PREDICTION OF FORCES AND MOMENTS FOR

FLIGHT VEHICLE CONTROL EFFECTORS-- FINAL REPORT

PART I: VALIDATION OF METHODS FOR PREDICTING

HYPERSONIC VEHICLE CONTROLS FORCES AND MOMENTS

i

NASA Grant NAG 1-84g

The National Aeronautics and Space Administration

Langley Research Center

ATTN: Dr. John Shaughnessy

MS 494

Hampton, VA 23665

u

M. Maughmer, L. Ozoroski, T. Ozoroski, D. Straussfogel

Department of Aerospace Engineering

The Pennsylvania State University

233 Hammond Building

University Park, PA 16802

May 1990

n

(NASA-CR-186571} PREDICTION OF F_RCES AND

MOMENTS FOR FLIGdT VEHICLE CONTROL

EFFECTORS- PART 1: VALInATInN OF WETHOOS FOR

PREDICTING HyPeRSONIC VEHICLE CONTROLS

FORCES AND MOMENIS Final Report

G3107

N90-2173_

Uncles

0278027
Abstract

Many types of hypersonic aircraft configurations are currently being studied for their

feasibility of future development. Since the control of the hypersonic conflguratins throughout thespeed range has a major impact on acceptable designs, it must be considered in

the conceptual design stage. Part I of this report examines the ability of the aerodynamic

analysis methods contained in an industry standard conceptual design system, APAS II,

to estimate the forces and moments generated through control surface deflections from low

subsonic to high hypersonic speeds. Predicted control forces and moments generated by

various control effectors are compared with previously published wind tunnel and flight

test data for three configurations: the North American X-15, the Space Shuttle Orbiter,

a_d a hypersonic research airplane concept. Qualitative summaries of the results are given

for each longitudinal force and moment and each control derivative in the various speed

ranges. Results show that all predictions of longitudinal stabiltiy and control derivatives

are acceptable for use at the conceptual design stage. Results for most lateral/directional

control derivatives are acceptable for conceptual design purposes; however, predictions at

supersonic Mach numbers for the change in yawing moment due to aileron deflection and

the change in rolling moment due to rudder deflection are found to be unacceptable. Including shielding effects in the analysis is shown to have little effect on lift and pitching

moment predictions while improving drag predictions. Overall, lateral/directional control

derivatives show better agreement when shielding effects are not included.

In Part II of this report, an investigation of the aerodynamic control effectiveness

jot highly swept delta planforms operating in ground effect is presented. A vortex, lattice

:computer program incorporating a free wake is developed as a tool to calculate aerodynamic

:stability and control derivatives. D_ta generated using this program are compared to '

experimental data and to d_tafrom other vortex-lattice programs. Results show that an

elevon deflection produces-greater increments in CL and CM in ground effect than the

same deflection produces out of ground effect and that the free wake is indeed necessary

forg0od predictions near the ground. .................

u

m

__ PRECEDING PAGE BLANK NOT FILMED
i

=

r

Table of Contents

Page

Overview ................................. 1

Part I:Validation of Methods for Predicting Hypersonic Vehicle

Controls Forces and Moments

Introduction ................................ 3

North American X-15 Research Aircraft .................. 4

Low Spee,d: Moo = .056 ...................... 4

Transonic: Moo = 0.80, 1.08, 1.18 ................. 4

Supersonic: .Moo = 2.96 ...................... 7

Hypersonic: Moo = J.65, 6.8_ ................... 8

Summary/Results with Mach Number ................ 9

Hypersonic Research Airplane ...................... 9

Low Speed: Moo = 0.20 ...................... 9

Transonic: Moo = 0.80, 0.98, 1.20 ................. 10

H!/persbnic: _Moo = 6.00 ...................... 11

Summary Results with Mach Number ................ 12

Rockwell Space Shuttle Orbiter ..................... 12

Low Speed: Moo = 0._0 ............ .......... 12

Transonic: Moo = 0.80 ...................... 14

Hypersonic: Mc¢ = 5.00, 20.0 ................... 14

Summary Results with Mach Number ................ 16

Conclusions and Recommendations ...................... 17

References ................................. 19

Nomenclature ............................... 21

Tables .................................. 22

Figures .................................. 24

Appendix: Bibliography of Experimental Force and Moment Data

For Hypersonic Vehicle Configurations

Ill
. l

u

m

Table of Contents (cont'd)

Page

Part H: An Analysis of Delta Wing Aerodynamic

Control Effectiveness in Ground Effect

Introduction ................................ 1

Motivation for the Investigation ..................... 1

Scope of the Research ......................... 2

Review of Previous Work ........................ 3

Theoretical Predictions ....................... 3

Experimental Results ....................... 4

Theoretical Developments .......................... 6

The Vortex-Lattice Method ...................... 6

The Biot-Savart Law ....................... . . 11

Determination of the Influence Coefficients ............... 13

Calculation of the Lift Coefficient ................... 16

The Free Wake ........................... 18

The Ground Effect ........................ 23

The Suction Analogy ....................... 27

Total Lift ........................... 27

Total Pitching Moment ..................... 30

Flap Deflections ........................... 31

Results and Discussion .......................... 33

Verification of Results .......................... 33

VLM-FIG Predictions ........................ 49

Summary and Conclusions ......................... 64

References ................................ 66

Appendix A: Computer Program Description ................ 69

Programming Method ........................ 69

Output ............................... 79

Appendix B: Computer Program Listing .................. 80

Nomenclature .............................. 109

iv
r

w

m

m

Overview

The purpose of this final progress report is to present the findings of two research

activities which focus on hypersonic vehicle configurations. The first of these activities,

discussed in Part I of this report, is aimed at validating methods, suitable for conceptual design work, which can be used for predicting flight control forces and moments over the Mach

number range from low subsonic through hypersonic. Specifically, the subsonic/supersonic

panel methods of Woodward, called the Unified Distributed Panel method (UDP) 1, are

examined for subsonic and supersonic flight Mach numbers, while an enhanced version

of the Hypersonic Arbitrary Body Program (HABP) of Gentry 2 is considered for hypersonic speeds. AI1 of these methods have been incorporated into a single analysis program

called the Aerodynamic Preliminary Analysis System H (APAS) 3'4, which is used in this

study. Although HABP has been widely used for conceptual design activities since the

early 1970's and a number of studies have been undertaken to examine its ability to predict the overall vehicle aerodynamics, no comprehensive, systematic study has been found

which has explored its ability to predict forces and moments generated by aerodynamic

flight controls. Thus, the goal of the present effort is to determine the accuracy and range

of validity of the simple local surface inclination methods for predicting control forces and

moments for a variety of configurations.

The approach used in this validation effort is to compare predicted and experimental

results for several vehicle configurations which cover a broad range of proposed hypersonic

vehicle configurations and for which wind-tunnel data are available.

The second research activity, considered in Part II of this report, is directed at studying

the flight control behaviour of highly swept delta wing planforms in close proximity to the

ground. Because of the possibility that the stability characteristics and/or the control

effectiveness of such configurations is altered significantly by ground effect, this is an

important area of concern in considering the take-off and landing of proposed hypersonic
flight vehicles. Furthermore, any analysis of such configurations in ground effect must

examine the coupling that exists between deflecting a control surface to achieve a proper

moment for trim and the change that such a deflection causes in the total lift generated.

In order to explore these issues, a vortex-lattice program, which includes a free-wake and a

reflective image plane to model ground proximity, has been developed. With this program

it is possible to examine the change in moment of the entire vehicle, as well as the change in

control effectiveness, due to ground effect. In this way, it can be determined whether or not

sufficient control power for trim in ground effect is available and if not, the vortex-lattice

code should be useful in evaluating innovative ways of generating the required moments

to trim.

r .

w

r_

m
w

r--

--i

w

Part I: Validation of Methods for Predicting Hypersonic Vehicle

Control Forces and Moments

Introduction

Before a great deal of activity is undertaken in assessing different types of flight

control systems for hypersonic vehicles, it is important to understand the strengths and

limitations of the prediction tools most likely to be used in this effort. Towards this end,

a number of supersonic/hypersonic methods have been evaluated. Specifically, because

their inherent simplicity makes them ideally suited to conceptual design work, among

the tools under examination are the classical local surface inclination methods, including Newtonian theory, tangent-wedge/tangent-cone methods, and shock expansion techniques. These methods are all part of an industry-standard computer program called

the "Hypersonic Arbitrary Body Program (HABP)," originally prepared by Gentry 2

and now part of a more encompassing program, the "Aerodynamic Preliminary Analysis System (APAS)," detailed in Refs. 3,4. Although HABP has been widely used for

conceptual design activities since the early 1970's and a number of studies have been undertaken to examine its ability to predict the overall vehicle aerodynamics, Refs. 5-9 for

example, it is apparent that no comprehensive, systematic study has explored its ability

to predict forces and moments generated by aerodynamic flight controls. Thus, the goal

of the present effort is to determine the accuracy and range of validity of the simple local surface inclination methods for predicting control forces and moments for a variety of

configurations.

In addition to exploring the validity of the hypersonic capability of APAS, results

using the subsonic and supersonic panel methods of that program, including a wettedarea drag prediction, were also compared with experimental data. Although more sophisticated methods are available for predicting control forces and moments in subsonic

and supersonic flows, few offer the generality, simplicity and speed of the panel methods.

3
The approach taken in this validation effort is to examine several vehicle configurations which cover a broad range of proposed hypersonic vehicle configurations, and

for which wind-tunnel data are available. For each configuration, the theoretical and

experimental results are compared across the entire speed range of subsonic, transonic,

supersonic, and hypersonic Mach numbers. The configurations analyzed are the X-15,

shown in Fig. I, the Hypersonic Research Airplane, a wing-body vehicle, as shown in

Fig. 2, and the Space Shuttle Orbiter, presented in Fig. 3. Details concerning the geometry specification and the program implimentation for this study are contained in Ref. 10.

The experimental data to which the theoretical results are compared are included in the

Appendix of this report, "Bibliography of Experimental Force and Moment Data for Hypersonic Vehicle Configurations," which is a rather extensive survey of available sources

for such data.

:North American X-15 Research Aircraft

r_

r

The North American X-15 research airplane was developed in the late 1950's and

flown in the early 1960's. It was designed to reach flight velocities of 6,600 ft/sec and

altitudes of 250,000 feet. The wind tunnel data used to compare with theoretical predictions are taken from Refs. 11-14.

Low Speed: Moo = .056

Comparisons of the longitudinal subsonic panel method results of APAS are compared to low-speed experimental data (Moo = .056) in Figs. 4-8. In each figure, theoretical zero-flap deflection results are given by the solid line, while the dashed line is

for a 40 degree flap deflection. From the lift-curve data of Fig. 4, it is seen that over

the linear range up to about 25 degrees angle of attack, the predicted results are within

approximately 15% of the experimental ones. As expected, because separated flow is

unaccounted for in the method, the rounding off of the lift curve is not captured. This

likewise probably contributes to the lift-curve slope not being well predicted. Over the

4
w

r _

linear range of the data, the theoretical method predicts the change in lift due to a flap

deflection reasonable well, although for the most part the change in lift is slightly underestimated.

The general behavior of the drag as it depends on angle of attack is predected fairly

well by the wetted-area build-up method used in APA$, Fig. 5, although the absolute

drag values, particularly at higher angles of attack because separation drag is not taken

into account, are underpredicted. The errors in lift and drag as they depend on angle

of attack compensate one another in the plot of CD against C_, Fig. 6. In this case, up

until the drag increases due to separated flow become appreciable, the method does a

good job when compared to the experimental results.

Pitching moment coefficient versus angle of attack, Fig. 7, and versus lift coefficient,

Fig. 8, demonstrate the well-known difficulty of predicting pitching moments using methods which do not account for separation. In fact, only the predictions of the zero-lift values of the moment coefficient with zero flap deflection are even close to the experimental

values. The change in moment with flap deflection is, in particular, not well predicted.

Above about 20 degrees angle of attack where the results are largely affected by separation, the method, of course, produces a very poor estimation of the experimental results.

Results are presented for a downward elevator deflection of 5 degrees in Figs. 9-

13. In this case, while the absolute values of the coefficients are somewhat in error, the

changes due to the control deflection are all reasonably well predi_:ted. In particular,

the change in pitching moment due to the elevator deflection is predicted quite well, although other features of the pitching moment curves, Figs. 12 and 13, are not even approximated well.

In the comparisons of the predicted and experimental lateral/directional coefficients,

Figs. 14-19, the different line types correspond to different treatments of the leading and

side-edge suction forces. Included are purely potential flow with no edge treatment, potential flow plus the leading edge suction force, and potential flow plus leading and side5
L

r

edge modifications. The change in side force, yawing moment and rolling moment with

differentially deflected elevons are presented as they depend on angle of attack in Figures

14-16, respectively. Although not accounting for flow separation probably contributes

to the discrepency at higher angles of attack between the theoretical and experimental

results for C,_,,, Fig. 15, all of the other results are certainly good enough for conceptual

design work. In addition, for these cases, the inclusion of leading edge and/or side-edge

suction analogies does not make a great deal of difference in the results.

The change in side force, yawing moment, and rolling moment with deflection of

the vertical tail are presented with their dependence on angle of attack in Figures 17-19.

Although not as good as the previous results, these too are probably close enough for

many conceptual design studies. Note that in these figures, only the potential flow plus

leading edge suction force results have been included. Unfortunately, some of the trends

with changes in angle of attack are not captured and, in particular, it should be noted

that the theoretical method misses the loss in control effectiveness at higher angles of attack, as seen in Fig. 18, probably because the method does not account for the separated

flow off the fuselage which blankets the vertical tail.

Transonic: Moo = 0.80, 1.08, 1.18

For transonic Mach numbers of Moo = 0.80, 1.03, and 1.18, the subsonic and supersonic panel method results are compared for experimental data only for the longitudinal

aerodynamic derivatives at elevon deflections of 0 degrees, -3 degrees, and -6 degrees.

The plots of these comparisons are presented in Figs. 20-31. In all cases, the angle of attack range is limited to less than 20 degrees such that the influence of separated flow on

the results is limited.

As observed in Figs. 20, 24, and 28, both the trends and absolute levels of the lift

curves are predicted quite well. In addition, the change in lift coefficient with elevon deflection is predicted well. Figs. 21, 25, and 29 demonstrate that the drag curves are well

predicted, particularly at the lower angles of attack, for Mach numbers of 0.80 and 1.18.
L_

L_

= :

w

=

Near Moo = 1.00, however, the drag coeffcient is over-predicted and some unexplained

discontinuities exist in the drag curve. In all cases, the drag increment due to a control

deflection is not resolved well, although the trends are correct. Similar comments apply

to the drag polar £urves, Figs. 22, 26, and 30.

The pitching moment coefficient is plotted against the lift coefficient for the three

Mach numbers in Figs. 23, 27, and 31. The zero-lift values of the pitching moment coefffcient, as well as the change in moment coefficient with elevon deflection, are possibly predicted close enough for some purposes. The slope of the curve (static stability, aCm/aCL) is not predicted all that well; however, it improves as the Mach number

moves away from Moo = 1.00.

Supersonic: Moo = _.96

Comparisons of experimental and theoretical results for longitudinal coefficients at

a Mach number of 2.96 are presented in Figs. 32-35. Lift and drag comparisons are not

unlike those at lower speeds. The prediction of the slope of moment coefficient against

lift coefficient curve, Fig. 35, is better than at lower speeds. Also, it is seen that the

prediction of the change in moment coefficient with elevon deflection is not too bad for

smaller deflections, where presumably the amount of separated flow on the surface is

small, but deteriorates at larger deflection angles.

The prediction of lateral/directional coefficients with a differential elevon deflection

are presented in Figs. 36-38. In these plots the potential flow result overlays that with

the leading-edge modification included. While the predictions entirely miss some of the

non-linear behavior observed in the experimental results, they are probably good enough

for conceptual design work. Similar comments can be made for the vertical tail deflection results, Figs. 39-41. In particular, the directional control effectiveness, Fig. 40, is

predicted quite well provided the vehicle angle of attack is not too great. Above an angle of attack of 25 degrees, the control effectiveness falls off rapidly, most likely due to

the blanketing of the vertical tail in separated flow off the body.

7
=

Hypersonic: Moo = 4.65, 6.83

For Mach numbers of 4.65 and 6.83, experimental results are compared with theoretical results obtained using the Hypersonic Arbitrary Body Program (HABP) which

is contained in APAS. When using HABP, the user has the choice of whether or not to

ignore the aerodynamic contributions of components "shielded" (or shadowed) in the

wake of upstream parts of the vehicle. For the X-15 runs, it was found that the use of

shielding had very little impact on the results. Consequently, except where noted, the

predictions presented are calculated without the shielding option.

The comparison of theoretical and experimental reults of the longitudinal data for

a Mach number of 4.65 are presented in Figs. 42-45. The lift-curve, Fig. 42, clearly

demonstrates the non-linear behavior typical of hypersonic flow. While the general

shape of this curve, as well as those of Figs. 43-45, is captured adequately, the effect of a

symmetrical elevon deflection is not. Because the effect of a control deflection on the CD

vs. a curve, Fig. 43, is small compared to the effect on CL,the underprediction of the effects of a control deflection observed in Figs. 44 and 45 is likely due to the problems of

predicting the effect on lift.

Theoretical and experimental comparisons of lateral/directional coefficients are

given in Figs. 46-51. Near zero angle of attack, the side force and yaw due to an unsymmetrical elevon deflection, Figs. 46 and 47, respectively, axe predicted reasonably

well; however, at higher angles of attack, only the sign of the force and moment is predicted correctly. Although the rolling moment due to the unsymmetrical elevon deflection is not predicted too badly, the control effectiveness is underpredicted, particularly

at higher angles of attack. The prediction of side force, yawing moment, and rolling moment due to a vertical tail deflection is relatively good and certainly adequate for conceptual design studies.

Comparisons are presented for a Mach number of 6.83 in Figs. 52-64. The comments made for the M_ = 4.65 case apply to these results as well. Note the negligible

8
= =

impact of shielding, Figs. 56-58, on lift, drag, and moment coefficients.

Summary Result8 with Mach Number

Some of the preceding lateral/directional comparisons are replotted versus Mach

number in Figs. 65-70. The change in side force, yawing moment, and rolling moment

due to an unsymmetrical elevon deflection are presented for two angles of attack in

Figs. 65-67. Away from Moo = 1.00, the zero angle-of-attack predictions are quite good.

From Fig. 67 it is seen that at higher Mach numbers the roll control power is considerably less than it is at low speeds. Although not predicted by HABP, it does increase

somewhat as the Mach number increases. As should be expected, the change in angle of

attack has little impact on the change in side force and yawing moment with unsymmetrical elevon deflection, Figs. 65 and 66, and the trends are predicted fairly well. Unfortunately, in the case of change in rolling moment with unsymmetrical elevon deflection, the

effect of an angle of attack change is not predicted.

The changes in the lateral/directional control derivatives for a ;vertical tail deflection

are presented in Figs. 68-70. For the Mach numbers shown and zero angle of attack, the

predicted results agree quite well with those obtained experimentally.

i _

Hypersonic :Research Airplane

Comparison of theoretical predictions with experimental data for the Hypersonic

Research Airplane (HRA), Ref. 15, are now considered. This blended wing/body aircraft has many configuration characteristics typical of proposed hypersonic vehicles. The

wind-tunnel results which will be used for this comparison are taken from Refs. 15-17.

Low Speed: Moo = O.PO

Experimental results are compared with APAS subsonic panel method predictions

of the longitudinal aerodynamic characteristics of the HRA in Figs. 71-75. Overall, the

agreement between the experimental values with those calculated is similar to that obtained for the X-15 aircraft. The lift coefficient versus angle of attack, Fig. 71, is like

_9
the X-15. The lift coefficient is underpredicted at higher angles of attack. Possibly, the

extent of vortex lift is not fuly taken into account. The change in lift coefficient with a

symmetrical elevon deflection, CL,_, is predicted well enough for conceptual design activities. Also as was the case for the X-15 and observed in Fig. 72, the drag coefficient

is increasingly underpredicted as the angle of attack increases. The agreement between

experiment and theory in the CD--CL curve, Fig 73, reflects the characteristics of the CL

and CD versus angle of attack results.

The pitching moment coefficient predictions, Figs. 74 and 75, are much better for

the HRA than for the X-15. Also, as is the case for the X-15 and more important from a

flight controls point of view, the prediction of the elevon control power, Crn_, is reasonable. It is observed, however, that the magnitude of the control power for the X-15, possibly because of the all-moving tail and the longer moment arm, is considerably greater

than that of the HRA.

In considering the prediction of the lateral/directional coefficients, Figs. 76-81, results are presented for calculations using potential flow, potential flow with leading edge

suction analogy, and potential flow with leading and side edge suction analogies. In the

case of the change in side force, yawing moment, and rolling moment with differentially

deflected elevons, Cy_, Cr,_, and Cl_ the predictions using potential flow plus the leading edge modification agree quite well with the experimental results. The changes in the

coefficients with a vertical tail deflection, Cy6,, Cn_., and Cz6° are not well predicted

and are typically in error by 50% or more. In all cases, use of potential flow plus the

leading edge suction treatment gives the best results. The inclusion of the side edge

modification makes little difference or causes the agreement between the experimental

and predicted results to be worse.

Transonic: Moo = 0.80, 0.98, 1._0

For the case of Moo = 0.80, the comparison between the experimental and predicted

results, presented in Figs. 82-89, is very similar to that for Mc¢ = 0.20 and additional

10
commentsare not necessary. As the Mach number increases,the prediction of CL_e deteriorates, Fig. 90, as does that of elevator control power, Cm6o Figs. 93 and 94. The

drag coefficient, Fig. 91, is further underpredicted when compared to the lower Mach

number cases. The predicted values of the change in the side force, yawing moment,

and rolling moment coefficients with differential elevon deflection, Cr6=, Cn6o, and Cl_=,

Figs. 95-97, are certainly adequate for conceptual design activities. The deterioration

of Cr_6_ and C,_, predictions noted for Moo = 0.80 continues as the Mach number increases to Moo = 1.20. Comparison of experimental and theoretical results for this Mach

number are presented in Figs. 98-105.

Hypersonic: Moo = 6.00

Comparisons of experimental results with predictions using HABP without shielding for the longitudinal aerodynamics are presented in Figs. 106-110. While the CL vs.

a data shown in Fig. 106 appears better than that for the X-15, it should be noted that

the values of CL, as well as CL_ are considerably less. In any case, with the exception

of missing that CL_ decreases as the angle of attack decreases, the magnitudes of the

forces and most of the trends are predicted reasonably well. In the case of the drag coefficient, Figs. 107 and 108, the trends are captured fairly well although, opposite of what

happens in the case of the X-15, the drag coefficient is consistently overpredicted. The

moment coefficient predictions for this case, Figs. 109 and 110, are also not too bad and

probably good enough for conceptual design activities. Certainly the trends, as well as

the values of the elevator control power, C,m_, are reasonably well predicted.

The longitudinal cases just presented without shielding are repeated with shielding

in Figs. 111-115. In all cases here, the use of shielding improves the predictions considerably. In particular, incorporating shielding produces much better agreement between

the experimental data and the calculated coefficients at lower angles of attack. As a consequence, the values of the control derivatives, CL6_ and C,_ are much improved.

Comparisons of wind-tunnel findings with HABP predictions with and without

11
shielding for the changes in the lateral coefficients with differentially deflected elevons

are given in Figs. 116-119. For these cases, the predictions without shielding agree

somewhat better with experimental data than those with shielding. In either case, while

the predictions are good, it should be noted that the control power, Cz,_, is very low.

Similar comments apply to the change in coefficients due to a vertical tail deflection,

Cy6,, Cr,_,, and Cl6,, as presented in Figs. 119-121.

Summary Results with Mach Number

Summaries of the lateral control derivatives as they depend on Mach number and

at two angles of attack are given in Figs. 122-124. Overall, Cy6_, C,_6_, and Cz_. are

predicted well at low speeds, reasonably well in the transonic range, and very well at

M_ = 6.00. With the exception of the prediction at M_ = 6.00, this is not the case

for Cr,,, C,_,, and Cz_,, presented for zero angle of attack as a function of Mach number in Figs. 125-127. In fact, the trends as well as the values of these derivatives are not

well predicted. This was not the case for the X-15, which achieved directional control by

means of very large all-moving, wedge cross-sectioned surfaces both above and below the

vehicle centerline.

Rockwell Space Shuttle Orbiter

The third hypersonic vehicle configuration for which theoretical and experimental

results are compared is the Rockwell Space Shuttle Orbiter. All of the experimental data

used for this comparison is taken from Ref. 18.

Low Speed: M_ = 0._0

Longitudinal aerodynamic coefficients as predicted using the subsonic panel method

of APAS are compared with experimental findings in Figs. 128-132. Included in these

comparisons are results for full-span elevon deflections of-20, -10, 0, and 10 degrees. As

observed in Fig. 128, the predicted lift coefficients as a function of angle of attack are,

for the most part, within about 10%. While both the predicted and experimental results

12
are closeto linear for the range of anglesof attack presented,the lift-curve slopesare

not well predicted. This causesfairly large descrepanciesbetweenthe predicted and experimental results at low anglesof attack and large negativeelevondeflections. In spite

of theseproblems, the flap effectivenessof the elevon,CL6o, is predicted reasonably well

except at low angles of attack and large negative deflection angles for which not accounting for separated flow evidently causes the control effectiveness to be overpredicted.

ComParisons of predicted and experimental drag coefficients ae presented versus

angle of attack in Fig. 129, and versus lift coefficient in Fig. 130. In both cases the predicted results are probably good enough for conceptual design activities.

Predicted and experimental pitching moment coefficients are presented as a function

of angle of attack in Fig. 131, and as a function of lift coefficient in Fig. 132. In both

cases the predicted values of the coefficients are not too good. In fact, the slopes of the

predicted and experimental curves are sometimes of the opposite sign. Although slightly

overpredicted, however, the predicted elevon control effectiveness, C,me, is reasonably

close to that found experimentally. Predicted values of the lateral control derivatives for

the Shuttle are compared with experimental values in Fig. 133-135. These panel method

predictions for differentially deflected elevons are made using potential flow only, potential flow plus leading-edge effects, and potential flow plus leading- and side-edge effects. As was the case for the other vehicles examined, the potential flow, modified with

the leading-edge suction analogy only, yields the best agreement with experiment. This

agreement is probably acceptable for conceptual design work.

Predicted and experimental comparisons for directional control derivatives, Cry,,

C,_6,, and Cz6,, are shown in Figs. 136-138. In these cases, only the potential flow plus

leading-edge effects are presented. As before, the predictions are not too bad. At worst,

the control effectiveness, C,_, is predicted to be 30% greater than the value found experimentally.

13
=

Transonic: M_ = 0.80

Predicted lift curves for different elevon deflections are compared with experimental results at M_ = 0.80 in Fig. 139. Although similar to the Mo_ = 0.25 comparison,

the problem areas are more pronounced. For example, the lift-curve slopes are generally

not well predicted, the bending over of the curves due to flow separation is not captured,

and the control effectiveness, C16o, while not too bad for 6e - :kl0 °, is poorly predicted

for 6e = -20 °. In the case of CD versus a and CD versus CL, Figs. 140 and 141, the

comparisons are similar to those for M_ = 0.25. While significant errors are present as

separated flow strongly influences the results at higher angles of attacks, the drag predictions capture the trends correctly and are acceptable for conceptual design activities. In

the case of CM plotted against a, Fig. 142, and against CL, Fig. 143, the 6e = 0 prediction is reasonably close to the experimental data. Other than that, about all that can be

said is that the predicted trends are more or less correct.

As in the Met -- 0.25 case, the best lateral control derivative predictions are obtained using the potential flow results modified with Ieading-edge effects. The predicted

values of Cyea, Cr,_,, and Cz_a as they depend on angle of attack, Figs. 144-146, are reasonably close to the experimental values. The directional derivatives, Cy6, and C,e,,

however, presented in Figs. 147 and 148, are in error by over 100% in some cases. The

predicted values of Cz_,, as observed in Fig. 149, are somewhat better.

Hypersonic: M_ --5.00, 20.0

Values of CL, CD, and Cm predicted using HABP without employing the shadowing of downstream components are compared with experimental results for M¢¢ = 5.00

in Figs. 150-154. The predicted lift coefficient as it depents on angle of attack, Fig. 150,

is in excellent agreement with the experimental data at angles of attack below 20 °, and

is not too bad even up to 40 °. The flap effectiveness for a symmetrical elevon deflection,

Cl6_, is predicted very well. Likewise, comparison of the predicted CD- (_ and CD- CL

curves with experiment, Figs. 151 and 152, are adequate for conceptual design studies.

14
k__

L_

For the comparisons with experiment of the predicted Cm - a, Fig. 153, and Cm - CL,

Fig. 154, certainly the characteristic trends are captured; however, the control effectiveness is significantly overpredicted for large negative elevon deflections.

Predicted and experimental changes in CL, CD, and C,_ due to deflecting the Shuttle body flap are presented in Figs. 155-157. In each case, the overall trends are more

or less predicted by the methods in HABP. Most likely, the descrepancies in predicting

these results are due to the fact that the effects of flow separation are not taken into account.

The influence on the longitudinal aerodynamic coefficients of shadowing panels

which are shielded from the oncoming flow by upstream elements is demonstrated by

the results presented in Figs. 158-162. It is observed that, while shielding does not make

a critical difference in the cases of lift and drag coefficients, Figs. 158-160, the shielded

results are uniformly better. For the pitching moment coefficient, Figs. 161-162, the

shielded results are better in the low angle of attack and lift coefficient range, while the

unshielded results are better at higher angles of attack and lift coefficients. In view of

the fact that for these cases the use of shadowing produces superior results, at least for

lift and drag, Figs. 163-167 repeat the previous predictions for lift, drag, and moment

coefficients but with downstream components shielded.

Comparisons between HABP predictions for the lateral control derivatives and experiment are presented in Figs. 168-170. None of the calculated results employ shielding; however, the predictions are seen to capture both trends and magnitudes quite

well. Unfortunately, this success is not repeated for the directional control derivatives,

Figs. 171-173. It should be noted, however, that the magnitudes of these derivatives are,

in general, quite small.

The comparisons generated for M_ = 5.00 are repeated in Figs. 174-197 for

Moo =20.0. Basic unshielded results are given in Figs. 174-178 while corresponding

shielded results are presented in Figs. 187-191. For the most part, the comments di15
rected to the Moo = 5.00 comparisons also apply to the Moo = 20.0 results.Ifanything,

the HABP predictions generally improve as the Mach number increases. As before, the

predicted values of the pitching moment coefficients,Figs. 177-178 and 190-191, are not

good; however, the values of C,_6. are reasonable. Also, itshould be noted that the directional control derivativepredictions (without shielding)are considerably better at

Moo = 20.0 than at Moo = 5.00. Again, itshould be noted that the control effectivenesses,Cz6, and Cn_,, are quite low.

Summary Results with Mach Number

Plots of the predicted and experimental values of the lateral/directional control

derivations at fixed angles of attack as they depend on Math number are presented in

Figs. 198-203. Each of these plots indicate that the prediction improves as Math number increases although, with the exception of the transonic region, they are reasonably

good elsewhere. In every case, it should be noted that the control derivatives, including

those of control effectiveness, decrease as the Math number increases.

16
Conclusions and Recommendations

Z==.

w

The suitability of the methods contained in APAS for flight control force and moment prediction are summarized in Table 1 for longitudinal quantities and Table 2 for

lateral/directional derivatives. Based on the comparisons between experimental and predicted results made in this study, the assessment of predicting a particular quantity in

a given speed range is ranked as: (1) capable of predicting magnitudes within 40%, as

well as the proper trends, over a significant portion of the operating range; (2) capable

of predicting trends reaonably well but not magnitudes; (3) generally unacceptable for

conceptual design work.

In considering the results presented in Table 1, it is concluded that, provided the

results are interpreted carefully and major decisions are not based on predictions which

are known to be suspect, all of the longitudinal results can be of some use in conceptual

design studies. As expected, the most questionable calculations are those in which neglected separated flow can significantly influence the results. In the case of predicting

lateral/directional control derivatives, some results are found to be unacceptable. Fortunately, however, for the most part the primary control derivatives, Cl6_ and Cn_,, are

found to provide reasonable results.

Given that for conceptual design work, the trade-off between accuracy and speed of

computation leans very heavily toward speed, all of the prediction methods contained in

APAS are reasonable. It should be noted, however, that in the case of the subsonic and

supersonic methods, a great deal of flexibility exists in controlling the trade-off between

accuracy and speed. The panel methods considered are basically as fast and robust as

possible at the expense of accuracy. If more accuracy is desired, a number of improvements are possible between panel methods and a full-blown Navier-Stokes computation.

For example, in the direction of increasing complexity and accuracy, a more complete

boundary layer model could improve the drag predictions and, combined with empiri17
r

= ,

v

cal correlations, could greatly improve the derivatives strongly influenced by separation

effects. If better accounting for the effects of compressibility is required, the flowfield

could be modelled using the Euler equation rather than the linearized potential equation.

While in the case of subsonic and supersonic Mach numbers there is a great deal of

flexibility for achieving a desired level of accuracy by implementing a sufficiently rigorous theoretical model, this is not the case in the hypersonic flight regime. Essentially,

there is nothing in the way of a more sophisticated theoretical model between the flow

inclination methods contained in HABP and a full Navier-Stokes computation. Unfortunately at present, the time required for setting up computational grids, as well as execution requirements, are much too great for the full computational methods to be routinely used for conceptual design activities. Thus, methods such as those contained in

HABP will likely remain the primary tools for preliminary flight control force and moment prediction for some time. For such work these methods can be expected to provide

reasonable estimates provided they are applied to situations not dominated by strong

viscous/inviscid interactions, real gas and/or rarefied gas effects, or fiowfields containing

extensive regions of separated flow.

Finally, while little can be done to remove the noted limitations of the hypersonic

flow inclination methods, the following should be considered to improve their usefulness

and accuracy:

(1) guided by the existing data base, make empirical adjustments in the methods to

account for the anomalies seen between predicted and experimental results

(2) add a more sophisticated boundary-layer theory, including a transition model, to

better account for viscous effects

(3) develop an unsteady Newtonian Theory (based on piston theory) to allow the computation of dynamic stability and control derivatives.

18
References

.

.

o

Q

o

.

o

°

.

10.

11.

12.

13.

Woodward, F., "Analysis and Design of Wing-Body Combinations at Subsonic and

Supersonic Speeds," Journal of Aircraft, Nov.-Dec. 1968.

Gentry, A.E., "Hypersonic Arbitrary-Body Aerodynamic Computer Program

(Mark III Version). Vol. 1-User's Manual," Report DAC 61552, McDonnell-Douglas

Corporation, April 1968.

Bonner, E., Clever, W., and Dunn, K., "Aerodynamic Preliminary Analysis System II-Part I, Theory," NASA CR-165627, 1981.

Sova, G. and Divan, P., "Aerodynamic Preliminary Analysis System H-- Part II,

User's Manual, North American Aircraft Operations, Rockwell International.

Middleton, W.D., Lundry, J.L., and Coleman, R.G., "A System for Aerodynamic

Design and Analysis of Supersonic Aircraft, Part 1-Development," NASA CR 3352,

1980.

Fisher, C.M.E., "Experience Using the Mark IV Supersonic Hypersonic Arbitrary

Body Program," Aerodynamics of Hypersonic Lifting Vehicles, ACARD CP-428,

Paper 31, April 1987.

Covell, P.F., Wood, R.M., Bauer, S.X., and Malker, I.J., "Configuration Trade and

Code Validation Study on a Conical Hypersonic Vehicle," AIAA Paper 88-4505,

September 1988.

Moore, M.E. and Williams, J.E., "Aerodynamic Prediction Rationale for Analyses

of Hypersonic Configurations," AIAA Paper 89-0525, January 1989.

Cruz, C.I., and Wilhite, A.W., "Prediction of High-Speed Aerodynamic Characteristics Using The Aerodynamic Preliminary Analysis System (APAS), AIAA Paper 89-

2173, August 1989.

Ozoroski, L.P., "Evaluation of Methods for Predicting the Forces and Moments

Generated by Control Surface Deflections on Hypersonic Vehicles," M.S. Thesis,

Pennsylvania State University, 1990.

Boisseau,Peter C., "Investigation of the Low-speed Stability and Control

Characteristics of a 1/7-Scale Model of the North American X-15 Airplane,"

NACA RM L57D09, 1957.

Osbourne, R.S., "Aerodynamic Characteristics of a 0.0667 Scale Model of the N.A.

X-15 Research Airplane at Transonic Speeds," NASA TM X-24, 1959.

Penland, J.A. and Fetterman, D.E. Jr., "Static Longitudinal, Directional, and Lateral Stability and Control Data at a Mach Number of 6.83 of the Final Configuration of the X-15 Research Airplane," NASA TM X-236, April 1960.

19
14.

15.

16.

17.

McKinney, R.L.a.ud Lancaster, J.A., "Investigation of the Aerodynamic Characteristics of a 0.02-ScaleModel of the X-15 Airplane at Mach Numbersof 2.96, 3.96,

and 4.65 at High Angles of Attack," NASA TM X-820, June 1963.

Dillon, J.L. and Pittman, J.L., "Aerodynamic Characteristics at Mach Numbers

from 0.33 to 1.20 of a Wing-Body Design Concept for a Hypersonic Research Airplane," NASA TP 1044, 1977.

Dillon, J.L. and Creel, T.R. Jr., "Aerodynamic Characteristics at Mach Number

0.2 of a Wing Body Concept for a Hypersonic Research Airplane," NASA TP 1189,

1978.

Dillon, J.L. and Pittman, J.L., "Aerodynamic Characteristics at Mach 6 of a WingBody Concept for a Hypersonic Research Airplane," NASA TP 1249, August 1978.

18. "Aerodynamic Design Data Book," Vol. I, Shuttle Orbiter Vehicle, October 1978.

m.

2O

w
Nomenclature

CD =

CL =

Cy -"

Cl =

Cm =

Subscripts

drag coefficient

lift coefficient

side force coefficient

rolling moment coefficient

pitching moment coefficient

yawing moment coefficient

12 _"

6<: =

6I=

6h =

6v "-

aileron or differential elevon deflection

body flap deflection

elevator or synchronous elevon deflection

flap deflection

horizontal tail deflection

rudder or vertical tail deflection

21
E

rO

r,..)

E

ro

,..,1

ro

ro

,,.1

,ro

0 _ _ 0 0

_ 0 O 0

_ 0 0

0 0 0 0 0

O 0 0 0 0

0 0 0 0 0

°_Izl aO• °_

_ _o ._v _v _v

S;_ _v _v =v .=v

o _ ,---]"_ l;zl a6

0

_ 0

0 0

0 _

•- _ • _

0

,,,i

._ "_

_ ._-_

_._ ._

0 _ • _"°

0

0

0

0

0

0

0 _j

cO

_y-_
i

r3

ro

ro

,<,.)

#

%,)

0 0

0 0 0

(9 0 O 0

0 0 0 0

0 0

0 0 0 _ 0

O0

ov

•-_ v 7= v o v _ v

_v _v _ov ,..,:v

ao

t.-i 0

.0 0

0

u 0

_._ .__ o

N o =_o2

_ u

o

"_ _ ."=_ o

'-_ o o

0 _ • _

o

2;

>

o

o

o

o_

o

o

o_

°,._

_ o

4
L_w

0

o_

f_O

L_

Vml

tm
w

0

q_0

"00

.-0

°_
wm

i1111

jll,_

:I:] LIIJj IV

I1 "J

I!

ItJA

.ooU

°_
w

3.5OO --

CL

3. 000

2. 500

2.000

1. 500

1.000

0.500

0.000 (

0 10 20 30 40

Angleof a.tt.a_ -.- degrees

Fig. 4 Comparison Between Predicted and Experimentally Measured Lift on

X-15 Vehicle at Low Speed Including the Effect of Flap Deflection.

X-15;CLvs. a;M=.056;6/ =0 ° and 40 °

w

CD

2.0

1.5

1.0

0.5

0.0

-- " O

I

exp. pred. 65 t_) /

0 " -- 0° .,,,," /

_. .... 40° _ '/

/

/

--: -- , ",..J I I I | ! ] | [ 1 l ] J [ [ : ]

0 10 20 30 40

Angleof at;t_k ~ degrees

Fig. 5 Comparison Between Predicted and Experimentally Measured Drag on

X-15 Vehicle at Low Speed Including the Effect of Flap Deflection.

X-15 ;CDVS. a;M= .056;6/ =0 ° and 40 °
2.0

1.5

CD I .0

0.5

0.0

0.0

m

0

/

_ p,ed. 6S "/

i

0 -- 0° _ --/,"

/x - - - 40" _ "

%

V Jff

! ! ! ]

0.5 1.0 1.5 2.0 2.5 3.0

CL

Fig. 6 Comparison Between Predicted and Experimentally Measured Drag as a

Function of Lift on the X-15 Vehicle Including the Effect of Flap Deflection.

X-15 ; CDvs. CL ;M=.056;6,--0 ° and 40 °

Cm

0.3

0.2

0.1

0.0

-0.1

-0.2

1

A

O

A

O

exp. pred. 6!

O 0 °

A 40 °

_ , _ I I I I , , I , ,o' ' ? ,, , , I

0 10 20 30 40

Angle of attack ~ degrees

Fig. 7 Comparison Between Predicted and Experimentally Measured Moment

on X-15 Vehicle at Low Speed Including the Effect of Flap Deflection.

X-15 ; Cmvs. a;M=.056;6f =0 ° and 40 °
m

Cm

0.3

0.2

0.1

0.0

-0. I

exp. pred. 6!

0 0°

40°

_ 0 0

0-', O_

-0.2 l_,,l,,,,l,11,1,,,,l,,,,[,,,,l

0.0 0.5 I .0 I.5 2.0 2.5 3.0

CL

Fig. 8 Comparison Between Predicted and Experimentally Measured Moment

as a Function of Lift on the X-15 Vehicle at Low Speed Including the

Effect of Flap Deflection.

X-15 ; Crn vs. CL ; M = .056 ; 6f -- 0 ° and 40 °

3. 500 --

3. 000 --

exp. pred. 6,,

2.000 -- [] ........ 5 0 .-_ -""

I.500

Cz

I.000

O. 500

O. 000

-0.500 I t I J I , _ x , I i l i , I , , , , ]

0 10 20 30 40

Angle ofattack~ degrees

Fig. g Comparison Between Predicted and Experimentally Measured Lift on X15 Vehicle at Low Speed Including the Effect ofHorizontal Tail Deflection.

X-15 ; CLvs. a;M=.056;6_ =0 ° and -5 °
2.0

1.5

CD I.0

0.5

0.0

- _-p. pred. 6_o _/

- o :. /

-- _ _'"'"," I I , I i i I

: __r_

0 10 20 30 40

A=gle of at_a_ - degTe_

Fig. 10 Comparison Between Predicted and Experimentally Measured Drag on

X-15 Vehicle at Low Speed Including the Effect of Horizontal Tail Deflection.

X-15 ;CDvs. a ;M=.056;6h =0 ° and-5 °

CD I.0

Ck

Fig. 11 Comparison Between Predicted and Experimentally Measured Drag as a

Function of Lift on the X-15 Vehicle at Low Speed Including the Effect

of Horizontal Tail D,flection.

X-15 ;CDVS. CL IM =.056;6h =0 ° and -5 °
=

/

0.2

0.1

C,,, 0.0

-0. I

-0.2

" ° " " - - - . . . ° o o o

. o

D

D

D

exp. pred. 6h

O 0°

I"I ........ 5°

- °

- °

° °

° + °

- . .

° .

0 o

0

0 o

0

- 0 0

0 10 20 30 40

Angle of a_;T.a_.k-.. degrees

Fig. 12 Comparison Between Predicted _nd Experimentally Measured Moment

on X-15 Vehicle at Low Speed Including the Effect of Horizontal Tail

Deflection.

X-15;Cmvs. a;M=.056;6h =0 ° and -5 °

C,,, O. 0

rl

• °

C,,5

Fig. 13 Comparison Between Predicted and Experimentally Measured Drag as a

Function of Lift on the X-15 "vehicle at Low Speed Including the Effect

of Horizontal Tail Deflection.

X-15 ; Cmvs. CL ;M=.O56;6h =0 ° and -5 °
= =

w

Cys.

0.000

-0.002

-0. 004

.... potentia! oniy

pot. r,lus i.e. sue:ion

..... pot. plus I.e. and side edge effects

0

0

__,. ,-, c_--o : ---q-----

- o .W'o . o o. o . . . ..... o . . ° - -

0

I I 1 , I _-, , , 1 , , , , I I I i IJ

10 20 30 40

Angle of av.ack ~ degrees

Fig. 14 Comparison Between Predicted and Experimentally Measured Side Force

Due to Aileron Deflection on the X-15 Vehicle at Low Speed Including

the Effect of Different Modelling Techniques.

X-15 ; Cy6. vs. a ; M = .056

C_ 4

0.002

0.000

-0.002

.... potemtiaJonly

pot. pills l.e, suction

..... po=. plus I.e. _nd side edge effects

m

0

' - ""C_'-" :_+ -' -(5 ....... _-

0

-- 0

0

,,,, I,,,, I,, ,, I, ,,,

10 20 30

Angle of attack ,,- degrees

0

I

40

Fig. 15 Comparison Between Predicted and Experimentally Measured Yawing

Moment Due to Aileron Deflection on the X-15 Vehicle at Low Speed

Including the Effect of Different Modelling Techniques.

X-15 ; Cn_ vs. a ; M = .056
O. 004

pot enti,,] eniT

pot. ?i_ ].e. sucdon

pot. plus ].{. zmd side ed=e eEec_

Cl,. O. 002

0.000

I i z., I i LI I I I t I 11 _ ' f '?

0 10 20 30 40

A.n$le of at_z_ ~ cle_ees

Fig. 16 Comparison Between Predicted and Experimentally Measured Rolling

Moment Due to Aileron Deflection on the X-15 Vehicle at Low Speed

Including the Effect of Different Modelling Techniques.

X-15 ; Cz_, vs. _ ; M = .056

= =

0.000

Cy,. _0. 002

-0.004

- 0 0 0 0 0 0 __

0 10 20 30 40

Amgle of at_acic ~ degrees

Fig. iT Comparison Between Predicted and Experimentally Measured Side Force

Due to Vertical Tail Deflection on the X-15 Vehicle at Low Speed.

X-15 ; Cy_, vs. c_ ; M = .056
=

m

Cl,_&,

O. 002

O. 000

-0.002

0 0

0

0

0

0

0

l I I I, i I t I I

10

0

0

Ilil,l,,,,l

20 30 40

Angle of _tu_ck ~ degrees

n

Fig. 18 Comparison Between Predicted and Experimentally Measured Yawing

Moment Due to Vertical Tail Deflection on the X-15 "vehicle at Low Speed.

X-15 ; C,_s, vs. a ; M = .056

i O. 004

C_,, O. 002

O. 000

L

C_

I I I i

o _ o o o

10 20 30 40

Angle of z,t_ ~ deg'rees

Fig. 19 Comparison Between Predicted and Experimentally Measured Rolling

Moment Due to Vertical Tail Deflection on the X-15 Vehicle at Low Speed.

X-15 ; Cz_, vs. a ; M = .056
-.,__-

CL

1. 500

1.25O

1. 000

O. 750

O. 500

O. 250

O. 000

-0. 250

-0. 500

D

!

-5

op.

0 0°

[] 3"

•_ _6•

,,,I,,,_I,,,,I,,,,I,,,,I

0 5 10 15 20

X_zg|eof attack ~ clegrees

Fig. 20 Comparison Between Predicted and Experimentally Measured Lift on X15 Vehicle at High Subsonic Speed Including the Effect of Horizontal Tail

Deflection.

X-15;CL vs. a;M=0.8;6h =0 °,-3 °,and-6 °

m

CD

0.6

0.4

0.2

0.0

O

exp. ],red. 6a /

0 0° 0

/.

0 5 10 15 2O

Angle of attack ~ degrees

Fig. 21 Comparison Between Predicted and Experimentally Measured Drag on

X-15 Vehicle at High Subsonic Speed Including the Effect of Horizontal

Tail Deflection.

X-15,'CDvs. a;M=0.8;6h =0 °,-3 ° , and -6 °
0.6 D

O

_p. pred. & _m'_//

0.4 0 o"

Co r't 3° & _ t/,.

.2"

//

0.2 - _/_ J

_ ° "_f' f I

/,.

0.0 ' ' t I , I I

-0.,5 0.0 0.5 1.0 1.5

Cz,,

Fig. 22 Comparison Between Predicted and Experimentally Measured Drag as a

Function of Lift on the X-15 Vehicle at High Subsonic Speed Including

the Effect of Horizontal Tail Deflection.

X-15 ;CD vs. CL ;M=0.8;6h=0 °,-3 °,and-6 °

0.4 w

Cflt

0.2

0.0

-0.2 -

Q----_-_-_..__..E... _

0 0 _

O, U rl

0

exp. pred. 6h

0 0°

[] -3"

..... 6 °

_0

0

0

0

-0.4 I t

-0.5 0.0 0.5 ] .0 1.5

CL

Fig. 23 Comparison Between Predicted and Experimentally Measured Moment

as a Function of Lift on the X-15 Vehicle at High Subsonic Speed Including the Effect of Horizontal Tail Deflection.

X-t5 ;C._vs. CL ;M =0.8;6h =0 °,-3 °,and-6 °
CL

I.500

I.250

I.000

O. 750

O. 500

O.250

O.000

-0. 250

-0. 500

_ ...<.Q>" _P. p_ed. ,_

_ _"

,,,,I,,,,I,,,,I,,_,I,,,, I

-5 0 5 10 15 20

Angle of atuLc.k -- degrees

Fig. 24 Comparison Between Predicted and Experimentally Measured Lift on X15 Vehicle at Transonic Speed Including the Effect of Horizontal Tail

Deflection.

X-15 ;CL vs. ,_;1V[= 1.03 ; 6h =0 °,-3 ° , and -6 °

0.6 -- __.

O. 4 exp. pred. 6._

C D 0[] o" 3 ° /_ / . . ,/,/

/% 6° ./"/"I " "

0.2 - S¢-"

z , , , I i , , , 1 , x , , ] I ' ' ' I-' ' ' ' I

0.9

-5 0 5 10 15 20

Angle of a_t_k ~ degrees

Fig. 25 Comparison Between Predicted and Experime_.:ally Measured Drag on

X-15 Vehicle at Transonic Speed Including the Effect of Horizontal Tail

Deflection.

X-15", CD vs. a;M= 1.03;6h=0 °,-3 °,and-6 °
0.6

CD

0.4

0.2

o.,/

exp. pred. _

O _ 0° / ,"

I:3 .... 3° / ." r ."

0.0 i

-0.5 0.0

lllIlll

0.5 1.0 1.5

C_

Fig. 26 Comparison Between Predicted and Experimentally Measured Drag as a

Function of Lift on the X-15 "vehicle at Transonic Speed Including the

Effect of Horizontal Tail Deflection.

X-15,'CDvs. CL ,'M= 1.03;_h = 0 °,-3°,and-6 °

0.4

0.2

i

i

-0.2 --

A

g-- .0-- -0_ -.__

r7

0 0 0

exp. pred. 6h

0 0°

-3*

Z_

D

_---- ----= _ _L._. 0

0 D

0

0

1 I

-0.,_ ' ' ' , I : .,, , I , : , i I , , 1

-0.5 0.0 0.5 1.0 1.5

CL

Fig. 27 Comparison Between Predicted and Experimentally Measured Mo: ent

as a Function of Lift on the X-15 \-ehicle at Transonic Speed Incluaing

the Effect of Horizontal Tail Deflection.

X-15 ; Cm vs. CL _ M = 1.03 ; 5h = 0 °, --3°, and -6 c
CL

.5O0

I.250

I.000

O. 750

O. 500

O. 250

O.000

-0. 250

-0. 500 = !

-5

_/_.'/ exp. pred. _h

_ _ 0 0°

'." D 3"

, , , I I , , , I , ,., , I , , , , I I , ..t , I

0 5 10 15 20

Angleof ,,t',.a_ .-- degrees

Fig. 28 Comparison Between Predicted and Experimentally Measured Lift on X15 Vehicle at High Transonic Speed Including the Effect of Horizontal

Tail Deflection.

X-15 ;CLvs. a;M- 1.18;_h =0 °,-3 ° , and -6 °

CZ

0.6

0.4

0.2

0.0

exp. pred. 6h /0/"

0 0° /

D -3" //

S"

0 5 10 15 20

Angle of a=_ac_:",_degrees

Fig. 29 Comparison Between Predicted and Experimentally Measured Drag on

X-15 Vehicle at High Transonic Speed Including the Effect of Horizontal

Tail Deflection.

X-15 ; CD vs. a ; M = 1.18 ; 6h -- 0 ¢, -3°, and -6 °
0.6

0.4

0.0

-0.5

3

, I I , , , , I , , , , I , , , , t

0.0 0.5 1.0 1.5

CA

Fig. 30 Comparison Between Predicted and Experimentally Measured Drag as a

Function of Lift on the X-15 Vehicle at High Transonic Speed Including

the Effect of Horizontal Tail Deflection.

X-15 ; CDvs. C_ ;M= 1.18;8h =0 °,-3 ° , and -6 °

Cm

0.4

0.2

0.0

-0.2

exp. pred. 6h

0 O°

0

-0.4 I i J i

-0.5 0.0 0.5 I .0 1.5

C/.

Fig. 31 Comparison Between Predicted and Experimentally Measured Moment

as a Function of Lift on the X-15 "vehicle at High Transonic Speed Including the Effect of Horizontal Tail Deflection.

X-15 ;C,,,vs. CL ;M-- 1.18; 64=0 °,-3 c,and-6 °
2. 000

0 0

_:.j,. pred. bl, 0 _ A

I 500 0 c 0 A .._/" d

• , ....... .,o- xo- o

1.000 _ ._"O-b--'1"-----

c: o L%°'>""" X%>;

o.soo ox<._:>.'-

o.ooo

l"a_1 f I l J I l

-0.500 ,., , I , , ,, I i.l , , I , , , , I I i i , I i _ _ , 1

-10 0 10 20 30 40 50

Angle of a_ta_ -- degTees

Fig. 32 Comparison Between Predicted and Experimentally Measured Lift on X15 "vehicle at Supersonic Speed Including the Effect of Horizontal Tail

Deflection.

X-15 ;CL vs. c_;M= 2.96; 6h =0 °,-20 ° ,-35 °,and-45 °

w

2.0

1.5

1.0

0.5

0.0

I

O

- exp. pred. 6h 0

-- 0 o" 'I_

_ ,, ....... 2co 0

- n -3_o @

o .S/?-.

-I0 0 10 20 30 40

A.ngie of a,_ck ~ 6egree_

5O

Fig. 33 Comparison Between Predicted and Experimentally Measured Drag on

X-15 Vehicle at Supersonic Speed Including the Effect of Horizontal Tail

Deflection,

X-15 ;CD vs. a;M= 2.96:6h =0 °,-20 °,-35 °,and-45 °
C_

2.0

1.5

1.0

0.5

O

OD

exp. pred. _. 0

- 0 O' 0 0

....... 20© 0

O D

A _

D 35=

45 r"

D I,

-_ 0 ; "

o.o -, i_, ,_, 1,,,, I, ,,, L

-0.5 0.0 0.5 I.0 1.5 2.0

CL

Fig. 34 Comparison Between Predicted and Experimentally Measured Drag as a

Function of Lift on the X-15 Vehicle at Supersonic Speed Including the

Effect of Horizontal Tail Deflection.

X-15 ; CD vs. CL ; M = 2.96 _ _h = 0 °, --20 ° , --35°, and -45 °

. o

Fig. 35

C_

0.5

0.0

-0.5

i:--..

"_..^ "-... "..

exp. prea. _. 0

o 0. o

....... 20 ©

..... 35"

<> 45"

L

-1.0 , , , , I , _ _ _ 1_, _ , l. I , , I

-0.5 0.0 0.5 1.0 2.0

0

0

o%

!l!]lil

1.5

CComparison Between Predicted and Experimentally Measured ._omen_

as a Function of Lift on the X-15 _'ehicie at Supersonic Speed Including

the Effect of Horizontal Tail Deflection.

X-15 ;Cm vs. CL ;M =2.96; _h =0 °,-20 ° , --35 ° , and -45 °
Cy_°

O. 000

-0.001

.... _,olenti=! c,ni:"

pot. T,iUs l.e. suc_ior,

..... po:. pJus 1.e. Lnc_ si6e e6_e ef[ec_

- o.c>-c--oc-o"6bO o o

-0.002 ,,, i I,,,, I,,,, I,,,,_I,,,, I

- 10 0 10 20 30 40

Angle of attar.k ~ degrees

Fig. 36 Comparison Between Predicted and Experimentally Measured Side Force

Due to AiJeron Deflection on the X-15 "vehicle at Supersonic Speed Including the Effect of Different Modelling Techniques.

X-15 ; Cy_. vs. a ; M = 2.96

Cn_ I

O. 002 !

0.001

0.000 _ i i r I i _ i i

.... poxemia! c,nI.r

pot. _]u_ ',.e. sue:ion

..... po:. p]_ 1.e. _ si6e e6_e eft'cots

0 0 0

0 0 0 0 0 0 0

° - - . . o . 0

° .... " ..... t I • •

-10 0 10 20 30 40

Angle of s:ta_ ~ 6e_rees

Fig. 37 Comparison Between Predicted and Experimentally Measured Yawing

Moment Due to Aileron Deflection on the X-15 Vehicle at Supersonic

Speed Including the Effect of Different Modelling Techniques.

X-15 ; C_o vs. _ ; M = 2.96
=

C16.

O. 002

0.001

0.000

0

I I. l

-10

.... potcr,;L_! onir

pou ?ius !.e. suction

pot pi_ I.e. _.,nd side edi_e efi'ect$

0

0 0 0 0 0 0 0

, I ,,l,],,,Ll,,, I I ,,, , I

0 10 20 30 40

_gle of at_a_ ~ c_e_zee_

Fig. 38 Comparison Between Predicted and Experimentally Measured Rolling

Moment Due to Aileron Deflection on the X-15 Vehicle at Supersonic

Speed Including the Effect of Different Modelling Techniques.

X-15 ; CI_ vs. a ; M = 2.96

w

0.000

-0. OO5

-O.010

.... pot_dz] oniy

-- pot. pi_ I.e. s_c;ior.

i ..... po:. pi,,_ l.e._.n_side ed_-e eEe:_

0

0

c_ c_ o 0 0 0 0 0

I, I [ I [ I L' '

-10 0 10 2O 30 4O

Angle of att_k ~ degrees

Fig. 39 Comparison Between Predicted and Experimentally Measured Side Force

Due to Vertical Tail Deflection on the X-15 Vehicle at Supersonic Speed

Including the Effect of Different Modelling Techniques.

X-15 • Cy_,. vs. a ; M = 2.96
w

C_ D

O.O04

O. 002

.... pmer.'J_:en)y

po:. ?]_ i.e. su.'tion

..... pot. p}m l.e. r_n_sideedgeef_eczz

_-_b _ d" b _ _ o ........

0

0

0

-10 0 10 20 30 40

Angle ofs_tsc_ _ degrees

Fig. 40 Comparison Between Predicted and Experimentally Measured Yawing

Moment Due to "vertical Tail Deflection on the X-15 Vehicle at Supersonic

Speed Including the Effect of Different Modelling Techniques.

X-15 ; Cm, vs. a ; M = 2.96

Fig. 41

C:t_

0.001

O. 000

-0.001

I°J ! ! I

.... po_e,n_i_]vniy

po:. p]'_ l.e. sue'ion

..... pc_:. p]m i.e. an6 side edge eEecu

O

O O O

O

O C

C O

O

-10 0 10 20 30 40

Comparison Between Predicted and Experimentally Measured Rolling

Moment Due to Vertical Tail Deflection on the X-15 Vehicle at Supersonic

Speed Including the Effect of Different Modelling Techniques.

X-15 ; Cz_, vs. a; M = 2.96
C/.

I. 750

i .dO0

I .250

I. 000

O. 750

O. 500

O. 250

O. 000

-0. 250

- 0

0

0

exp. pred. 15_ />_"//_) _r_O 0

0 _ 0_ 0 /_%_'0.7-

A ....... 20° v/_';."_,__

0

[] ..... 35" 0/_'._"- 0

o ...... oj/ z"g

_"_ V _8 . _ _

,I

-10 0 10 20 50 40 50

Angle of att_ ~ degrees

Fig. 42 Comparison Between Predicted and Experimentally Measured Lilt on X15 Vehicle at High Supersonic Speed Including the Effect of Horizontal

Tail Deflection.

X-i,5 : CL VS. O' ; M -- 4.65

2.0

1.5

1.0

0.5

0.0

0

exp. pred. 6h O/ _

_ o o.

i _ ....... 20" " 9/,/

-- 0 ...... 45" _/

I _ I 1 I I I I I

- 10 0 10 20 30 40

Angle of attack ~ degrees

i

i

5O

Fig. 43 Comparison Between Predicted and Experimentally Measured Drag on

X-15 Vehicle at High Supersonic Speed Including the Effec: of Horizontal

Tai] Deflection.

X-15 ; Co vs. a ; M = 4.65
CD

1.5

1.0 --

0.5 -

0.0

-0.5

& Oo

exp. pred. _z /

'/'_C)

0 0_ _ /,,

A ....... 20 ° I_ /

13 ..... 35° 0 D/,)_/

0 ...... 45° 0 u '._ 0

0 o _/0

• °°°°" 76

0.0 0.5 1.0 1.5

C£

0

0

Fig. 44 Comparison Between Predicted and Experimentally Me'_sured Drag as a

Function of Lift on the X-15 Vehicle at High Supersonic Speed Including

the Effect of Horizontal Tail Deflection.

X-15 ; CD vs. CL ; M = 4.65

r_

Fig. 45

C_r.

0.5

0.0

-0.5

i

o

o

exp. pred. _ - 0 ""_ "

0 _ 0 "¢,.,,

A ....... 20° 0

13 -35 °

O

0 45*

-1.0 K

-0.5 C.O 0.5 1.0 1.5

CO

2.0

CL

Comparison Between Predicted and Experimentally .Measured Moment

as a Function of Lift on the X-15 Vehicle at High Supersonic Speed Including the Effect of Horizontal Tail Deflection.

X-15 ; Cm vs. Cz, : M = 4.65
w

w

t',

1%.

O. 000

-0.001

-0.002

I 0 0 0

,,, I,,,, I,,,,I,,,,I,,,,I

- 10 0 10 20 .30 40

A.ug]e of at_c._ ~ ¢_egrees

Fig. 46 Comparison Between Predicted and Experimentally Measured Side Force

Due to Aileron Deflection on the X-15 Vehicle at High Supersonic Speed.

X-15 ; Cy_ vs. a ; M = 4.65

w

Clot, I

0.002

0.001

0.000

0 0 0 0 0

,,_ I_,,, i, Ii, Jl!!i I ,,, , I

0 10 20 30 40

Angie of _.:_ck ~ degree_

Fig. 47 Comparison Between Predicted and Experimentally Ivleasured Yawing

Moment Due to Aileron Deflection on the X-15 Vehicle at High Supersonic

Speed.

X-15 ; C,_ vs. a ; M = 4.65
O.O01

0.000

-0.001

0

0 0 0 E 0 0 0 0 0

0

- 10 0 10 20 30

Angle of ,,t_a,ck ..- degrees

0

I

4O

Fig. 48 Comparison Between Predicted and Experimentally Measured Rolling

Moment Due to Aileron Deflection on the X-15 Vehicle at High Supersonic

Speed.

X-15 ; Cz_ vs. a ; M = 4.65

w

0.000

C_%, !

-0.004 0 _- 0 _'--'-- '-"0 " 0

J 0 0 0

I I I i J

-0.008

-I0 0 10 20 30 40

Amgieof_ack ~ degree_

Fig. 4g Comparison Between Predicted and Experimentally Measured Side Force

Due to Vertical Tail Deflection on the X-15 Vehicle at High Supersonic

Speed.

X-15 ; C1_,, vs. a : M = 4.65
O. OO4

O. 002

O. 000

0

0 0 O. O_

,,, , I,,lllJ,,, I,,,, I,,,, I

- 10 0 10 20 30 40

A.n_Jeof _t_l- -- degrees

Fig. 50 Comparison Between Predicted and Experimentally Measured Yawing

Moment Due to Vertical Tail Deflection on the X-15 Vehicle at High

Supersonic Speed.

X-15 ; C,_, vs. a ; M = 4.65

--=

0.002 -

0.000

-0.002

-t0

F 0 0 0 0 0

0

- C c'> o 0

0

_,,_ I,_,, l_j_, I:,111,,_,, I

0 10 20 30 40

AngJe of a_aci: _ degre_

Fig. 51 Comparison Between Predicted and Experimentally Measured Rolling

Moment Due to Vertical Tail Deflection on the X-15 Vehicle at High

Supersonic Speed.

X-15 ; Ct,,,, vs. a ; M = 4.65
= =

v

I

r--

I.000

E ,x O. 800 gxp pred. 6_. z,

A ........ --:15 _ _.._,.'"

o o • r

O.60C -- 0 o A O'J_/.,.70

I'I _ - lO° . ._._; "

- _. 20° _ -_

C]. 0,400 -- V --- 30"A ._ A y :

.., .- "" 0

0.200 --_,_ .. _a_-5"_I"_> v

o. ooo _ v

-0.200 _,,,, l,,,,l,,_,I,,,,I,,,,l,,,,l

-5 0 5 10 15 20 25

Angle of _tt_k -- degrees

Fig. 52 Comparison Between Predicted and Experimentally Measured Lift on X15 Vehicle at Hypersonic Speed Including the Effect of Horizontal Tail

Deflection.

X-15 ; CL VS. a ; h_ = 6.83

CD

0.4

0.3

0.2

0.1

0.0

B Z_

'D

exp. pred. 6_ "'I

A ........ +15 ° ,'//_/

- o ..... o.

[] ....... lO o A ._...,

¢ ......... 20" ; .,_

- _, _ _ -30 = ,//

_-

-5 0 5 10 15 20 25

Angle of at;ack ~ degre_

Fig. 53 Comparison Between Predicted and Experimentally .Measured Drag on

X-15 \:ehic]e at Hypersonic Speed Including the Effect of Horizontal Tail

Deflection.

X-15 ; CD vs. a ; M = 6.83
= =

CD

Cz,

Fig. 54 Comparison Between Predicted and Experimentally Measured Drag as a

Function of Lift on the X-15 Vehicle at Hypersonic Speed Including the

Effect of Horizontal Tail Deflection.

X-15 ; CD vs. Cr ; M = 6.83

=

C_,_

0.2

0.0

-0.2

A - "0''. _

exp. pred. 6_. /'

A ........ +15 © ,,,

0 0°

..... I0°

20°

x7 _ _ -30 °

0

A

-0.8 i,! _ I 1 ] , I I I I J

-0.2 0.0 0.2 0.4 0.6 0.8 1.0

CL

Fig. 55 Comparison Between Predicted and Experimentally Measured Moment

as a Function of Lift on the X-15 Vehicle at Hypersonic Speed Including

the Effect of Horizontal Tail Deflection.

X-15 ; C,n vs. Cc ; M = 6.83
I .00 --

C1

0.75

O. 50

0.25

0.00

-0.25

0 i

withoutshield;n_ 0 /

,,,,I,, ,, I,,,, I,,,, I,,,,I,,, ,J

-5 0 5 10 15 20 25

.,_ag|e of attack -- degrees

Fig. 56 The Effect of Shielding on the Predicted Lift on the X-15 "vehicle at

Hypersonic Speed.

X-15 ; CL vs. a ; M = 6.83 ; with and without shielding

w

0.4 i

o3 //

withou= shielding /_O

. . .... • . o ..

C9

0.2

--

0.1

0.0 , l , J , I : I , I

-0.2 0.0 C.2 0.4 0.6 0.8

CL

Fig. 57 The Effect of Shielding on the Predicted Drag on the X-15 Vehicle at

Hypersonic Speed.

X-15 ; C D vs. CL ; M "- 6.83 ; with and without shielding
i

w

O. _

C_

0.0

-0.1

-0.2

-0.3

O

without shie|din_

...... with shielding;

0

,. I 1 I , t i I ,.. d

-0.2 0.0 0.2 0.4 0.6 0.8

C.T.

Fig. 58 The Effect of Shielding on the Predicted Moment of the X-15 Vehicle at

Hypersonic Speed.

X-15 ; Cm vs. CL ; M = 6.83 ; with and without shielding

m

m

Cy,..

0.001

O. 000

0.001

{-c_----c_ o

E °

I ' _ I ! [.| I I ! [ I I I

0

,.i,,,,rg,,,o._. .

-5 0 5 10 15 2O

A_ie of _:=aci; ~ de_-ees

Fig. 59 Comparison Between Predicted and Experimentally Measured Side Force

Due to Aileron Deflection on the X-15 "vehicle at Hypersonic Speed.

X-15 ; CY, o vs. a ; M = 6._3
Cet_ a

0.001

C.O00

0.001

- 0 0 0 0

m

i,,, !,,,, I,,,, I,:,,),LL, I

-5 0 5 10 15 20

Angle of _:_r_ ~ degrees

Fig. 60 Comparison Between Predicted and Experimentally Measured Yawing

Moment Due to Aileron Deflection on the X-15 Vehicle at Hypersonic

Speed.

X-15 ; C._ vs. a ; M = 6.83

m

.002 --

.001

•000

0

0

0

0

_,z , I , J ,,0 I _ , _ , I, , , , I , , , , I

-5 0 5 10 15 20

Angle of attack ~ degrees

Fig. 61 Comparison Between Predicted and Experimentally Measured Rolling

Moment Due to Aileron Deflection on the X 15 Vehicle at Hypersonic

"_eed.

., 15 ; Cl_,, vs. a : M = 6.83
C)%,

.0050

.0025

.0000

l

0 0

0 0 0 0

.__0

"t I t t I i i i , I i i I I I i I I i _ ! I i I I

-5 0 5 10 15 20

Amgieof _t_ac.k --degrez_

Fig. 62 Comparison Between Predicted and Experimentally Measured Side Force

Due to Vertical Tail Deflection on the X-I5 Vehicle at Hypersonic Speed.

X-15 ; C}%, vs. ¢z ; M - 6.83

w

O. 000 --

C_L,

-0.002

-0.004

m

0 O 0 o 0

0 0

-5 0 5 10 15 20

Angle of a_=ack ~ degree_

Fig. 63 Comparison Between Predicted and Experimentally Measured Yawing

Moment Due to Vertical Tail Deflection on the X-15 "vehicle at Hypersonic

Speed.

X-15 ; Cn_, vs. a : M = 6.83
Cl_,,

0.001 --

ZO

O.O00 --

-0.001 I , , I

-5

0 0

0

0 0 0

5 10 15 20

Fig. 64 Comparison Between Predicted and Experimentally Measured Rolling

Moment Due to Vertical Tail Deflection on the X-15 Vehicle at Hypersonic

Speed.

X-IS ; Cz,, vs. a ; M = 6.83

C_%.

O. O00

-0.002

-0. 004

..-0""" "9"" o

• °

• .., .-,, " exp. pred• a

',;'- 0 ........ o°

r [] 12 °

!

l , I , ] , I , I

0 2 4 6 8

Mach number

Fig. 65 Comparison Between Predicted and Experimentally Measured Side Force

Due to Aileron Deflection on the X-15 Vehicle as a Function of Mach

Number.

X-15 ; Cy_,, vs. M ; a = 0° and 12 °
Cr,£ m

O. 003

O. 002

0.001

O. 000

0

I I l

exp. pred. a

0 ........ O_

[] 12°

..0

(_ " " •

1

0 2 4 6

M_ numb_

I

B

Fig. 66 Comparison Between Predicted and Experimentally Measured Yawing

Moment Due to Aileron Deflection on the X-15 Vehicle as a Function of

Mach Number.

X-15 ; Cn_. vs. M ; a = 0 ° and 12 °

CI_ a

0.002

0.001

0.000

9cb,,

, %

" . %

• .%

f

I 1 , "

0 2

exp. pred. a

O ........ 0°

0 12 °

D

I o' ..... I.... , I

4 6 8

Math number

Fig. 67 Comparison Between Predicted and Experimentally Measured Rolling

Moment Due to Aileron Deflection on the X-15 Vehicle as a Function of

Mach Number.

X-15 ; Ct_ vs. M ; o_ = 0 ° and _2 °
C_'1, •

O. 000 V,

-0. 002 _p-,

-0.004 _--

-0.006 r

-0

O. ° o - ° " °

_0 o - w

, 1 , I I 1 I

0 2 4 6

Math number

I

8

Fig. 68 Comparison Between Predicted and Experimentally Measured Side Force

Due to Vertical Tail Deflection on the X-15 Vehicle as a Function of Mach

Number•

X-15 ; Cy,,, vs• M ; a = 0 °

=

_7

C_.4,,

•006

.004

• 002

• 000

.Q.

"0

i _ I a I I _l

2 4 8 8

Mac.h number

Fig. 69 Comparison Between Predicted and Experimentally Measured Yawing

Moment Due to Vertical Tail Deflection on the X-15 Vehicle as a Function

of Mach Number•

X-15 ; C,_, vs. M ; c_ = 0 °
F.

O. 004 002 ._.

O. 000 Lo.oo2 ] '

"'- 0 ........ 0

::)" " " ° " "

I I ! , 1 I I

2 4 6 8

M_eh number

Fig. 70 Comparison Between Predicted and Experimentally Measured Rolling

Moment Due to Vertical Tail Deflection on the X-15 Vehicle as a Function

of Mach Number.

X-15 ; Ct_o vs. M ; a = 0 °

r

1. 000

0.800

0.600

CL 0.400

0.200

0.000

-0.200

-0. 402

-- exp. pred. 6.

,,_ ....... + 5 ° ,,.,

-- 0 0° Q

L"3 5 °

z_

_ <> 10 ° 0 . - "_

.8

_

- >t

-5 0 5 10 15 20 25

Angle of a:;ack -_ degrees

Fig. 71 Comparison Between Predicted and Experimentally Measured Lift on HypP -onic Research Airplane at Low Speed Including the Effect of Elevator

E ection.

Hn, A;CLVS. a ;M= 0.2; _, = +5 °,0 °,-5 °,-10 °,and-15 °
= =

Fig. 72

Up

0.3

0.2

0.1

0.0

-- O

exp. pred. 6, ,'_,

i

A ....... +S ° A "' //,

o _ 0[ . a'/_//.'

, o......

- o ....... lo" .."/_///

LA/ __

" v __ -is ° " ./.,','/

_ _'/.///

_

Ii1,1111_1 isl,l,,,,I,,,,l_IIII

-5 0 5 10 15 20 25

Angle of att__k ~ de_rees

Comparison Between Predicted and Experimentally Measured Drag on

Hypersonic Research Airplane at Low Speed Including the Effect of Elevator Deflection.

HRA;CD vs. a;M=0.2;6e = +5 °,0 °,-5 °,-10 °,and-15 °

Fig. 73

CD

0.5 , V

, m

exp. pred. 6, / O

A ....... + _° /

02 O -_ 0° ,/ O"

[] ...... 50 ,.?A" r_

O ....... lo ° , _'/ _>" _

/,//

X7 .... 15 ° /,-//_'&,

0 1 ,_ O_

0.0 , I , I

-0.4 -0.2 0.0 0.2 0.4 0.6 0.8

CZ.

Comparison Between ?redicted and Experimentally Measured Drag as

a Function of Lift on the Hypersonic Research Airplane at Low Speed

Including the Effect of Elevator Deflection.

HRA_ CD vs. CL ;M=0.2;6e = ÷5 °,0 °,-5 °,-10 °,and-15 °
C_

0.10

0.05

0.00

-0.05

-0.10

-5

exp. pred. 6,

"% ....... + 5*

0 O*

[]...... 5°

_ O -_o °

:L- __-- &

°.....i o-7--q_--+."

"0-.. 0

0

- _ C)''''-O

0

, ,, , I,,,, I,,,, I,, __, I ,,, , I,,,_'I

0 5 10 15 20 25

Angle of at_az.k - degrees

Fig. 74 Comparison Between Predicted and Experimentally Mcasured Moment

on Hypersonic Research Airplane at Low Speed Including the Effect of

Elevator Deflection.

HRA ;Cmvs. a;M=0.2 ; 6, = +5 °,0 °,-5 °,-10 °,and-15 °

Fig.

0.10

0.05

O. O0

-0.05

exp. pred. 6,

I _ ....... + 50

0 0"

[] ...... ,5*

- _ O m"

"_. _" _. V 15°

'b- <;..._- - v

- - " "+'a-v._qT'---_ _? " -.

°_+ _:.

_'" _

h'. 0 v

"'0

o

A

, ,I , I , I , I , I , I

-0.10

-0.4 -0.2 0.0 0.2 0.4 0.6 0.8

Cz,

75 Comparison Between Predicted and Experimentally Measured Moment

as a Function of Lift on the Hypersonic Research Airplane at Low Speed

Including the Effect of Elevator Deflection.

HRA ; C,_vs. Cc ;M=0.2;6, = +5 ° ,0 °,-5 ° ,-10 °,and-15 °
L

0.001

O.000

-0.001

.... po_en_iz! only

poL p]us '..e.. s_ct]o=

..... po;. pius i.e.r.ad si,_e e_ge e_e¢_

t'_ . "

_'"_'-"__ 0 _0

,,,,I,,,, l,l,,l,,,,l,,,,l

-5 0 5 10 15 20

Angle of at_t.k - degrees

Fig. 76 Comparison Between Predicted and Experimentally Measured Side Force

Due to Aileron Deflection on the Hypersonic Research Airplane at Low

Speed Including the Effect of Different Modelling Techniques.

HRA ; Cy_. vs. a ; M = 0.2

C_6.,

0.001

O. 000

-0.001

potcr._ia] oni._"

. ..... po'. pi_ I.e. and side e6ge e_e¢_s

• o 0 -0 -0

-5 0 5 10 15 2O

Amg]e of at:ack .-,. ,5.e&,'r_

Fig. 77 Comparison Bet-,een Predicted and Experimentally Measured Yawing

.Moment Due to Aileron Deflection on the Hypersonic Research Airplane

at Low Speed Including the Effect of Different Modelling Techniques.

HRA ; C,_ vs. a ; M = 0.2
O. 002

0.001

O. 000

.... potentia] on}y

------1. pot. pl_ l.e, suction

...... pot. plus I.e. _d side edge efi'e¢_

.

B

• o •

o °

"tlllJt l

-5 0

11 I,,,, I,,,, I,,,, I

5 10 15 20

.A.ugleof attack ~ d_'t_

Fig. 78 Comparison Between Predicted and Experimentally Measured Rolling

Moment Due to Aileron Deflection on the Hypersonic Research Airplane

at Low Speed Including the Effect of Different Modelling Techniques.

HRA ; C16. vs. _ ; M = 0.2

0. 004

O. 002

O. 000

.... pote_ti,] eniy

--------- pot. plus I.e. suction

...... pot. plus l.z. _4_d side edge efi'ec:.s

........... . . . .

7' 0 0 0 0 O

,, ,, 1,_,,, I,,,, I,,,, I , ,,, I

-5 0 5 10 15 20

Angleof_:a_ ~ de_'ees

Fig. 79 Comparison Between Predicted and Experimentally Measured Side Force

Due to Rudder Deflection on the Hypersonic Research Airplane at Low

Speed Including the Effect of Different Modelling Techniques.

HRA ; C}%. vs. ,_ ; M = 0.2
- =

CR._ 9

0.000

-0. 002

-0.004

m

-5

.... pmentia! vniy

pot. plus ].e. suction

..... po:. plus I.e. andsideefigeeffects

0 0 0 0 0

j i ii ...... . . . -

,,,, I, I,, I,,,, I,,,, l,l, , I

0 5 10 15 20

Angle of attack ~ degrees

Fig. 80 Comparison Between Predicted and Experimentally Measured Yawing

Moment Due to Rudder Deflection on the Hypersonic Research Airplane

at Low Speed Including the Effect of Different Modelling Techniques.

HRA ; C,_, vs. _ ; M = 0.2

Fig. 81

0.001

0.000

-0.001

.... po:entiz] only

--------- ._o_.plus Le. sue:ion

..... pot. l_i'cs1.e.endsideedgee_ec_s

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

0 0 0 0 0

Ill i , I,,,,I,,,, I,,,, I,,,, t

-5 0 5 10 15 20

.tngieof attack ~ degrees

Comparison Between Predicted and Experimentally Measured Rolling

Moment Due to Rudder Deflection on the Hypersonic Research Airplane

at Low Speed Including the Effect of Different Modelling Techniques.

HRA ; Ct_, vs. a ; M = 0.2
C£

1.000 --

Fig. 82

0.750

0.500

0.250

0.000

-0. 250

-0.500

-- 0

exp. pred. 6, 0 Z_

0 0 ° A m"

-- A 10" O /

_ ,,0o 0 _ /

u " 0 ._a/"'u ./'/

A.ff

-5 0 5 10 15 20

A_gle of attack -- degrees

Comparison Between Predicted and ExperimentalIy Measured Lift on Hypersonic Research Airplane at High Subsonic Speed Including the Effect

of Elevator Deflection.

HRA; CLvs. a;M=0.8;6_ =0 °,-10 °,and-20 °

CD

Fig. 83

0.3

0.2

0.1

0.0

exp. pred. 6. 0

F 0 o°

- A ..... ,o" /#"

[].... =o. ?/°

. • o"-_ .

-5 0 5 10 15 20

Angte of attack _ 6eg-ree5

25

Comparison Between Predicted and Experimentally Measured Drag on

Hypersonic Research Airplane at High Subsonic Speed Including the Effect o[ Elevator Deflection.

HRA ; CD vs. a;M=0.8;5_ =0 °,-10 °,and-20 °
!

CD

0,5

0.2

0.1

0.0

- exp. pred. _, 0

- ////0 0

,/

1 I J I l I J ] l I J

-0.4 -0.2 0.0 0.2 0.4 0.6 0.8

CL

Fig. 84 Comparison Between Predicted and Experimentally Measured Drag as a

Function of Lift on the Hypersonic Research Airplane at High Subsonic

Speed Including the Effect of Elevator Deflection.

HRA; CD vs. CL ;M=0.8;_e =0 °,-10 °,and-20 °

Fi_. 85

C_

0.10

0.05

0.00

-0.05

-0. 10

A "-... W D

A A" " _ . __..

0

0

r__ o 0° o

- _ ..... 10" 0

E_ 20" 0

,,,,I,,ilt,,,,!,,,,I,,,,T,,,-I

-5 0 5 10 15 20 25

Angle of _t=ack _ degrees

Comparison Between Predicted and Experimentally Measured ._Ioment

ou Hypersonic Research Airplane at High Subsonic Speed Including the

Effect of Elevator Deflection.

HRA ; Cmvs. a;M=0.8; 5e =0 °,-10 °,and-20 °
_w

=--

CTrl

0.10

O. 05 --_

0.00 -

-0.05 -

-0.10

-0 4

D _

h-,,,, o _

exp. prea. & - _ ,,

O O" _ a

•'_ ..... lO° 0

D 20° O

p I , I , I _ [ L I , !

-0.2 0.0 0.2 0.4 0.6 0.8

Cz,

Fig. 86 Comparison Between Predicted and Experimentally Measured Moment

as a Function of Lift on the Hypersonic Research Airplane at High Subsonic Speed Including the Effect of Elevator Deflection.

HRA ; C,_vs. CL ;M=0.8;_,-0 °,-10 °,and-20 °

w

C};.

o. OOl

0.000

-0.001

m

) O O -"

o

potc._ti= _.o.-.]y

?_:. pi,'.s !.t. sue.ion

po'. pi,'.s i.e. _.nd side cdBe er,,ec_s

°

_..o._o_o o o o o

,I,l,,,,l,,L,T,x,,I,,,,l

-5 0 5 10 15 20

Angle of a:-ack~ degrees

Fig. 87 Comparison Between Predicted and Experimentally .Measured Side Force

Due to Aileron Deflection on the Hypersonic Research Airplane at High

Subsonic Speed Including the Effect of Different Modelling Techniques.

HRA ; Cv_,, vs. a ; M = 0.8
0.001

C. 000

-0.001

.... ;..o_en:iz! on}y

po=...',ius La. s_c:iom

-" " - °

_ u C"''O © 0 0 0 0 0 0 0

,,,, I LI_, I,,rl Illlz l,ii jl

-5 0 5 10 15 20

Ang}e of a_zsc.k ~ derrees

Fig. 88 Comparison Between Predicted and Experimentally Measured Yawing

Moment Due to Aileron Deflection on the Hypersonic Research Airplane

at High Subsonic Speed Including the Effect of Different Modelling Techniques.

HRA ; C... vs. a ; M = 0.8

O. 002

0.001

0.000

.... po_en;iz! ._ni.v

po:. piu, '..e. s_¢:io---

..... ?o:. pi'_ !.-. a=_ si_e e_ge e£e:u

. o "

, o

I-, I_l I,,,, I, ,,, I,,,, L,,' , I

-5 0 5 10 15 2O

.Engle of attack--degrees

Fig. 89 Comparison Between Predicted and Experimentally Measured Rolling

Moment Due to Aileron Deflection on the Hypersonic Research Airplane

at High Subsonic Speed Including the Effect of Different Modelling Techniques.

HRA ; Ct_., vs. c_ ; M = 0.8
CL

1. 000

O. 750

O. 500

0.250

0.000

-0.250

-0.500

I exp. pred. 6, 0

o _ o

_

_

I/'_, , , I ,, , , I , ,, , I , , , i I , i , , I , , ,

-5 0 5 10 15 20

._ileof a._:k ~ degrees

Fig. 90 Comparison Between Predicted and Experimentally Measured Lift on

Hypersonic Research Airplane at Transonic Speed Including the Effect of

Elevator Deflection.

HRA;CLvs. a;M=0.98;6e =0 °,-10 °,and-20 °

CD

0..3

0.2

0.1

0.0

A

_- ,:<p.pred. & O_

0 0° 0

- _ -10 ° _ /

-- El -20" D/

f- a .

i

_,,_ _/.,-5-"

:,o-...__, _..-L<..>"

-5 0 5 10 '5 20

Angle of a_:ack ~ de_rees

i

! I

25

Fig. 91 Comparison Between Predicted and Experimentally Measured Drag on

Hypersonic Research Airplane at Transonic Speed Including the Effect of

Elevator Deflection.

HRA ; CD vs. a;M=0.98;6, =0 °,-10 °,and-20 °
CjD

0.2

0.2

0.1

exp. pred. _,

0 0°

_. 10 °

D -_-0 ° /z_

/.X

o.o /, I , I , L._, I , 1 • I

-0.4 -0.2 0.0 0.2 0.4 0.6 0.8

13

O

C£,

A

O

Fig. 92 Comparison Between Predicted and Experimentally Measured Drag as a

Function of Lift on the Hypersonic Research Airplane at Transonic Speed

Including the Effect of Elevator Deflection.

HRA;CDvs. CL ;M=0.98;6e =0 °,-10 °,and-20 °

0.10

0.05

0.00

-0.05

D

-0.10 i i i i I i J , t

-5 0

exp. pred.. _, 0 _ []

0 o° 0 '_

..... I0°

0

20° 0

I,,_,+,,,,f,,,,I,,,, _..

5 10 15 20 25

AngJe of ax:azk _ degrees

Fig. 93 Comparison Between Predicted and Experimentally Measured Moment

on Hypersonic Research Airplane at Transonic Speed Including the Effect

of Elevator Deflection.

HRA ; Cm vs. a;M=0.:8 ; 6e =0 °,-10 °,and-20 °
OR

0.10

O. 05

0.00

-0.05

-0.10

-0.4

- "o- 2 o

"8 "e.&.,. - []

_xp. p,-,,d. _, ""m._ o

" 0 , o" "'0"<"

.% lO °

[3 _ _ -20" 0

, I , I L r , [ , I , !

-0.2 0.0 0.2 0.4 0.6 0.5

C/,,

Fig. 94 Comparison Between Predicted and Experimentally Measured Moment

as a Function of Lift on the Hypersonic Research Airplane at Transonic

Speed Including the Effect of Elevator Deflection•

HRAiCmvs. CL ;M= 0.98;6, =0 °,-10 °,and-20 °

"' %"

0.001

O. 000

-0.001

.... pu:e.'.qz! _niy

po:. ?}'_!.t. s_::3o=

..... po'. p}".._!.:. zar/ ,,iciee_e effe:',s

- ,.. "'" _-__-g-O--O--q-om

-5 0 5 10 15 20

.-ingleOf a.t_&ck _ de_rees

Fig. 95 Comparison Between Predicted and Experimentally Measured Side Force

Due to Aileron Deflection on the Hypersonic Research Airplane at Transonic Speed Including the Effect of Different Modelling Techniques.

HRA ; Cy_o vs. a ; M = 0.98
;.i V* •

O. 000

-0.001

-0.002

oo o O"-P

,,l , [I I I I [1

-5 0 5

?o_e.'.:i_.!oni::

?o:. _i'_ '..t. s==:ic=

0 0 0 O

o 0

, ] ' J , ' ,'J [ ' ] | I

10 15

Angle of attack _ desrees

I

2O

Fig. 96 Comparison Between Predicted and Experimentally Measured Yawing

Moment Due to Aileron Deflection on the Hypersonic Research Airplane

at Transonic Speed Including the Effect of Different Modelling Techniques.

HRA ; C,_6= vs. a ; M = 0.g8

0.002 --

0.001

0.000

poten;;Zor.i.v

po=. :_]=,. i.e. s'.'::ion

..... pot. pi_ I.e. _=_ side edge eSe¢_a

o

-, . -

* " . • - " "

i

,,,,I,illl,,,, I,,,,I,_,, I

-5 0 5 10 15 20

Angle of at:ack ~ degrees

Fig. 97 Comparison Between Predicted and Experimentally Measured Rolling

Moment Due to Aileron Deflection on the Hypersonic Research Airplane

at Transonic Speed Including the Effect of Different Modelling Techpiques.

].:RA ; Cl,_ vs. a ; M = 0.98
1.000 --

exp. pred. 6, O. 750 --

0 0_

..... zoo 9+_ d,,,%_

- []..... 20" _'_-_

O. 500 _ . ./

C'z. 0.250 - I" " /

°.°°° -0. 250

-0.500 PW, , , I , , , , I , , , I I , , , , I , , , i I i , , , Y

-5 0 5 10 15 20 25

._gle of a.U,_r._~ dellrees

Fig. 98 Comparison Between Predicted and Experimentally Measured Lift on

Hypersonic Research Airplane at Supersonic Speed Including the Effect

of Elevator Deflection.

HRA ; CL vs. _ ; M "- 1.20 ; 6, = 0 °, -10", and -20 °

CD

0.3

0.2

0.1

0.0

exp. pred. 6,

O 0° O_

& ..... I0 °

i

D -20° p /

\\ /L7

['_ _l I f, , ,, !, , i, r, _ , _ f,, i j I , , _ ! .

-5 0 5 I0 15 20 _5

Angie of_:_ack ~ degrees

Fig. 99 Comparison Between Predicted and Experimentally Measured Drag on

Hypersonic Research Airplane at Supersonic Speed Including the Effect

of Elevator Deflection.

HRA ; Co vs. a;M= 1.20;6, =0 °,-10 °,and-20 °
CD

0.3

0.2

0.1

0.0

I exp. o pred. -- E,o• /,_,.b,_)

A ..... i0•

-- [].... 20° '7 O

- ///'o'_

.___ i///

- "%_ _ o

T I ..,.I , I ' I ._. I , ]

-0.4 -0.2 0.0 0.2 0._ 0.6 0.8

Or_

Fig. 100 Comparison Between Predicted and Experimentally Measured Drag as

a Function of Lift on the Hypersonic Research Airplane at Supersonic

Speed Including the Effect of Elevator Deflection.

HRA ; C D vs. C L ; M = 1.20 ; 6e = 0 °, -10 °, and -20 °

Fig. 101

0.10

O. 05

O. O0

-0.05

-0.10

F..A 13 ,.-,

k _ o "2"-_-- ,_ o _" -_ _

F_ "_' 0 "" z_-- a _

L _o o a'_a _-

[- exp.._:ed. _, "",4,....,,,_ C]

I_ . 20°

l-I ' T I _I l ' " l 'i r I I I I ,, I L', I ' I I i ._ .' I !

-5 0 5 10 15 20

Ang}e of a_:.ack ~ degrees

I

' I

25

Comparison Between Predicted and Experimentally Measured Moment

on Hypersonic Research Airplane at Supersonic Speed Including the Effect of Elevator Deflection.

HRA ; Crnvs. a;M= 1.20;6e =0 °,-10 °,and-20 °
0.10

0.05

O.GO

-0.05

• ,,,

"_. ,-,

.o

e.,:b, pred. _, _.,- 0"_ [] D

,_ 10_ _0

D 20 °

-o. lo . ,, . I , I , J j I _ t .,_ I

-0.4 -0.2 0.0 0.2 0.4 0.6 0.8

CL

Fig. 102 Comparison Between Predicted and Experimentally Measured Moment

as a Function of Lift on the Hypersonic Research Airplane at Supersonic

Speed Including the Effect of Elevator Deflection.

HRA ; C._vs. Cc ;M= 1.20;6e =0 °,-10 °,and-20 °

O. 002

0.001

0.000

.... po{en',ia; oniy

oo:. "qu_1.e. sue:ion

po:.piu-_ Le._nd sideedgeeffec:s

F

0 © -C- - " - --

-5 0 5 10 15 20

AngJe of at:ack _ degrees

Fig. 103 Comparison Between Predicted and Experimentally Measured Side Force

Due to Aileron Deflection on the Hypersonic Research Airplane at Supersonic Speed Including the Effect of Different Modelling Techniques.

HRA ; C}_ vs. a ; M = 1.20
t.,s

O. 000

0

-0.001

0"0

po:. r,iu.* Le..*u::ion

..... po-. p]=s i.e. -'.rid sit_e edge e.ffec:_

0 0 0 Q n _ 0 o

, , t I J I I t i _,1 , , _ I , , I -0.002 _ z,, I , J

-5 0 5 10 15 20

Angle of _t.t, ac.k ~ degrees

Fig. 104 Comparison Between Predicted and Experimentally Measured Yawing

Moment Due to Aileron Deflection on the Hypersonic Research Airplane

at Supersonic Speed Including the Effect of Different Modelling Techniques.

HRA ; C,,. vs. a ; M = 1.20

0.002

0.001

O. 000

.... pote,-,:i_.::t.]y

po'. ?iu._ 2._.s_c:ior,

..... pot. piu_ .Le. t.ni side edge effe:_

. ° o

booo oo 0 0 0 0 0

I

illlt

-5

,I, ::, Ill i, I,,,, I :,, , l

0 5 10 15 20

.&ngie of at_a¢i: _ degrees

Fig. 105 Comparison Between Predicted and Experimentally Measured Rolling

Moment Due to Aileron Deflection on the Hypersonic Research Airplane

at Supersonic Speed Including the Effect of Different Modelling Techniques.

HRA ; Cir. vs. a ; M = 1.20
O. 4O0

0.300

0.200

0.100

0.000

-0.100

I j

e>:,.,, pred. L, /'_r

_¢ " "

/.7 ._

! o :'_;Y

. _. ...... s,< . _/"

0 .... ]o:, ._/. • _/

_ v ,-- -1_. /..2jj"

t x ,!tT,l.i I_ i , i! q , i[I ,,,_ I

-5 0 5 10 15 20

Angle of at_ck ~ degrees

Fig. 106 Comparison Between Prediction Without Shielding and Experimentally

Measured Lift on the Hypersonic Research Airplane at Hypersonic Speed

Including the Effect of Elevator Deflection•

HRA;CLvs. a;M=6.0;de =+10 ° , 5 ° ,0 °,-5 °,-I0 °,and-15°;

without shielding

CD

0.20 --

0.15

ezp. pred. E,

-- . -- .I0 c

/

[] ....... _,_

c oo

O. 10 0 ]o-" 1.

/.

• ,+

0.05 !l__

0.00 F , , I .... I , , I , , , , I , , , , I

-5 0 5 10 !5 20

Angle of &ttxck ~ degrees

Fig. 107 Comparison Between Prediction Without Shielding and ExperimentaIly

Measured Drag on the Hypersonic Research Airplane at Hypersonic Speed

Including the Effect of Elevator Deflection.

HRA ; CD vs. a ; .M = 6.0; _e = +10 ° , 5°- 0°, -5°, -10°. and -15°;

without shielding
lJ

m

CD

0.20

0.!5

0.10

0.05

F

exp. pre4. _,

z, -,1C)"

/

0 0° 2'

"- _ ---- I,, _,*¢_

-0.1 0.0 O. 1 0.2 0.3 0.4

C.L

Fig• 108 Comparison Between Prediction Without Shielding and Experimentally

Measured Drag as a Function of Lift on the Hypersonic Research Airplane

at Hypersonic Speed Including the Effect of Elevator Deflection.

HRA; CDvs. CL ;M=6.0;Se = ,-i-10 °,5 °,0 °,-5 ° ,-10 °,and-15°;

without shielding

exp. pred• _,

A _10 °

E

0.02 m ...... _"

0 ...... 0

0.01 Lr._ _ 0 ....... 10=

c_ °'°° k --_" _iZ "--_.__-'-e. ---=

"'--_ ""-2 _"----L._- -" - .o

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

IO o _ _

F

E

-0.02 rE

-0.o3 I-, I I , , f I !

,,' ' _ I ' ' _ _ _ ; , ; I 1 ,' ',

-5 0 5 10 15 20

Angle of attack _ degrees

Fig. 109 Comparison Between Prediction Without Shielding and Experimentally

Measured Moment on the Hypersonic Research Airplane at Hypersonic

Speed Including the Effect of Elevator Deflection.

HRA ; Cm vs. a ; M = 6.0; 5e = +10 °, -5°: 0 °, -5 °. -10 °. and -15°;

without shielding
0 G2

0.01

C_ 0.00

-0.01

-0.02

-0.03 ;

-0. 1

exp. pred,

F [] ......

-10"

O'

_0_-

-15 _

, I I I i I , I I I ' r a I _ I , , I r , , LJI

-0.100

0.0 0.1 0.2 0.3 0.4

CL

Fig. Ii0 Comparison Between Prediction Without Shielding and Experimentally

Measured Moment as a Function of Lift on the Hypersonic Research

Airplane at Hypersonic Speed Including the Effect of Elevator Deflection.

HRA;Cmvs. CL;M=6.0;6e =+I0 °,5 °,0 °,-5 ° ,-I0 °,and-15°;

without shielding

0.400 -- exp. pred. _,

A -- " -- "+I0"

D ...... + 5" /"

o. 300 - o o, .

* - ,° .4 7

Cz 0.200 -- v -15 ° .f._"

O. 100

O.OOO

1 , } : I , , ; =', ,,, , : I ' I 1 i

-5 0 5 10 15 20

Angie of attack~ degrees

Fig. 111 Comparison Bezween Prediction With Shielding and Experimenta]]y Measured Lift on the Hypersonic Research Airplane at Hypersonic Speed Including the Effect of Elevator Deflection.

HRA ; CL vs. a ;M = 6.0; 6_ = +10 °,5 °,0 °,-5 ° , -10 s , and-15°;

with shielding
CD

0.!5

0.10

0.05

O. O0

i e_:p. pred. /._ /

i ..... /..j. ...... 5_ . ./',_

- ' .'7¢

....

-5 0 5 10 15 20

Angle of attack - degrees

Fig. 112 Comparison Between Prediction With Shielding and Experimentally Measured Drag on the Hypersonic Research Airplane at Hypersonic Speed

Including the Effect of Elevator Deflection.

HRA; CDvs. a;M=6.0;6_ = +10 °,5 ° ,0 °,-5 °,-10 °,and-15°;

with shielding

CD

0.15

0.10

0.05

m

._:_. pred. Er z_ -/

#. -.I0° _//

1:3 -"..... "_ 5_

0 ..... 0° //_

- _ ......... 5_

L

--_.:.00 •

t , , i I z I : : I : , , , , , : _

-01 • 0.0 O_, • 02, 0 . _,_ 0.4

C'r

Fig. 113 Comparison Between Prediction With Shielding and ExperimentaIly hfeasured Drag as a Function of Lift on the Hypersonic Research Airplane at

Hypersonic Speed Including the Effect of Elevator Deflection.

HRA ;CD vs. CL ;M=6.0;5_ =-rl0 °,5 ° .0 °,-5 °,-10 °,and-15°;

with shielding
0.02

0.01

0. O0

C,,

-0.01

-0,02

-0.03

exp. pred. _,

A .... r ]0_

C' [:'

_:.,..._.... 0 ]O"

_-':. L..._ -- - ..__.___..... _ r°

- Z "'-. T.

- _._ .... -.._

"- k

I,,Jl,,,,l,,i,IL,,,lj,j,J

-5 0 5 10 15 20

Angle of attack ~ degrees

Fig. 114 Comparison Between Prediction With Shielding and Experimentally Measured Moment on the Hypersonic Research Airplane at Hypersonic Speed

Including the Effect of Elevator Deflection.

HRA •, Cm vs. a ; M = 6.0; 6e = ,-4-10° , 5 ° , 0 °, -_°," -10 °, and -15°;

with shielding

exp. pred. 6t

A +I0°

O. 02 r_ ...... + s"

0 .... o"

0.01 _1' ..... 5 c

.... 10c

_ .... 15 °

O. O0 C,.n \-5.

-0.01 -- "_ ,a, "" " " _'_3"-

"_A -.

-0.02 ,-- _-

-0.03 , , . . . I _ . , ,

-0. 1 O.O 0. i 0.2 0.3 0.4

C:

Fig. 115 Comp&rison Between Prediction With Shielding and Experimentally Measured Moment as a Function of Lifton the Hypersonic Research Airplane

at Hypersonic Speed Including the Ef[ect of Elevator Deflection.

HKA ; C,_ vs. Cr ; M = 6.0; 6c = _-10c, 5°, 0°, -5 °, -10 _, and -15°;

with shielding
C}:_,

0.0302

0.0001

0.0000

0

--- withou: s?.ielding

...... with shielding 0

0 ......... 0 0

-5 0 5 10 15 20

Angle of attack _ de_rees

Fig. 116 Comparison Between Prediction With and Without Shielding and Experimentally Measured Side Force Due to Aileron Deflection on the Hypersonic Research Airplane at Hypersonic Speed.

HRA ; Cy_. vs. a ; M = 6.0 ; with and without shielding

v_l_o

0.0000 --

-0.0002

-0.000_

..... C) ° -

-- w_:ho'.:: s::ieiE.L-.g

i ...... v,'l:._".";_e_cl_g

' _ _ , •I I ! } { I r , I, , I ¢ , ,I ,

-5 0 5 10 _5

Angie of a::a:i: _- degree.,

O

©

,, I

20

Fig. 117 Comparison Between Prediction \Vith and \Vithout Shielding and Experimentally Measured Yawing Moment Due to Aileron Deflection on

the Hypersonic Research Airplane at Hypersonic Speed.

HRA ; Cn_, vs. a ; M = 6.0 ; with and without shielding
C_¢o

0.001 --

0.000

-0. O01

wi:.hou: -_hieldinG

I( ..... ,,

...... WII_. S_',, T,el ¢. In g

- o --'v----e----°

_) . .':"*: . - . _ .... . . . ...........

I-',,, I:.:,, Ill,, lts,i FIz:_ I

-5 0 5 10 15 2O

Angle of _ttack ~ degrees

Fig. I18 Comparison Between Prediction With and Without Shielding and Experimentally Measured Rolling Moment Due to Aileron Deflection on the

Hypersonic Research Airplane at Hypersonic Speed.

HRA ; Clio vs. a ; M = 6.0 ; with and without shielding

Cy,,

0.001

0.000

-0.001

I

I

E

]

r'._

-5

0 0 0

whhm:" shi-.id!n_.

I l , t l y I

5 _0 !5

._._.g;e of a_ac_: ~ degrees

,I

2O

Fig. 119 Comparison Between Prediction With and Without Shielding and Experimentally Measured Side Force Due to Rudder Deflection on the Hypersonic Research Airplane at Hypersonic Speed.

HRA ; Cy_ vs. a ; M = 6.0 ; with and without shielding
Cn_

O. 000

-0. Of-,

-0. 002

0 0 b

I-" _ whheu: -hie]dinl;

-5 0 5 10 15 20

Angle of au,ack ~ de_;ree_

w

Fig. 120 Comparison Between Prediction With and Wkhout Shielding and Experimentally Measured Yawing Moment Due to Rudder Deflection on

the Hypersonic Research Airplane at Hypersonic Speed.

HRA ; C,t, vs. a ; M = 6.0 ; with and without shielding

0. 0002

0.0001

0.0000

wi;hou, e'_: ' ";-,e

_ . . , "',

t..-- ©

/

1

[ c @

_" I",":.I

-5 0 5 10 !5 20

kngieo_at_ac_:_ degree_

Fig. 121 Comparison Between Prediction V¢ith and Without Shielding and Experimentally Measured Rolllng Moment Due to Rudder Deflection on the

Hypersonic Research Airplane at Hypersonic Speed•

HRA ; Cl,:_ vs. a ; M = 6.0 ; with and without shielding
C_'lo

0.001

O. 000

-0.001

e>:i- !.,red, u

[3 12_

'OL"....... -'-" z - -.

w

i I , t I 1

2 4 6

Mach number

Fig. 122 Comparison Between Predicted and Experimentally Measured Side Force

Due to Aileron Deflection on the Hypersonic Research Airplane as a Function of Mach Number at Two Angles of Attack.

HRA ; Cry,, vs. M ; a = 0 ° and 12 °

0.001

O. 000

-0.00_

e:::,, pred.

0 ........ O_

[] ...... !2 _

e%

0 2 " 6

Fig. 123 Comparison Between Predicted and Experimentally Measured Jawing

Moment Due to Aileron Deflection on the Hypersonic Research Airplane

as a Function of Mach Number at Two Angles of Attack.

HRA; C._ vs. M ;a= 0 ° and 12 °
C_ a

O, 002

0.:30 _

0.000

e::!,, pred, a

0 ........ O_

[] 1 "..'_

, ......t , 1 ,'-,- g

0 2 4 6

Math number

Fig. 124 Comparison Between Predicted and Experimentally Measured Rolling

Moment Due to Aileron Deflection on the Hypersonic Research Airplane

as a Function of Mach Number at Two Angles of Attack.

HRA;Cz_, vs. M;a=0 °and 12 °

0.004 F.''" ,

C_'. I " "

0.002

C,.O00 0.. !

o .

"0

i t

0 2 _- 6

.Mac_ n u.--.,be:

Fig. 125 Comparison Between Predicted and Experimentally Measured Side Force

Due to Rudder Deflection on the Hypersonic Research Airplane as a Function of Mach Number.

HRA; Cv_,. vs. M ; a =0 °
o.ooo f °°° " "

-0. 002 . - - "

t "

-0.004 , I _ I

.0

0 2 4 6

Mach number

Fig. 126 Comparison Between Predicted and Experimentally Measured Yawing

Moment Due to Rudder Deflection on the Hypersonic Research Airplane

as a Function of Mach Number.

HRA;C_, vs. M;a=0 °

0.00: _. ....

- _:9 0 " .... 0

_':'" o. coo E- co

L- i , t , ! i

-0. O0"

0 2 4 5

h'achnu.'nber

Fig. 127 Comparison Between Predicted and Experimentally Measured Rolling

Moment Due to Rudder Deflection on the Hypersonic Research Airplane

as a Function of Mach Number.

HRA; CI_ vs. M;a=0 °
ORIGINAL PAGE IS

OF POOR QUALITY

Oz.

2

-1

EL

-5

t :.;,, pre".. E,

F, -IC:

,...,

.... ] q:

_,.-_ .._ .i._ _ ...-"

.-'''t'_._-I" ..9.-'"

. -

° -

o J

° •

0 5 10 15 2O

.kngie of a.t.t,_.ck,-, 8e@'rees

Fig. 128 Comparison Between Predicted and Experimentally Measured Lift on

Shuttle Orbiter at Low Speed Including the Effect of Elevator Deflection.

Shuttle ; CL vs. c_ ; M = 0.25 ; 5_ = +10 °, 0 °, -10 °, and -20 °

C_

0.4

0.3 --

0.2

0.1

C , t_w

¢

exp. pred. _= p

<> +i0 = _ 0

0 O° "

F

L _ -10 = ,

L _ ........ 20" ,' //20

F /,/ /

z / / / - /A

e"-.. ?'/ .--'"..-

,p_. --.. ,z -.. _._-_.._.__ /..-

;_-.-:.___ ---......_ :_ . _ :- •

I

u _ ; i I ' . . . 1 . i i r f p I _ i _ : :

-_.,.., r.. ,= 10 15 25

..kngie of z-:.aci: ~ degree._

Fig. 129 Comparison Between Predicted and Experimentally Measured Drag on

Shuttle Orbiter at Low Speed Including the Effect of Elevator Deflection.

Shuttle ; CD vs. a ; M = 0.25 ; 6e = +i0 °,0 °, -i0 °, and -20 °
ORIGINAL PACE '_

OF. POOR QUALITY

CD

C_

Fig. 130 Comparison Between Predicted and Experimentally Measured Drag as a

Function of Lift on the Shuttle Orbiter at Low Speed Including the Effect

of Elevator Deflection.

Shuttle ; CD vs. CL ; M = 0.25 ; 6c = +10 °, 0 °, -10 ° , and -20 °

CTn

0.4

0.2

O.O

"" ° " ° " ° " " " " " - = ° . . . . ° ° . . . . . . °

r_. A A A

[] O ''-"',_-- -

1

exp. pred. 6,

<> +10 °

0 O°

-10 _

Z_ ........ 20°

° " " " ° - . . . . . /k

o ©

0 0

-0.2 ' ' ' '

-5 0 5 10 i5 2E

Angle of a_,a¢_: _ degrees

Fig. 131 Comparison Between Predicted and Experimentally Measured Moment

on Shuttle Orbiter at Low Speed Including the Effect of Elevator Deflection.

Shuttle ; Cm vs. a ; M = 0.25 ; 6, = +10 °, 0 °, -10 °, and -20 °
=

ORIGINAL PAGE IS

OF POOR QUALITY

Cyt_

0.4

0.2

I

0.0 -

-0.2 v

-1.0

0 _T"--4_....0 0

"_r"'_%_...._... o 0

O O"

Im - ]0 _

A ........ 20 _

1 I 1 I I ! l I 1 I ? 1 I 1 I I I I I I ! I , I

-0.5 0.0 0.5 1 .0 1 .5

C£

Fig. 132 Comparison Between Predicted and Experimentally Measured Moment

as a Function of Lift on the Shuttle Orbiter at Low Speed Including the

Effect of Elevator Deflection.

Shuttle ; Cm vs. CL ; M = 0.25 ; 6e = +10 °, 0 °, -10 °, and -20 °

0.005

.... p oI e.-.:ip-)eni.v

-- .vo-. _}u; :.,.._-=c:}c-

..... po:,. pi"_ Le. r._d side edge erie:::

,r

0.000

-0. 005

_" . " "

r _

- O

C- O 0 0

F ,, ' ; ',_, ' , I ' ' ' ; !, ) ,i ', l , , ' , '

-5 C' z%,: ,._m 10 I =`

.&ng}e of &t_ac]: "-- deg-_e- _

Fig. 133 Comparison Between Predicted and Experimentally Measured Side Force

Due to Aileron Deflection on the Shuttle Orbiter at Low Speed Including

the Effect of Different Modelling Techniques.

Shuttle ; Cy_ vs. a ; M = 0.25
ORIGINAL PAGE IS

OF. POOR QUALITY

0.00_

-o. oo 1

F

0 0

-5 0 5 I0 15 20

Angle of _ttac_ ~ degrees

Fig. 134 Comparison Between Predicted and Experimentally Measured Yawing

Moment Due to Aileron Deflection on the Shuttle Orbiter at Low Speed

Including the Effect of Different Modelling Techniques.

Shuttle ; C,_ vs. a ;M = 0.25

m

i

C'

.OlO

.005

•ooo

F

m

I

F

F !, _ I ;

-5

..-,o:en-i_) oniy

po:,pi'.'sLe, su:tio:.

..... po:.pi'.'sLs, _._ sidse_.l:ee.=.e:'.s

o

° °

° -

° °

I I

' I ' I' : I ....

, (,, ; . ,, , , : l

O 5 iO "='-- 20

Angie of _.'taCi: ~ degree-'-

Fig. 135 Comparison Between Predicted and Experimentally Measured Rolling

Moment Due to Aileron Deflection on the Shuttle Orbiter at Low Speed

Including the Effect of Different Modelling Techniques.

Shuttle ; Cz_ vs. a ; M = 0.25
ORIGINAE PAGE IS

POORQUALITY

w

.010

.025

.000

..... :,c-. ;.u; ..-.. ;:._ .,id-. ee;e ,.Se:'r

EI

r

m

r

0 0

F

-5 0 5

0 0

I,,I,III'I I

10 15 2O

Angle of atuL_ ~ 6egrees

Fig. 136 Comparison Between Predicted and Experimentally Measured Side Force

Due to Rudder Deflection on the Shuttle Orbiter at Low Speed.

Shuttle ; Cy_. vs. a ; M = 0.25

0.000 F

E

L

r i

-0.001 LI

-0. 002

.... ?o: e.-.:iz] rn]y

_:-. _._ :.e. --_ si "_-. e@: e_e:_,-_

0 O 0

- 5 0 i0 I =`

.&n[_.i_. of at:at', ~ degree- ¢

_C

Fig. 137 Comparison Between Predicted and Experimentally Measured Yawing

Moment Due to Rudder Deflection on the Shuttle Orbiter at Low Speed.

Shuttle ; C,_, vs. a ; M = 0.25
ORIGINAL PAGE IS

Of POOR QUALITY

•oo2

• oo"

• ooo

..... .:'=-'..="._.....' • ::;dsi6e edI_ee-e:'.•

O 0 0 0

.l ,, ! llij _ !111, jrll i J, t IT J

-5 0 5 _0 _5 20

Angle of _tt_ck ~ degrees

Fig. 138 Comparison Between Predicted and Experimentally Measured Rolling

Moment Due to Rudder Deflection on the Shuttle Orbiter at Low Speed.

Shuttle ; Cl_, vs. c_ ; M = 0.25

Cr

2

0

exp. pred. 6,

O -10"

0 0° ,'"

s

[] -!0 _ . - -

_ v,, ........ 20 ° . - _O

_- -0 _ / ..--

.- _.1t-,.., .I _.."

Y" / " . -

¢ P, . ..

._- ./ •

°

_f .

-5 O "_ 1"' _5 _'_

.-kngie of zt_.ac_ _ 6et:ree •.

Fig. 139 Comparison Between Predicted and Experimentally Measured Lift on the

Shuttle Orbiter at High Subsonic Speed Including the Effect of Elevator

Deflection.

Shuttle ; CL vs. c_ ; hJl = 0.8 ; 6e = -r-10 °, 0 °, -10 _. and -20 °
OF POOR QUALITY

_'D

C.6

0.4

0.2

0.0

/

P

C.: C'_ "

" C',

_-- -IC" "

A ........ 20" ,," /0

/ -

/ i e'_

i

. / r.

.'__--7-:__-_-..--:'/-_:'_.-_._..----"

-5 0 5 10 15

.A._leo'a_ack- degree-_

/

2O

Fig. 140 Comparison Between Predicted and Experimentally Measured Drag on

the Shuttle Orbiter at High Subsonic Speed Including the Effect of Elevator Deflection.

Shuttle ; Co vs. a ; M = 0.8 ; 6e = +10 °, 0 °, -10 °, and -20 °

CD

0.6

0.4

I

o._ L

I

I

r"

I

i

I

?-_ y- I

. • v

-1 .0

exp. pred. 6,

<b -10'

C 0_

.... 3o_ 0

k_ ........ 20° 0 /'

, . ;;:¢-

,_. ,L .f,.-""_..e#

_=. :-,,___._z -_

-_,.5 0.0 n 5 ".,0

I

I

I

I

C:.

Fig. 141 Comparison Between Predicted and Experimentally Measured Drag as a

Function of Lift on the Shuttle Orbiter at High Subsonic Speed Including

the Effect of Elevator Deflection.

Shuttle ; CD vs. CL ; M = 0.8 ; 6, = ÷10 °, 0 °, -10 °. and -20 °
ORIGINAL PAGE IS

OF POOR QUALITY

0._

. ° . , _ ° . ° ° , . . , . - . ° . .

exJ,. !srt-_. c,

.,........ lo_

:":_ .... ]U-

° . . ° . °

, . . . . .

• ..,....

] A " "--- ._...

[3 O [] " ---- .A

.m

I 0

0.0 -- 0--_'-""_-_-.._.__ 0

"----0.___ 0 0 0

-0.2 i, I l..l !', T I I I, ! (,, I ! I. I 5"_,,-I

-5 0 5 10 15 20

Anl;leof a:_ack ~ degrees

Fig. 142 Comparison Between Predicted and Experimentally Measured Moment

on the Shuttle Orbiter at High Subsonic Speed Including the Effect of

Elevator Deflection.

Shuttle ; Cm vs. a ; M = 0.8 ; 6e = +10 °, 0 °, -10 °, and -20 °

=

Fig. 143

Cy1_

0.6

C.4

0.2

0.£,

-_.2

r

F ° . ° . . . .

VF

1

V

, . • !

°

.

exp. pred. 6,

K'> +10 °

0 0_

-10 °

L ........ 20 =

r- /" & """ " _-

_.- ,._ -- Z_ " --..... .

0" "-0.. C. (> 0

...... : - T'_-'

-1.C -0 5 C.O 0.5 ;.C' :.5

C:

Comparison Between Predicted and Experimentally Measured Moment

as a Function of Lift on the Shuttle Orbiter at High Subsonic Speed

Including the Effect of Elevator Deflection.

Shuttle ; Cm vs. Cc ; M = 0.8 ; 6e= -10 °, 0 °, -10¢: and -20 °
ORrGfNALPAGE IS

Of POOR QUALITy

Cy_,

0.005

0.000

-0.005

.... ".'_lCLl;r. (''; *

..... •.,o-, "..m, .._..::;C sioe cdI_e e£e::._

. _ _L'_.__.- _ .... _

0 0 0

--©

-5 0 5 I0 15 20

Angle of attack _ degrees

Fig. 144 Comparison Between Predicted and Experimentally Measured Side Force

Due to Aileron Deflection on the Shuttle Orbiter at High Subsonic Speed

Including the Effect of Different Modelling Techniques.

Shuttle ; Cy_, vs. a ; M - 0.8

0.081

0.000

-2.CO_

r-"

¢,

.... ]_ol,enzhL)on]v

--. po:. y,ius i.e. suc:io.-.

..... poz. pius ;.a. and side edge effe:t._

O O

" " - --''"_- " .... P..... 0

I ,,,1 ,,_ I , I .... :

A;._ie c."a::,ack ~ degreez

Fig. 145 Comparison Between Predicted and Experimentally Measured Yawing

Moment Due to Aileron Deflection on the Shuttle Orbiter at High Subsonic Speed Including the Effect of Different Modelling Techniques,

Shuttle ; C,_o vs. a ; M = 0,8
ORIGINAL PAGE IS

OF POOR QUALITY

C_r o

pc:...r.i'.',, he. ;::c side ed;e e_e:;. •

.010 --

F ,

.000 _ , , , , , , = , I I , , , I , , , i I,.I I ,

.

. .

0

-5 0 5 10 15

An_-ie of at.=a.ck ~ degrees

0

i

2O

Fig. 146 Comparison Between Predicted and Experimentally Measured Rolling

Moment Due to Aileron Deflection on the Shuttle Orbiter at High Subsonic Speed Including the Effect of Different Modelling Techniques.

Shuttle ; Cz_= vs. a ; M = 0.8

Cyl,

.006 i"-

1.-

I

.004 _-

I

• 002 Of

r-

• OOO

r_ C' 0 C'

i r : | , i , _ I , , I , • 1

-: [_ =" 1 _ "D _'_

.-.-.;Je of at:acJ: ~ degrees

Fig. 147 Comparison Between Predicted and Experimentally Measured Side Force

Due to Rudder Deflection on the Shuttle Orbiter at High Subsonic Speed.

Shuttle ; Cy_,. vs. a ; M = 0.8
ORIGINAL PAGE f$

OFPOOR QUALITY

0.000

C,., -0.00_ ,_

.

-0. 002

-0.003

0 0 0 0

-5 0 5 10 15 20

AngJe of au.ack _ de_rees

Fig. 148 Comparison Between Predicted and Experimentally Measured Yawing

Moment Due to Rudder Deflection on the Shuttle Orbiter at High Subsonic Speed.

Shuttle ; C,_, vs. a ; M = 0.8

C

0.002

F

0.001

-C'. OC_

o 0 o o

I I l

-5 O =, 10 "5

.&n_}e of zt:ack _ 6.e_re-_

Fig. 149 Comparison Between Predicted and Experimentally Measured Rolling

Moment Due to Rudder Deflection on the Shuttle Orbiter at High Subsonic Speed.

Shuttle ; Cl6,, vs. c_ ; M = 0.8
C'L

1.5

1.0

0.5

0.0

-0.5

exp. pred. 6,

O ÷I0 °

0 0 ° ._

A ........ 20 ° ._f"

.jj

-10 O 10 20 ,:30 40

A.ugte of attack ~ degrees

Fig. 150 Comparison Between Predicted and Experimentally Measured Lift on

the Shuttle Orbiter at Hypersonic Speed Including the Effect of Elevator

Deflection.

Shuttle; Ccvs. a;M= 5.0," 6_ = +10 °,0 °,-10 °,and-20 °

CD

1 25

0

0

0

0

O0

75

5O

25

O0

_ pred. £,

ex'_.

<> +10 ° t

,/ [] -I0 ° //.'_

A ........ 20°

] l L I I I I

-10 0 10 20 30 40

Ang,_e of a_mc.k _ degrees

Fig. 151 Comparison Between Predicted and Experimentally Measured Drag on

the Shuttle Orbiter at Hypersonic Speed Including the Effect of Elevator

Deflection.

Shuttle ; CD vs. a : M = 5.0 ; 5, = ÷10 °, 0 °, --10 °, and -20 ¢
CD

I .25

I .00

O. 75

O. 50

0.25

exp. pred. _,

<> +I0 °

- 0 O" O/

[] -1o"

- _ ........ 20" • .._

..3

I l I (

0.00 0.25 0.50 0.75 1.00 I .25

CI,

Fig. 152 Comparison Between Predicted and Experimentally Measured Drag as a

Function of Lift on the Shuttle Orbiter at Hypersonic Speed Including

the Effect of Elevator Deflection.

Shuttle ; CD vs. CL ; M = 5.0 ; 5e = +10 °, 00, -100, and -20 °

C_

0.05

O. 00

-0.05

-0.10

-0.15

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

o

exp. pred. _, ".. 0

0 +10" _,

O 0° " %

%

rl -10 ° .,

........ 20° <>

, , , , I , , , , l , ,., , I , , , , I , , _ , I

-10 0 10 20 30 40

Angle of ,,-t.t.ack .-. deip'ees

Fig. 153 Comparison Between Predicted and Experimentally Measured Moment

on the Shuttle Orbiter at Hypersonic Speed Including the Effect of Elevator Deflection.

Shuttle ; Cm vs. a ; M = 5.0 ; be = -+-10 °, 0 °, -10°, and -20 °
Cl_

O. O0

-0.05

-0. I0

O.05 --

- ......" . & Z_--- A

_. o'_'-_.. D_ .

: %-'oo o ..... ..

. ex_. pred. 6_ 0 "

- _ +10 ° -

- 0 O° " -

[] - l0° 0

A ........ 20°

, l , , l , ] , , i i I , i i , I l , I l 11 I , , I

--0.15 _ i i I l

-0.25 0.00 0.25 0.50 0.75 I .00 I .25

CL

Fig. 154 Comparison Between Predicted and Experimentally Measured Moment

as a Function of Lift on the Shuttle Orbiter at Hypersonic Speed Including

the Effect of Elevator Deflection.

Shuttle ,"Crn vs. CL ,"M = 5.0 ; 6e = +10 °, 0 °, -10 °, and -20 °

ACL

0.04

0.02

0.00

-0.02

exp. pred. 6bf

0 0

_ ....... 11.5 °

D ---- ÷10.0 °

<> +22.5° . .D......

." 1:3 - ID

....0. ---'_

0

o

-. A

_LLJ , I , , , , I , , , , I , , , , I , , ' ' ]

-10 0 10 20 30 40

Angle of sttack - degrees

Fig. 155 Comparison Between Predicted and Experimentally Measured Lift on the

Shuttle Orbiter at Hypersonic Speed Including the Effect of Body Flap

Deflection.

Shuttle ; ACL vs. c_ ; M = 5.0 ; 6bf = --ll.B °, +10.0°, and +22.5 °
0.08

O. 06

0.04

ACD

0.02

O. O0

-0.02 I f I

-10

exp. pred. 6b:

Z_ ....... II.$°

D +i0.0 °

0 .+,22.5"

s

_[_'-"-" : ......... Z_...

r'J o

O

m

- o

0 10 20 30 40

Angle of_.ttack .- degrees

Fig. 156 Comparison Between Predicted and Experimentally Measured Drag on

the Shuttle Orbiter at Hypersonic Speed Including the Effect of Body

Flap Deflection.

Shuttle ; ACD vs. a ; M = 5.0 ; (_bf = -11.5°, + 10.0°, and +22.5 °

O. 04 --

0.02 -

0.00

AC,,

--O.02 --

--0.O4 --

-0.06 ' ]

-10

exp. pred. 6b_

A ....... 11.5°

D +I0.0 Q

<> +22.5 °

°

o

....

--. O ._.%.

°

13

i , J 1 I , , J , i i i I i i i I _1 I I i _

0 10 20 30 40

Az_gleofatt_k _ degrees

Fig. 157 Comparison Between Predicted and Experimental})' Measured Moment

on the Shuttle Orbiter at Hypersonic Speed Including the Effect of Body

Flap Deflection.

Shuttle ; ACre vs. a ; M = 5,0 ; 6bf = --11.5 ° , +10.0 °, and +22.5 °
1.5 --

C_

1.0

0.5

0.0

-0.5

i

,,,,I,,,,l,,,,ll,lfl,,,,I

-I0 0 I0 20 50 40

Angleof a_taci ~ degrees

Fig. 158 The Effect of Shielding on the Predicted Lift on the Shuttle Vehicle at

Hypersonic Speed.

Shuttle ; CL vs. a ; M -- 5.0 ; with and without shielding

CD

1.00 --

0.75 -

0.50

0.25

O.O0

- without shiel_in$ _0"

, , , , I , + , , I ,, ,,, , I , , , , I L, , , I

-10 0 10 20 3O 40

Angle of _t_mk *- degrees

Fig. 159 The Effect of Shielding on the Predicted Drag on the Shuttle Vehicle at

Hypersonic Speed.

Shuttle ; CD vs. a ; M = 5.0, with and without shielding
C' D

1.00

0.75

0.50

0.25

0.00

/

..... ;

-0.25 0.00 0.25 0.50 0.75 1.00 1.25

CA

Fig. 160 The Effect of Shielding on the Predicted Drag as a Function of Lift on

the Shuttle Vehicle at Hypersonic Speed.

Shuttle ; CD vs. CL ; M = 5.0 ; with and without shielding

Cm

Fig. 161

O.O0

-0.02

-0.04

-0.06

-0.08

...... with shielding

-10 0 10 20 50 40

Amgle of a:_aek ~ degre_

The Effect of Shielding on the Predicted Moment on the Shuttle Orbiter

at Hypersonic Speed.

Shuttle , Cm vs. a ; M = 5.0 ; with and without shielding
C_

0.00

-0.02

-0.04

-0.06

-0.08 LI,,I,,,IlIII,I,,,,I,LI,I,,,

-0.25 0.00 0.25 0.50 0.75 1.00

C.L

Fig. 162 The Effect of Shielding on the Predicted Moment as a Function of Lift

on the Shuttle Orbiter at Hypersonic Speed.

Shuttle ; C,_ vs. Cz, ; M = 5.0 ; with and without shielding

7.5 i

CL

1.0

0.5

0.0

exp. pred. _e

+I0"

_ o

z_ ........ 20° ._-'" "--..,_"_" "

I Izl!,,,!,,,, f,,,,l,,,,1

0 10 20 30 40

Angle of at_ac_ ~ degrees

Fig. 163 Comparison Between Prediction With Shielding and Experimentally Measured Lift on the Shuttle Orbiter at Hypersonic Speed Including the Effect

of Elevator Deflection.

Shuttle ; CL vs. a ; NI = 5.0 ; Ee = -10 °. 0 °, -I0 °. and -20 ° ; with

shielding
Fig. 164

CD

1.00

0.75

O. 50

0.25

7

m

O. O0

-10

exp. pred. _, _/_

0 _ 0° ,//(/'

/,--

.... ,/'

f -

0 10 20 30 40

A=i_}eof Ltt_ck ~ degrees

Comparison Between Prediction With Shielding and Experimentally Measured Drag on the Shuttle Orbiter at Hypersonic Speed including the

Effect of Elevator Deflection.

Shuttle ; CD vs. a ; M = 5.0 ; 6, = +10 °, 0 °, -10 °, and -20 ° ; with

shielding

C_

1.00

0.75

0.50

0.25

0.00

" exp. pred. $, ¢y/Q/

: o : - -

_ 0 _ O" ;.7'

- D .... 10 ° ,j//

" z_ 20" ,_

" , 1 _ I i ! _. I z I i 1

-0.2 0.0 0.2 0.4 0.6 0.8 1.0 1.2

G"z,

Fig. 165 Comparison Between Prediction With Shielding and Experimentally Measured Drag as a Function of Lift on the Shuttle Orbiter at Hypersonic

Speed Including the Effect of Elevator Deflection.

Shuttle ; CD vs. C L ; M - 5.0 ; _, = --riO v, 0 ¢, -100, and -20 ° ; with

shielding
C_N

0.05

0.00

-0.05

-0.10

-0.15

Z

J

-10

. ° = - o ° ° . . .

. °

..... £ - -.. o

O" O ".. ?

exp. pred. 6, "

O

O +10"

0 0 °

.... lo" 0

Z_ ........ 20°

'' I I J,_,, I, ,,, Ii I, II,,, I I

0 10 20 30 40

•_ml_leof at_ae.k~ decrees

Fig. 166 Comparison Between Prediction With Shielding and Experimentally Measured Moment on the Shuttle Orbiter at Hypersonic Speed Including the

Effect of Elevator Deflection.

Shuttle ; Cm vs. a ; M = 5.0 ; 5e =-rl0 °,0 °, -10 °, and -20 ° ; with

shielding

O. 05 --

O. O0

C,,

-0.05

-0.10

. °

° -

. # ° ,.,.....

.." i A Z_ _"

o

exp. pred. g,

- O

- 0 +I0°

- © 0°

"-] -I0" 0

........ 20°

-o.15 ] [ , I , I , I i I , I , I

-0.2 0.0 0.2 0.4 0.6 0.8 1.0 1.2

CL

Fig. 167 Comparison Between Prediction With Shielding and Experimentally Measured Moment as a Function of Lift on the Shuttle Orbiter at Hypersonic

Speed Including the Effect of Elevator Deflection.

Shuttle ; Cm vs. CL ; M = 5.0 ; 6e = +10 °, 0 °, -10 °, and -20 ° : with

shielding
0.001

0.000

-0.001

w

,,,, lJ,,, l,I,,i,,,, I,,,, I

-10 0 10 20 30 40

A.ng]eof _ack ~ degr_s

Fig. 168 Comparison Between Predicted and Experimentally Measured Side Force

Due to Aileron Deflection on the Shuttle Orbiter at Hypersonic Speed.

Shuttle ; CY_o vs. a ; M = 5.0

0.001 -

#s

0.000

-0.001

_ c_ © 0 _ °

,i],rT,z1(iI,, I,,,, I,,_, f

-10 0 10 20 30 40

Angle of a_ack ~ degrees

Fig. 169 Comparison Between Predicted and Experimentally Measured Yawing

YAoment Due to Aileron Deflection on the Shuttle Orbiter as Hypersonic

Speed.

Shuttle ; Cn_o vs. a ; M = 5.0
Cl_ a

.002

.O01

• 000 iLil,l,,,,I,,,,I,,,,I,,i,I

-10 0 10 20 .30 40

Angle o£ at,ta.r,.k.-.-degrees

Fig. 170 Comparison Between Predicted and Experimentally Measured Roiling

Moment Due to Aileron Deflection on the Shuttle Orbiter at Hypersonic

Speed.

Shuttle ; Cz_o vs. a ; M = 5.0

XlO -3

0.50

0.25

O. O0

0

0

0

0

0

i i i i J i _ i _ J l i i i I i , l i[_I i l'_'-_

-10 0 10 20 30 40

Angle of Ltta_k~ degrees

Fig. 171 Comparison Between Predicted and Experimentally Measured Side Force

Due to Rudder Deflection on the Shuttle Orbiter at Hypersonic Speed.

Shuttle ; Cy_o vs. a ; M = 5.0
CW,£ t

X10 -3

0.0

-0.2

-0.4

0

0

0 0

0 0

0

,,,, l,,,,1,1111,,,,],,I, I

-10 0 10 20 .30 40

Angle of a.t_Ic ~ degrees

Fig. 172 Comparison Between Predicted and Experimentally Measured Yawing

Moment Due to Rudder Deflection on the Shuttle Orbiter at Hypersonic

Speed.

Shuttle ; C,_, vs. c_ ; M = 5.0

•0002 [- 0

u 0

C_,. L 0 0

•0001

.0000 I I J I I

0

! I [ i ! I l I ! I !

-10 0 10 20 30 40

Ang}e of a:_ck ~ degrees

Fig. 173 Comparison Between Predicted and Experimentally Measured Rolling

Moment Due to Rudder Deflection on the Shuttle Orbiter at Hypersonic

Speed.

Shuttle ; CI_,. vs. cz ; M = 5.0
CL

1.5

1.0

0.5

0.0

-0.5

exp. pred. _,

<> _-I0°

0 0°

- i0•

A ........ 20°

- 10 0 10 20 30 40

._g_e of =Ltt.=ck -- degrees

Fig. 174 Comparison Between Predicted and Experimentally Measured Lift on the

Shuttle Orbiter at High Hypersonic Speed Including the Effect of Elevator

Deflection.

Shuttle ; CL vs. a ; M = 20.0 ; _ = +10 ° , 0 °, -10 °, and -20 °

CD

1.25

1.00

0.75

0.50

0.25

O. O0

i-

-10

exp. pred. 6=

+10 °

[] -10 ° /_/.. :

A ........ 20 ° " r : "_

,Y

f J.f •

r--o ?

0 10 20 30 40

Angle of _:ac]: ~ _,eg'ree.s

Fig. 175 Comparison Between Predicted and Experimentally Measured Drag on

the Shuttle Orbiter at High Hypersonic Speed Including the Effect of

Elevator Deflection.

Shuttle ; CD vs. a ; M = 20.0 ; be -- "10 ¢, 0 °, -10 °, and -20 °
CD

1.25 --

1.00

O. 75

0.50

0.25

exp. pred. &,

0 +]0 °

- 0 0° ./'

C3 -]0 ° ._'_a"

- Z_ ........ 20"

,._

o.oo - ' '_' I , I , I , I , !. , I

-0.2 0.0 0.2 0.4 0.6 0.8 1.0 1.2

CI.,

Fig. 176 Comparison Between Predicted and Experimentally Measured Drag as a

Function of Lift on the Shuttle Orbiter at High Hypersonic Speed Including the Effect of Elevator Deflection.

Shuttle ; CD vs. CL ; M = 20.0 ; 6e = +10 °, 0 °, -10 °, and -20 °

CIR_

0.05

O. O0

-0.05

-0.10 --

-0.15 i

-10

,., "-"':":'-...

exp. pred. 6,

<_ +:10 °

0 0°

D -10 °

z_ ......... 20°

%

%

\0

0 10 20 30 40

Angle of a_ack ~ _egrees

Fig. 177 Comparison Between Predicted and Experimentally Measured Moment

on the Shuttle Orbiter at High Hypersonic Speed Including the Effect of

Elevator Deflection.

Shuttle ; Cm vs. _ ; M = 20.0 ; _;_ = .10 ° .. 0 °,. -10 °, and -20 °
C_

O. 05

O. O0

-0.05

-0.10 t

-0.15

-0.2

..:....o.ool

• "" o o

-- ,'xp. pred. E, O" ",

<> +]0 ° ,

0 0 % °

[] .... ]0 • 0%,%,

A ........ 20 °

, I, I, I, I, I, I, I

0.0 0.2 0.4 0.6 0.8 1.0 1.2

Cz,

Fig. 178 Comparison Between Predicted and Experimentally Measured Moment

as a Function of Lift on the Shuttle Orbiter at High Hypersonic Speed

Including the Effect of Elevator Deflection.

Shuttle ; Cm vs. CL ; M = 20.0 ; 5e = +10 ° , 0°, -10°, and -20 °

L_C L

0.04

0.02

0.00

-0.02

-- exp. pred. 6_f

A ....... 11.5®

[] +102"

0 +22.5 °

j.

S

I

- _r-_---a---a..

0

0

0

.. A

" ° - - . . . . _

, ,, , I,,, , I,,,, I, ,, , I, ,, , ]

-10 0 10 20 30 40

Angle of attack ~ degrees

Fig. 179 Comparison Between Predicted and Experimentally Measured Lift on the

Shuttle Orbiter at High Hypersonic Speed Including the Effect of Body

Flap Deflection.

Shuttle ; ACL vs. a ; M = 20.0 ; 6bf = -11.5°, + 10.0°, and +22.5 °

i
Z_CD

O. 06

0.04

0.02

O. O0

-0.02

-10

exp. pred, 6br

A ....... ii.5°

D +102"

0 +22.._"

f

I

I

i

,r

: .......

O

.o"

D

" - . Ik

I_,,I,,,,I,,,,1,,,,1,,,,I

0 10 20 30 40

Angle ofaVt.a.ck.--.de_ee.s

Fig. 180 Comparison Between Predicted and Experimentally Measured Drag on

the Shuttle Orbiter at High Hypersonic Speed Including the Effect of

Body Flap Deflection.

Shuttle ; ACD vs. a ; M'= 20.0 ; 6br = -11.5 °, +10.0 °, and +22.5 °

_Cm

0.025

0.000

-0.025

-0. 050

-0.075

°

- .J_ rl

" O "-"" --.. D

".. <>

°

.

- • 2%,

exp. pred. 6_r

A ...... . _11.5° " - - .

D +i0.0o 0

0 +22.5 ° 0

,,,,l,,i, I, lill,,I,l,,, , I

-10 o lO 20 30 40

Angle of _ttar.k ~ degrees

Fig. 181 Comparison Between Predicted and Experimentally Measured Moment

on the Shuttle Orbiter at High Hypersonic Speed Including the Effect of

Body Flap Deflection.

Shuttle ; ACrn vs. a ; M = 20.0 ; 6bf = --11.5 °, +10.0 °, and +22.5 °
CL

1.0

0.5

0.0

-0.5

°'O

withou:_hie]ding . """'""

-10 0 10 20 30 40

Angle of attack ~ degrees

Fig. 182 The Effect of Shielding on the Predicted Lift on the Shuttle Vehicle at

High Hypersonic Speed.

Shuttle ; CL vs. a ; M = 20.0 ; with and without shielding

CD

I . 00

0.75

0.50

0.25

0. O0

--10

! P

witho:_ shielding /." '"

,,l,,,,I],lzl,lx!It,I II

0 10 20 30 40

Angle of _t'ack ~ degrees

Fig. 183 The Effect of Shielding on the Predicted Drag on the Shuttle Vehicle at

High Hypersonic Speed.

Shuttle ; CD vs. ct _ M = 20.0 ; with and without shieh: ng
CD

i .00

O. 75

O. 50

0.25

0.00

-0.2

withou_ shielding

...... with shiela_

I I , I i I I I i I , f

0.0 0.2 0.4 0.6 0.8 I .0

Fig. 184 The Effect of Shielding on the Predicted Drag as a Function of Lift on

the Shuttle Vehicle at High Hypersonic Speed.

Shuttle ; CD vs. CL ; M - 20.0 ; with and without shielding

OR

0.050

0.025

0.000

-0.025

-0.050

-0. 075

Lm

° - " ° ° ° . . °

"° °'.

...... v'i:h shkidi.._

,,,,I,,,, I,,,, I,,, ,I1.,l LI

-10 0 10 20 30 40

Angle of attack _ degrees

Fig. 185 The Effect of Shielding on the Predicted Moment on the Shuttle Orbiter

at High Hypersonic Speed.

Shuttle ; Cm vs. _ : M = 20.0 ; with and without shielding
!

C_

O. 050

0.025

0.000

-0.025

-0.050

I . ...........

O'

...... with shielding

-0.075 , I , f, , I , I , [ , I

-0.2 0.0 0.2 0.4 0.5 0.8 1.0

C/.

Fig. 186 The Effect of Shielding on the Predicted Moment as a Function of Lift

on the Shuttle Orbiter at High Hypersonic Speed.

Shuttle ; Cm vs. CL ; M = 20.0 ; with and without shielding

1.5

CL

1.0

0.5

0.0

-0.5

- 10 0 10 20 30 40

Angle of _t, taCk ~ degrees

Fig. 187 Comparison Between Prediction With Shielding and Experimentally Measured Lift on the Shuttle Orbiter at High Hypersonic Speed Including the

Effect of Elevator Deflection.

Shuttle ; CL vs. _ ; M = 9-0.0; de = "710 °, 0c: --10°; and -20=; with

shielding
CD

I.O0

0.75

0.50

0.25

exp. pred. 3,

0 +]0_ ,',_ 0 0° '

[] -i0°

A ........ 20° _L_"/

//

_/_ • S /w

0.00 , , , , I , , , , I ,...,, , I , , , , I , , , , I

-10 0 10 20 30 40

Angle of atta_ _ _egrees

Fig. 188 Comparison Between Prediction With Shielding and Experimentally Measured Drag on the Shuttle Orbiter at High Hypersonic Speed Including

the Effect of Elevator Deflection.

Shuttle ; CD vs. a ; M- 20.0; 6_ - +10 °, 0 °, -10 °, and -20°; with

shielding

CD

1. O0

0.75

0.50

0.25

- exp. pred. 6,

i O +1o" _>

o o[

- D .... 10° ./

- ,_- ........ 20 -_

" ,_

°

C.O0 _ J

-0.2 0.0 0.2 0.4 0.6 0.8 i.0

C,L

Fig. 189 Comparison Between Prediction With Shielding and Experimentally Measured Drag as a Function of Lift on the Shuttle Orbiter at High Hypersonic Speed Including the Effect of Elevator Deflection.

Shuttle ; C_ vs. CL ; M = 20.0: 6_ =-+-10 c,0 °, -10 °, and -20°; with

shielding
Cm

0.10

0.05

O. O0

-0.05

-0.10

u

exp. pred. 6,

-I0°

0 0° .........

D -lO _ ._:''"

.--S--..

0

0

i,,, , I, ,, , I, ,, , I, ,,, I, ,, ,$

-10 0 10 20 30 40

Angle of =track ~ deLn'ees

Fig. 190 Comparison Between Prediction With Shielding and Experimental]y Measured Moment on the Shuttle Orbiter at High Hypersonic Speed Including

the Effect of Elevator Deflection.

Shuttle ; Cm vs. a ; M = 20.0; 6e - +10 °, 0 °, -10 °, and -20°; with

shielding

C_fa

0.10

0.05

O. O0

-0.05

exp. pred. 6,

<> +I0"

i

" 0 0"

=.

[] -I0"

-- A ........ 20:i". _ ..-l-"

o--...

o o

O

-o.1o i , I . , I , l , l.. , I ,o I

-0.2 0.0 0.2 0.4 0.6 0.8 1.0

C,,..

Fig. 191 Comparison Between Prediction With Shielding and Experimentally Measured Moment as a Function of Lift on the Shuttle Orbiter at High Hypersonic Speed Including the Effect of Elevator Deflection.

Shuttle ; Cm vs. Cc ; M = 20.0; 5e = +10 _,0 °, -10 °,and -20°; with

shielding
C_,

0.001

0.000

-0.001

m

-10 0

--- without shielding

...... with shielding;

...... _0

0 0

,,,, [,,,, I ,,,i I,,,, I

10 20 30 40

Angle of _tu__k ~ de_eeJ

Fig. 192 Comparison Between Prediction With and Without Shielding and Experimentally Measured Side Force Due to Aileron Deflection on the Shuttle

Orbiter at High Hypersonic Speed.

Shuttle ; Cy_= vs. a ; M = 20.0 ; with and without shielding

0.001

O. 000

-0.001

-10

without shielding

...... with shielding

0 10 20 50 40

Angle of attack ~ degrees

Fig. 193 Comparison Between Prediction With and V,:ithout Shielding and Experimentally Measured Yawing Moment Due to Aileron Deflection on

the Shuttle Orbiter at High Hypersonic Speed.

Shuttle ; C,_ vs. a ; M = 20.0 ; with and without shielding
,J

Ch.

O. 005

0.000

L

-.! .I ! # _ ! ! !

without shie|din_

...... wi_h shie]din_ . - "

o

-0.005 Ill,, I ,, f , I,,, , }

-10 0 10 20 30 40

A_g}e of at_ac_ ~ degrees

Fig, 194 Comparison Between Prediction With and Without Shielding and Experimentally Measured Rolling Moment Due to Aileron Deflection on the

Shuttle Orbiter at High Hypersonic Speed.

Shuttle ; Cz_, vs. a ; M = 20.0 ; with and without shielding

C_.

0.001

0,000

-0.001

without shielding

....... with shielding

.

o

0 0 .O. 0"" o

-10 0 10 20 30 40

Angle of attack ~ degrees

Fig. 195 Comparison Between Prediction With and Without Shielding and Experimentally Measured Side Force Due to Rudder Deflection on the Shuttle

Orbiter at High Hypersonic Speed.

Shuttle ; Cy_. vs. a _ M = 20.0 _ with and without shielding
CRAw

0.001

0.000

-0.001

-- without shielding

...... whb shielding

6"6"'0"0--- o" o o

ii,,ll,,,l,,,,l,,,,l,,,,l

- 10 0 10 20 30 40

Angle of attack -- degrees

Fig. 196 Comparison Between Prediction With and Without Shielding and Experimentally Measured Yawing Moment Due to Rudder Deflection on

the Shuttle Orbiter at High Hypersonic Speed.

Shuttle ; C_, vs. a ; M = 20.0 ; with and without shielding

o.o01 _- ,"'"

I O. O.. O" " "n ,-, 0. 000 / _ 0

-- withou_ shielding

...... with shielding

-I0 0 10 20 30 40

Angle of _tt_.k ~ degrees

Fig. 197 Comparison Between Prediction With and Without Shielding and Experimentally Measured Rolling Moment Due to Rudder Deflection on the

Shuttle Orbiter at High Hypersonic Speed.

Shuttle ; Cl_,. vs. _ ; M = 20.0 ; with and without shielding
O. 0O2 -

C_

O. 000

-0. 002

-0. 004

. _ .-s _- -i r =--° °--.

i exp. pred. =

m

I

0

10 °

i,,I, ,,,I,,111 i , t ,

5 10 15

M I_..._numb_

I

2O

Fig. 198 Comparison Between Predicted and Experimentally Measured Side Force

Due to Aileron Deflection on the Shuttle Orbiter as a Function of Math

Number at Two Angles of Attack.

Shuttle ; Cry= vs. M ; a = 0 ° and 10 °

XlO -3

1.0

C,,. O. 5

0.0

-0.5

exp. pred. a

0 ....... 0 °

A I0u

g"-:_-.- : -.-:--- ":' " '" "-: :2

,,zilzl l,l,,,,J,IJ]l

0 5 10 15 2O

M,,eh number

Fig. 199 Comparison Between Predicted and Experimentally Measured Yawing

Moment Due to Aileron Deflection on the Shuttle Orbiter as a Function

of Mach Number at Two Angles of Attack.

Shuttle ; Cn,_ vs. M ; a = 0 ° and 10 °
CI&,

•0050

.0025

•0000

m

'}

,%

-0 '_

,%

! I I

0

exp. pred. c:

0 ....... 0 °

,., ID °

, _:,".?r""T-I:,re.r -T-'I:,r_:_

5 10 15 20

M_..hnumber

Fig. 200 Comparison Between Predicted and Experimentally Measured Rolling

Moment Due to Aileron Deflection on the Shuttle Orbiter as a Function

of Mach Number at Two Angles of Attack•

Shuttle ; Ct_= vs. M ; a = 0 ° and 10 °

cy_.

.0050

.0025

.0000

i

-0

0

I I I I -I.-I- I .t.j.L J_ I _I I L_I

5 10 15

Ma:.h number

2O

Fig. 201 Comparison Between Predicted and Experimentally Measured Side Force

Due to Rudder Deflection on the Shuttle Orbiter as a Function of Mach

Number.

Shuttle ; Cy_, vs. h<[ ; a = 0=
0.000

-0. 002

-0. 004

-C_ * -

D '

A.,,.,

,O" ...................... D

,,,, I,,,, I, ,,, 1_,, , , I

0 5 10 15 20

M_.h number

Fig. 202 Comparison Between Predicted and Experimentally Measured Yawing

Moment Due to Rudder Deflection on the Shuttle Orbiter as a Function

of Mach Number.

Shuttle ; C,_, vs. M ; a = 0°

• 002 --

.001

• 000

%

! I , 1 , J.L .1. :_L J_L_I j I a , i •

0 5 10 15 20

h _Lr./n nuzmbe:

Fig. 203 Comparison Between Predicted and Experimentally Measured Rolling

Moment Due to Rudder Deflection on the Shuttle Orbiter as a Function

of Mach ,Number.

Shuttle ; Cl_,, vs. M • a -- 0 °
AppendLx: Bibliography of Experimental Force and Moment Data

For Hypersonic Vehicle Configurations

1. Beeler, D.E., "The X-15 Research Program," AGARD Report 289, October 1960.

Brief description of the X-15 research program. Comparison of wind tunnel tests

and flight measured values of stability derivatives at Mach numbers from 0.0 to

8.0. Variation of control effectiveness with Mach number including horizontal tail,

vertical tail and ailerons.

2. Bernot, P.T., "Effect of Modifications on Aerodynamic Characteristics of a SingleStage-to Orbit Vehicle at Mach 5.9," NASA TM 84565, January 1983.

The model was based on control-configured stability concepts. Results are presented for elevons, body flap, and wing tip fin controllers. Model similar to that

in NASA TM X-3550 (item 3).

3. Bernot, P.T., "Aerodynamic Characteristics of Two Single-Stage-to-Orbit Vehicles at

Mach 20.3," NASA TM X-3550, August 1977.

Control deflection data are for elevons and a body flap. Most results are for high

angles of attack (between 16 ° and 50°).

4. Boissean,Peter C., "Investigation of the Low-speed Stability and Control Characteristics of a 1/7-Scale Model of the North American X-15 Airplane," NACA RM L57D09,

1957.

Early X-15 data from free flying model tests. Purpose was to evaluate the use of

the horizontal tail for roll control. Control deflection results are presented for the

wing trailing edge flap, all moving vertical tail, and symmetrical and differential

horizontal tail deflections.

5. Boyden, R.P. and Freeman, D.C. Jr., "Subsonic and Transonic Dynamic Stability

Characteristics of a Space Shuttle Orbiter," NASA TN D-8042, November 1975.

Dynamic and static stability are investigated. Dynamic results are presented to

show the effect of rudder flare in combination with body flap deflection. The static

lateral stability data show the effect of the vertical tail, combination body flap

and rudder flare, and body flap alone.

6. Brooks, C.W. Jr. and Cone, C.D. Jr., "Hypersonic Aerodynamic Characteristics of

Aircraft Configurations with Canard Controls," NASA TN D-3374, April 1966.

The investigation was done on a wing-body configuration with a 70 ° swept delta

wing at a Mach number of 10.03. Four different canards were each tested on various configurations. Results include canard effectiveness on longitudinal, lateral,

and directional characteristics.

7. Brooks, C.W. Jr., "Interference Effects of Canard Controls on the Longitudinal Aerodynamic Characteristics of a Winged Body at Mach 10," NASA TN D-4436, April 1968.

Effect of canard interference is studied by comparing body alone data with canard deflection data which appeared in NASA TN D-3374 (item 6) and TN D3728 (item 44).
w

8. Clark, L.D., "Hypersonic Aerodynamic Characteristics of an All-Body Research Aircraft Configuration," NASA TN D-7358, December 1973.

Experiment was done at Mach 6 on a lifting body configuration and compared

with theoretical models. The horizontal wing-tip-type control surfaces were adjustable in 5 ° increments from +15 ° to -30 °. HABP was used for theoretical

predictions. The tangent-cone method gave the best agreement at control settings between +5 ° and -5 ° and at positive lift coefficients except for directional characteristics. None of the methods predicted characteristics well at

negative lift coefficients and large control deflections.

9. Clark, L.E. and Richie, C.B., "Aerodynamic Characteristics at Mach 6 of a Hypersonic Research Airplane Concept Having a 70 ° Swept Delta Wing," NASA TM X3475, May 1977.

The study was a configuration build up and includes effect of elevon defleciton

on trim characteristics. Elevon data are given for deflections of 10 ° to -20 ° in

5 ° increments for seven different configurations. Speed brake deflection data are

also included.

10. Covell, P.F., Wood, R.M., Bauer, S.X., and Malker, I.J., "Configuration Trade and

Code Validation Study on a Conical Hypersonic Vehicle, _ AIAA Paper 88-4505,

September 1988.

Test Mach numbers were between 2.5 and 4.5. Effect of canard shape, vertical

tail shape, wing location, and wing incidence on aerodynamic characteristics included. A comparison is also made between the experimental results and three

theoretical analysis programs: HABP, LT (Linear Theory), and SIMP (Supersonic Implicit Marching).

11. Decker, J.P. and Spencer, B. Jr., "Low-Subsonic Aerodynamic Characteristics

of a Model of a Fixed-Wing Space Shuttle Concept at Angles of Attack to 76°,"

NASA TM X-1996, April 1970.

These tests were done on an early shuttle concept at a Mach number of 0.25.

Longitudinal stability and control are provided by a horizontal tail with an elevator. Elevator deflections of 20 ° to -20 ° were effective at low angles of attack

where flow separation is not a major problem.

12. Dillon, J.L. and Creel, T.R. Jr., "Aerodynamic Characteristics at Mach Number

0.2 of a Wing Body Concept for a Hypersonic Research Airplane," NASA TP 1189,

1978.

The experiment consisted of configuration buildup from the basic body by

adding a wing, center vertical tail and scramjet engines. The test angle of attack range was approximately -5 ° to 30 ° at constant angles of sideslip of 0 °

and 4 °. The elevons were deflected from 5 ° to -15 °. Roll and yaw control were

investigated. Also includes rudder deflection data.
w

w

13. Dillon, J.L. and Pittman, J.L., "Aerodynamic Characteristics at Mach Numbers

from 0.33 to 1.20 of a Wing-Body Design Concept for a Hypersonic Research Airplane," NASA TP 1044, 1977.

The tests were done at seven different transonic Mach numbers. Control deflection data includes: symmetrical elevon deflections of 0 °, - 10 °, and -20¢;

differentially deflected elevons at +20°; and rudder deflections of 0 ° and 15.6 °.

14. Dillon, J.L. and Pittman, J.L., "Aerodynamic Characteristics at Mach 6 of a WingBody Concept for a Hypersonic Research Airplane," NASA TP 1249, August 1978.

Similar configuration build-up of model in TP 1044 (item 13) and

TP 1189 (item 12) at Mach 6. The elevons were deflected from 10 ° to -15 °

for pitch control and yaw and roll control were also investigated. HABP was

used and gave good predictions for the longitudinal but not for the lateraldirectional aerodynamic characteristics.

15. Ellison, J.C., "Investigation of the Aerodynamic Characteristics of a Hypersonic

Transport Model at Mach Numbers to 6," NASA TN D-6191, April 1971.

Tests were done at Mach numbers from 0.36 to 6.0. Results for elevon deflections from 5 ° to -20 ° are included for all Mach numbers. The configuration

also had strakes which helped provide positive CM.

16. Fetterman, D.E. Jr., Penland, J.A., "Static Longintudinal, Directional, and Lateral

Stability and Control Data from an Investigation at a Mach Number of 6.83 of Two

Developmental X-15 Airplane Configurations," NASA TM X-209, March 1960.

Directional control data were obtained by testing vertical tail deflections of 0 °

and -5 °, and lateral control data were obtained by testing differential horizontal tail deflections of 0 °, -10 °, and -20 °. Results for speed brake deflections of

20 ° are also included.

17. Freeman, D.C. Jr., "Dynamic Stability Derivatives of Space Shuttle Orbiter Obtained from Wind-Tunnel and Approach and Landing Flight Tests,"

NASA TP 1634, April 1980.

Wind tunnel and flight test data were compared with ADDB values at subsonic

Mach numbers for the parameters of pitch, yaw and roll damping, as well as

the yawing moment due to rolling velocity and the rolling moment due to yawing velocity.

18. Freeman, D.C. and Boyden, R.P., "Supersonic Dynamic Stability Characteristics of

a Space Shuttle Orbiter," NASA TN D-8043, January 1976.

Similar results to NASA TN D-8042 (item 5) except for different Mach numbers. Elevon data are also included.

19. Freeman, D.C. and Fournier, R.H., "Static Aerodynamic Characteristics of a SingleStage-to-Orbit Vehicle With Low Planform Loading at Mach Numbers from 0.3 to

4.63," NASA TM 74056, November 1977.

Tests were run at eight different Mach numbers. At a Mach number of 0.9,

elevon deflections of 10 ° became completely ineffective at angles of attack above

6 °. The resulting nonlinearity in CM_, was also seen in the Space Shuttle.
w

20. Freeman, D.C. and Fournier, R.H., "Static Aerodynamic Characteristics of

a Winged Single-Stage-to-Orbit Vehicle at Mach Numbers from 0.3 to 4.63,"

NASA TP 1233, August 1978.

Tests were done to determine the static longitudinal stability and trim, the

static lateral-directional stability, and the aileron control effectiveness. Elevons

were deflected from 0 ° to -20 ° for all eight Mach numbers tested.

21. Freeman, D.C. and Jones, R.S., "Low -Speed Static Stability and Control Characteristics of Two Small-Scale, Hypersonic Cruise Configurations," NASA TM X-2021,

June 1970.

The first model was a distinct wing-body with a conventional rudder for directional control and differential deflections of the all-movable horizontal tail for

roll control. The second model was a blended wing-body with elevons for both

pitch and roll control and a center vertical rudder for directional control.

22. Freeman, D.C. and Spencer, B. Jr., "Comparison of Space Shuttle Orbiter LowSpeed Static Stability and Control Derivatives Obtained from Wind-Tunnel and

Approach and Landing Flight Tests," NASA TP 1779, December 1980.

The longitudinal stability, elevon effectiveness, lateral directional stability and

aileron effectiveness derivatives were compared from wind tunnel tests, approach and landing flight tests and ADDB values. Body flap and speed brake

deflections are included.

23. "Hypersonic Aerodynamic Characteristics of Two Delta-Wing X-15 Airplane Configurations," NASA TN D-5498, October 1969.

The effects of wing geometry and longitudinal position, wing fins, nose cant,

strakes, and speed brakes were looked at for elevon deflections to -45 °. The

experimental aerodynamic characteristics were compared with the analytical

results from HABP. At the time of the investigation, HABP was very new and

the results were not very good.

24. Kelly, M.W., "Wind-Tunnel Investigation of the Low-Speed Aerodynamic Characeristics of a Hypersonic Glider Configuration," NACA RM A58F03, September 1958.

The tests were done to investigate the adequacy of the low speed stability and

control characteristics for landing. Trailing-edge flaps at the wing tips supplied

both yaw and roll control. The effect of wing tip droop on lateral and directional stabiltiy is also reported.

25. McCandless, R.S. and Cruz, C.I., "Hypersonic Characteristics of an Advanced

Aerospace Plane," AIAA Paper 85-0346, January 1985.

Tests were run at Mach numbers of 6, 10, and 20. Results include elevator,

elevon and rudder deflection data. The experimental data were then compared

with APAS II predicted values.
26. McCandless,R.S., "Hypersonic Characteristics of an Advanced Aerospace Plane at

Mach 20,3," NASA TM 86435.

Aerodynamic control effectiveness was determined by deflecting the elevators,

the elevons, and the rudder. Tests were run at a Mach number of 20.3 at various Reynolds numbers.

27. McKinney, R.L. and Lancaster, J.A., "Investigation of the Aerodynamic Characteristics of a 0.02-Scale Model of the X-15 Airplane at Mach Numbers of 2.96, 3.96,

and 4.65 at High Angles of Attack," NASA TM X-820, June 1963.

Supersonic tests on the final X-15 configuration. Results include deflections of

the horizontal tail, asymmetric deflections of the upper and lower verticals, and

deflections of upper and lower speed brakes.

28. Mellinger, G.R., "Design and Operation of the X-15," Shell Aviation News,

April 1961, pp 14-21.

Includes a description of the X-15 design. The article discusses design decisions

such the need for the wedge airfoil for the upper and lower vertical tails.

29. Moore, M.E. and Williams, J.E., "Aerodynamic Prediction Rationale for Analyses

of Hypersonic Configurations," AIAA Paper 89-0525, January 1989.

A method selection rationale was developed for S/HABP. They suggest braking

the configuration into three basic parts: nose, body, and aerodynamic surfaces.

Analyses were done on the Space Shuttle, the FDL-7, and the X-24C-10D. No

comparisons for control deflections.

30. Nelms, W.P. and Ames, J.A., "Longitudinal Aerodynamic Characteristics of Three

Representative Hypersonic Cruise Configurations at Mach Numbers from 0.65 to

10.70," NASA TM X-2113, October 1970.

Two configurations were discrete wing-body concepts and the third was a

blended wing-body design. Effects of varying angle of attack, Mach number,

and configuration build-up were considered. There are no control deflection

data in this report.

31. Nelms, W.P. and Thomas, C.L., "Aerodynamic Characteristics of an All-Body Hypersonic Aircraft Configuration at Mach Numbers from 0.65 to 10.6," NASA TN D6577, November 1971.

The effectiveness of horizontal tail, vertical tail and canard stabilizing and control surfaces were investigated. The horizontal tail was deflected both symmetrically and differentially. The rudder was deflected asymmetrically and flared as

a speed brake.

32. Osbourne, R.S., "Aerodynamic Characteristics of a 0.0667 Scale Model of the N.A.

X-15 Research Airplane at Transonic Speeds," NASA TM X-24, 1959.

Tests were run at eight Mach numbers between 0.6 and 21.43. The tests were

not run on the final X-15 configuration (does not include vertical wedge airfoil).

Results are presented for symmetrical and differential deflections of the horizontal tail.
33. Penland, J.A., "Low-SpeedAerodynamic Characteristics for a Hypersonic ResearchAirplane Concept Having a 70° SweptDelta Wing," NASA TM X-71974,

August 1974.

Testswere conductedat a Mach number of 0.06 on a model like that in NASA

reports: TP 1252(item 38), TP 1552(item 39), TM X-3475 (item 9), and

TN D-8065(item 37). Eight model configurations weretested with various

elevonand aileron deflections.

34. Penland, J.A. and Fetterman, D.E. Jr., "Static Longitudinal, Directional, and Lateral Stability and Control Data at a Mach Number of 6.83 of the Final Configuration of the X-15 ResearchAirplane," NASA TM X-236, April 1960,

Data arepresentedin comparisonplots to showthe effectsof component breakdown and control deflection. Control surfacesinclude: vertical tail, horizontal

tail (symmetrical and differential deflections),speedbrakes.The configuration

geometry is well documentedin the report (0.02scalemodel).

35. Penland, J.A. and Creel, T.R. Jr., "Low-SpeedAerodynamic Characteristicsof

a Lifting-Body Hypersonic Research Aircraft Configuration," NASA TN D-7851,

February 1975.

Configuration is similar to that of NASA TN D-7358. The model was tested

with two sets of horizontal and vertical tip controls, a center vertical tail, and

two sets of canard controls.

36. Penland, J.A., Creel, T.R. Jr., and Howard, "Experimental Low-Speed and Calculated High-Speed Aerodynamic Characteristics of a Hypersonic Research Airplane

Concept Having a 65 ° Swept Delta Wing," NASA TN D-7633, August 1974.

Experimental low speed tests were done to determine lift and stability during

landing. Calculated results using HABP are presented for Mach numbers from

3 to 12. Results are given for elevon, aileron, and wing tip rudder deflections.

37. Penland, J.A., Fournier, R.H., and Marcum, D.C. Jr., "Aerodynamic Characteristics

of a Hypersonic Reseach Airplane Concept Having a 70 ° Swept Double-Delta Wing

at Mach Numbers from 1.50 to 2.86," NASA TN D-8065, December 1975.

A configuration build-up was done as well as the effect of elevon deflections.

Data for elevon deflections of 0 °, -10 °, and -20 ° are presented at four supersonic Mach numbers. Aileron effectiveness data are also presented.

38. Penland, J.A., Creel, T.R. Jr., and Dillon, J.L., "Aerodynamic Characteristics of a

Hypersonic Research Airplane Concept Having a 70 ° Swept Double-Delta Wing at

Mach Number 0.2," NASA TP 1252, September 1978.

Tests were done at a Mach number of 0.2 for various Reynolds numbers. The

elevons were deflected from 0 ° to -20 ° in 5 ° increments. Roll control was also

investigated.

=
39. Penland, J.A., Hallissy, and Dillon, J.L., "Aerodynamic Characteristics of a Hypersonic Research Airplane Concept Having a 70 ° Swept Double-Delta Wing

at Mach Numbers from 0.80 to 1.20, With Summary of Data from 0.20 to 6.0,"

NASA TP 1552, December 1979.

Wind tunnel data of static longitudinal, lateral, and directional stabilty characteristics of a hypersonic research airplane, for angles of attack from -4 ° to 23 °,

and at angles of sideslip of 0 ° and 5 °. The configuration variables included wing

planform, tip fins, and the center vertical tail. The second area is a summary of

the variations of the more important aerodynamic parameters with M = 0.2 to

6.0. Elevon deflections are included.

40. Pittman, J.L and Riebe, G.D., "Experimental and Theoretical Aerodynamic Characteristics of Two Hypersonic Cruise Aircraft Concepts at Mach Numbers of 2.96,

3.96, and 4.63," NASA TP 1767, December 1980.

Comparison of wind tunnel tests with results from various theoretical methods

includeng HABP. The control deflections studied were for horizontal and vertical tails.

41. Powell, R.W. and Freeman, D.C., Jr., "Application of a Tip-Fin Controller to the

Shuttle Orbiter for Improved Yaw Control," Journal of Guidance and Control,

AIAA Paper 81-0074R, 1982.

Looks at possibility of implementing tip-fins on the Shuttle Orbiter for imnproved yaw control. Results compared effectiveness of speed brakes on the two

configurations.

42. Powell, R.W. and Freeman, D.C. Jr., "Aerodynamic Control of the Space Shuttle Orbiter with Tip-Fin Controllers, _ Journal o] Spacecraft, AIAA Paper 84-0488,

September-October 1985.

Results show that the orbiter with tip-fin controllers can successfully perform

the required maneuvers during entry. They do however exhibit less control authority in some flight regimes than the current configuration.

43. Putman, L.E. and Trescot, C.D. Jr., "Hypersonic Aerodynamic Characteristics of

Plain and Ported Elevon Controls on a 75 ° Swept Modified Delta-Wing Configuration," NASA TM X-987, July 1964.

The tests were run at a Mach numbr of 10.03 and at various Reynolds numbers. The relative effectiveness of plain and ported elevons and the effect of

wing position on elevon effectiveness were reported. Newtonian impact theory

was used to predict the control characteristics but was not adequate for this

configuration.

44. Putnam, L.E. and Brooks, C.W, Jr., "Hypersonic Aerodynamic Characteristics

of Wing-Body Configurations with Canard Controls," NASA TN D-3728, December 1966.

Similar to work in NASA TN D-3374 (item 6) but with different canards.
45. Rainey, R.W., Fetterman,D,E. Jr., and Smith, R., "Summary of the Static Stability and Control Results of a Hypersonic Glider Investigation," NASA TM X-277,

May 1960.

Test Mach numbers of 6.7 to 18.4. Wing trailing edge flaps provided longitudinal and lateral control, and the wing-tip fins with rudders provided directinal

stability and control. Other control surfaces tested include delta tripanel tip

controls, pyramidal tip controls, and conical tip controls.

46. Romere, P.D. and Young, J,C., "Space Shuttle Entry Aerodynamic Comparisons of

Flight 2 with Preflight Predictions," AIAA Paper 82-0565, March 1982.

Control deflection data given only for the speed brakes.

47. Ross, A.J. and Thomas, H.H.B.M, "A Survey of Experimental Data on the Aerodynamics of Controls, in the Light of Future Needs," Aerodynamics of Controls -

Paper P, AGARD CP 262.

This paper is a good overview of both conventional and unconventional motivators. A bibliography of control data for pitch, roll, yaw, and lift motivators is

included. They are further broken down into different speed regimes.

48. Shuttle Performance: Lessons Learned, "Space Shuttle Entry Longitudinal Aerodynamic Comparisons of Flights 1-4 with Prefight Predictions," NASA CP 2283,

March 8-10, 1983.

Analysis results of the STS 2 and 4 maneuvers during entry indicate that the

hypersonic trim discrepancy is due to an error in the prediciton of the basic

vehicle pitching moment and not an error in prediction of the elevon and body

flap effectiveness. Speed brake data are included.

49. Small W.J., Kirkham, F.S., and Fetterman, D.E., "Aerodynamic Characteristics of a Hypersonic Transport Configuration at Mach 6.86," NASA TN D-5885,

June 1970.

Configuration was a low-wing, distinct wing-body with a vertical tail. Elevon

deflection data are included. The analytical prediction methods used were

found to be inadequate. At a = 6.85 ° the vertical tail became ineffective due

to interference and shielding effects.

50. Spearman, M.L. and Driver, C., "Longitudinal and Lateral Stability and Control

Characteristics at a Mach Number of 2.01 of a 60 ° Delta-Wing Airplane Configuration Equipped with a Canard Control and with Wing Trailing-Edge Flap Controls,"

NACA RM L58A20, March 1958.

51. Spencer, B. Jr., "Effects of Stabilizer Configuration on Transonic Aerodynamic

Characteristics of a Variable-Geometry High-Hypersonic-Performance Spacecraft,"

NASA TM X-1865, September 1969.

The configuration has an adjustable wing which is stowed at high speeds and is

deployed at low speeds. Elevon controls were deflected 0 °, -10 °, and -20 ° for

pitch control and differentially for roll control. The results are given for various

horizontal stabilizer configurations.
w

52. Spencer, B. Jr., Henry, B.Z. Jr., and Putnam, L.E., "The Transonic Longitudinal

and Lateral Aerodynamic Characteristics of a Low-Fineness-Ratio Elliptic Hypersonic Configuration Employing Variable-Sweep Wing Panels for Improving Subsonic

Lift and Performance," NASA TM X-768, March 1963.

Results are given for body-base flap deflections at various transonic Mach numbers. The wing leading edge sweep angle was also varied with body flap deflections.

53. Suit, W.T. and Schiess, J.R., "Lateral and Longitudinal Stability and Control Parameters for the Space Shuttle Discovery as Determined from Flight test Data,"

NASA TM 100555, February 1988.

Comparison of flight test data with predictions. Results include: rolling and

yawing moments due to aileron deflection; rolling and yawing moments due to

rudder deflection; and pitching moment due to elevon deflection. These results

are given versus Mach number from 25.0 to 0.

54. Syvertson, C.A., Gloria, H.R., and Sarabia, M.F., "Aerodynamic Performance and

Static Stability and Control for Flat-Top Hypersonic Gliders at Mach Numbers

from 0.6 to 18," NACA RM A58G17, September 1958.

The model was a 77.4 ° swept arrow wing with deflected wing tips, a retractable

ventral fin, plain trailing edge flaps, a rudder on the ventral fin, and body flaps

that could also be used for speed brakes. All of the test results are presented in

tabular form.

55. Trescot, C.D. and Spencer, B. Jr., "Hypersonic Aerodynamic Characteristics of a

Lifting Reentry Vehicle Model with Four Types of Longitudinal Control Surfaces,"

NASA TM X-1173, November 1965.

Tests were at a Mach number of 10.03. The baseline configuration is that of

TM X-768 (item 52) with the wing in the 75 ° leading edge sweep position. The

four types of controls tested were aft-mounted fins, a canard, a chin flap, and

two body trailing edge flaps (one each on upper and lower surfaces).

56. Underwood, J.M. and Cooke, D.R., "A Preliminary Correlation of the Orbiter Stability and Control Aerodynamics from the First Two Space Shuttle Flights (STS 1

and 2) with Preflight Predictions," AIAA Paper 82-0564, March 1982.

Data from STS-1 and STS-2 are compared with Shuttle Data Book values. Results are presented for aileron and rudder effectiveness.

57. Walker, H.J. and Wolowicz, C.H., "Stability and Control Derivative Characteristics

of the X-15 Airplane," NASA TM X-714, March 1962.

Results compare actual flight test data with wind tunnel test predictions. Control deflections include horizontal tail, upper vertical tail rudder, and speed

brakes.

58. Yancy, R.B., Rediess, H.A., and Robinson, G.H., "Aerodynamic Derivative Characteristics of the X-15 Research Airplane as Determined from Flight Tests for Mach

Numbers from 0.6 to 3.4," NASA TN D-1060, January 1962.

Flight test results are derived from pulse, pull-up and sideslip maneuvers and

are compared to wind tunnel results for corresponding conditions. Results are

given for aileron, vertical tail and horizontal stabilizer deflections.