIS:

TEST POINTS:

• FLUTTER

9 INTERMITTENT FLUTTER

O STABLE

.5

o

.4

.30.6 0.8 1.0 1.2 1.4

M

(a) Isolated Wing

Figure 19.— Variation of Mass Density Ratio, Flutter Frequency Index and Flutter Speed Index withMach Number for the 1 /80th-Scale Shuttle Model

55

10

.7

.6

.5

.4

o

ANALYSIS:

TEST POINTS:

• FLUTTER

O INTERMITTENT FLUTTER

O STABLE

8o

(b) Orbiter

Figure 19.— Continued

20

10

.7

.6

.5

.4

.6

ANALYSIS:

TEST POINTS:

• FLUTTER

9 INTERMITTENT FLUTTER

O STABLEO

.5

.4

0.6 0.8 1.0

M

(c) Orbiter/External Tank

Figure 19.- Continued

1.2 1.4

57

20

10

.7

.6

.5

- .30.6

ANALYSIS:

TEST POINTS:

• FLUTTER

9 INTERMITTENT FLUTTER

O STABLE

o

(d) Shuttle With Rigid Attachments

figure 19.— Continued

20

10o o

.7

.6

.5

.4

.6

ANALYSIS:

TEST POINTS:

• FLUTTER

Q INTERMITTENT FLUTTER

O STABLE

o O

.5

.4

.30.6 0.8 " 1.0

M

(e) Shuttle With Nominal Attachments

Figure 19.- Continued

1.2 1.4

59

20

10

.6

.5

.4

.6

.5

.4

.3

0.6

ANALYSIS:-

TEST POINTS:

• FLUTTER

0 INTERMITTENT FLUTTER

O STABLE

0.8 1.0

M

(f) Shuttle with Tuned Attachments

Figure 19.— Concluded

1.2 1.4

60

Vf

0.40.6 0.7 . 0.8

MACH NUMBERFigure 20.— Test Flutter Boundaries Adjusted to a = 3.0

0.6

0.4

IWING

ORBITERI

ORBITER/ETI

SHUTTLE

0.6 0.7 0.8

MACH NUMBER

Figure 21.- Analytical Flutter Boundaries at a = 3.0

0.9

1.0

1.0

61

Vfbr ur

0.6

0.5

0.4

0.30.6

ANALYSISI

•— TEST, RIGID COUPLING

TEST, NOMINAL FLEXURE

... TEST, TUNED FLEXURE

0.7 0.8

MACH NUMBER

0.9 1.0

Figure 22.- Shuttle Flutter Speeds at a = 3.0 for Various SRB Attachment Flexibilities

62

d

CACO

ci

•soIJ28

oo'*3

.n'5VI

5£

£Q.

CM0)

63

0.2

0.1

ACp

0.0

0.3

0.2

ACp

0.1

0.0

AO

TEST

WING A

ORBITER/ETQ

SHUTTLE n

a = -3

jf-04

0.2 '0.4 0.6

ANALYSIS

0.8 1.0

Figure 24.— Steady State Pressure Distributions on the Inner Winoat M = 0.6

0.4

0.3

ACp0.2

0.1

0.0

0.3

0.2

ACp

0.1

A1= o,

oTEST ANALYSIS

WING £

ORBITER/ET Q

SHUTTLE

-3

Figure 25.- Steady State Pressure Distributions on the Outer Wingat M = 0.6

65

0.2

ACp

0.1

0.0

0.3

0.2

ACp

0.1

0.0

o

D

\

O

TEST

AWING

ORBITER/ET Q

SHUTTLE n

0.0 0.2 0.4 0.6X,'c

ANALYSIS

0.8 1.0

Figure 26.— Steady State Pressure Distributions on the Inner Wingat M = 0.8

66

0.4

0.3'

ACp 0.2

ACp 0.2

Figure 27.— Steady State Pressure Distributions on the Outer Wing atM = 0.8

67

Axc

.20

.15

.10

.05

.00

0.0 25

WING

ORBITER

—•— ORBITER/EXTERNAL TANK

SHUTTLE

.50

Y/S

.75 1.0

Figure 28.— Calculated Spanwise Loadings and Centers of Pressure for VariousShuttle Configurations at Angle of Incidence

68

.7

.5

Axc

.3

.20

.15

.10

.05

.000.0 .25

WINGORBITER

— •— ORBITER/ET• • • • • SHUTTLE

.50

Y/S

.75 1.0

Figure 29.— Calculated Spanwise Loadings and Centers of Pressure for VariousShuttle Configurations — Bodies at Zero Incidence

69

TO

Figure 30.— Damping and Frequency as a Function of Airspeed for the Full-Scale WingEndplated at its Root.

Figure 31.— Damping and Frequency as a Function of Airspeed for the Full-Scale Wing inPresence of the Orbiter Fuselage

the

71

Wtn. 13IO. fr f ULLI TflLCA1OUT pnuIHJL ::: : :_;.. .. ^ H. _._. I . ! i-gr ' - . CT; . i . T . .1 . ; : . 1 iTTT. . . • I . . !

Figure 32.— Damping and Frequency as a Function of Airspeed for the Full-Scale Wing in thePresence of Other Shuttle Components (Fuselage, External Tank and Solid RocketBooster)

72

Figure 33.— Damping and Frequency as a Function of Airspeed for the Full-Scale Wing in thePresence of Other Shuttle Components (Fuselage, External Tank and Solid RocketBooster). Only Five Modes Used]

NASA-Langley, 1975 CR-2U88 73

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PENALTY FOR PRIVATE USE S3OO SPECIAL FOURTH-CLASS RATEBOOK

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