NASA Technical Memorandum 4756

Evaluation of the Low-Speed Stability and

Control Characteristics of a Mach 5.5 Waverider

Concept

David E. Hahne

Langley Research Center • Hampton, Virginia

National Aeronautics and Space Administration

Langley Research Center • Hampton, Virginia 23681-0001

May 1997
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Summary

Static force and moment tests of a 0.062-scale model

of a hypersonic vehicle study concept known as the

LoFLYTE rM1 configuration were conducted in the Langley 12-Foot Low-Speed Tunnel. These tests looked primarily at the low-speed static stability and control

characteristics of this configuration. Data were obtained

over an angle-of-attack range of-5 ° to 22 ° at sideslip

angles that ranged between -10 ° and 10°. Supporting

flow visualization tests were also conducted.

Two primary concerns before the test were whether

the configuration had adequate pitch control and adequate directional stability for low-speed operations. The

tiperons were sized to provide enough pitch control to

trim the vehicle to ot = 16 ° with no more than 10 ° of surface deflection; however, there still was some concern

about achieving this capability. The data obtained in this

test showed that 10 ° of tiperon deflection was nearly sufficient to trim the configuration to the desired angle of

attack. Because of the pitching-moment characteristics of

the LoFLYTE TM configuration, there is a reasonably high

level of unpowered trimmed lift at nominal takeoff and

approach to landing angles of attack that should allow for

acceptable takeoff and landing speeds for this vehicle.

Initial evaluation of the directional stability characteristics of this configuration showed a significant instability between tx = 10° and about t_ = 18°. This test

determined that the cause of this instability was the interaction of the wing leading-edge vortex with the vertical

tails. Moving the vertical tails either inboard or outboard

from the baseline location eliminated this unfavorable

interaction. The inboard location, however, is more

desirable because it also places the tails in a position

where they do not interfere with the tiperons. This location allows for the lower portion of the rudder to be

extended (the baseline rudders have a cutout to allow a

full 30 ° of tiperon deflection) and for the vertical tails to

be moved farther aft, thus resulting in an additional

increase in rudder power.

Introduction

The quest for faster and faster airplanes has been a

driving force in aeronautics since the Wright brothers'

first flight at the turn of the century. Man has always recognized the military and commercial benefits of decreasing the time it takes to travel between two points, and air

travel has always offered the most potential for improvements in the time-distance equation. Efforts to develop a

single-stage-to-orbit vehicle during the 1980's rekindled

1Trademark of Accurate Automation Corporation, Chattanooga,

Tennessee.

investigations into airplanes that could travel at hypersonic speeds (Mach > 4). While the X-30 program was

expected to leapfrog from vehicles capable of traveling

two or three times the speed of sound to velocities in

excess of Mach 15, it was understood that the technology

for developing a more modest hypersonic vehicle that

would travel at five to six times the speed of sound was

imbedded in the X-30 program. Advances made in hightemperature materials, lightweight structures, and high

thrust-to-weight ratio propulsion systems during the

X-30 program have made modest hypersonic vehicles,

such as the Mach 5.5 waverider of this study, not only

feasible but practical (ref. 1).

Waverider concepts have long been known to be

very efficient cruising vehicles in terms of lift-to-drag

ratio at their design Mach number (refs. 2 through 4).

While these configurations are well suited for operation

at design conditions, flying characteristics for off-design

conditions, particularly at low speed, can become undesirable. The waverider-derived configuration of the

present study is optimized for a Mach number of 5.5

(ref. 1), and as with all configurations designed for

hypersonic flight, it incorporates many highly swept

surfaces. Both the wing and the vertical tail of this configuration are swept to 75 ° . The vortical flows that dominate the aerodynamics of such highly swept surfaces can

result in complex and quickly changing flying characteristics at low-speed conditions. The present investigation

was intended to document the low-speed aerodynamic

characteristics of this configuration, with emphasis on

static stability and control, and to understand the

complex vortical interactions that dominate these

characteristics.

Symbols

All data initially were obtained in the body-axis system. Longitudinal forces and moments are presented in

the stability-axis system, and lateral-directional forces

and moments are presented in the body-axis system. All

aerodynamic moments were computed about a center-ofgravity location of 58 percent of the body length measured from the nose of the fuselage.

b wing span, ft

Lift

CL lift coefficient, ?/S

Ct rolling-moment coefficient, Rolling moment

_lSb

pitching-moment coefficient, Pitching moment

_ISI

yawing-moment coefficient, Yawing moment

glSb
Cy

LID

l

P

Ps

P

4

r

S

V

W

X,Y,Z

AQ

acm

Acn

ACt

_r

Side force

side-force coefficient,

#s

body axes moments of inertia, slugs-ft 2

lift-to-drag ratio

body length, ft

body axes roll rate, rad/sec

stability axes roll rate, rad/sec

roll-rate acceleration, rad/sec 2

pitch-rate acceleration, rad/sec 2

free-stream dynamic pressure, lb/ft 2

body axes yaw rate, rad/sec

yaw rate acceleration, rad/sec 2

wing-body planform area, ft 2

free-stream velocity, knots

vehicle weight, lb

body axes

angle of attack, deg

angle of sideslip, deg

incremental rolling-moment coefficient

incremental pitching-moment coefficient

incremental yawing-moment coefficient

incremental side-force coefficient

tiperon deflection, positive trailing edge down,

deg

rudder deflection, positive trailing edge left,

deg

Stability derivatives:

OCt _ OCn OCr

= -ff -

OC t OC n

Ct_" = O5 r Cns r =

OC t Of n

Ctst = -_t Cnst - O_t

Subscripts:

app approach

o value at zero angle of atack

TO takeoff

2

tr trimmed value (C m = O)

xw crosswind

Abbreviations:

BW body with blended wing

L left

N engine nacelle

R right

V vertical tails

WL waterline, in.

Model

The configuration tested is a 0.062-scale model of a

hypersonic vehicle study concept called LoFLYTE TM

(fig. 1). The configuration consists of blended wingbody, twin wing-mounted vertical tails, and an engine

nacelle package located on the underside of the body.

The model, 8.34 ft long with a 5.19-ft wing span, has an

initial leading-edge sweep of 75.13 °, which changes to

65 ° along the leading edge of the tiperons. The tiperons,

deflectable wingtips that also incorporate a more conventional trailing-edge surface, are used for both pitch and

roll control. Each vertical tail has a leading-edge sweep

of 75 ° and a full-height rudder with an aft-swept hing line

of 26.6 °. The baseline rudders have a cutout, or notch, at

the bottom to allow a full 30 ° of tiperon deflection. In

addition to the baseline position (fig. 1(a)), four alternate

vertical tail locations were tested: (1) 8 in. outboard of

the baseline location, (2) 4 in. inboard of the baseline

location, (3) a single centerline vertical tail, and (4) 4 in.

inboard and 4 in. aft of the baseline location. For the

fourth alternate vertical tail location, the notch in the rudder was filled. The engine nacelle was a simple flowthrough shell attached to the underside of the body. Geometric details of the model are given in table I.

Tests

A photograph of the model and test setup is shown in

figure 2. Tests were conducted in the Langley 12-Foot

Low-Speed Tunnel at a dynamic pressure of 4 lb/ft 2. This

dynamic pressure corresponds to a Reynolds number of

6

3.08 x 10 based on body length. Force and moment

coefficients were obtained over an angle-of-attack range

between -5 ° and 22 ° at sideslip angles of 0 °, -4 °, and 4 °.

Data obtained at sideslip angles of-4 ° and 4 ° were used

to compute lateral-directional derivatives. Data were also

obtained for some configurations over a sideslip angle

range from -10 ° to 10 ° at selected angles of attack.

Forces and moments were measured on an internal sixcomponent strain-gauge balance. Standard corrections

for wall interference and blockage, as described in
reference5,aswellasflowangularitycorrections, were

appliedto the data.Althoughcomparisonwith data

obtainedonthebaselineconfigurationitheLangley30-

by 60-FootTunnelshowedgoodagreementwith the

presentest,studieshaveshownthepotentialfor the

modelsupportsystemandthetunneldrivesystemto

affectthebreakdownof vortexflowsoverthemodel.

Limited laserlight-sheetflow visualizationstudies,

althoughnotshown,werealsoconductedto helpinterprettheforceandmomentdata.

Results

Longitudinal Characteristics

The effect of the configuration components on the

longitudinal aerodynamic characteristics of the model are

shown in figure 3. For the body-wing (BW) and the

body-wing-nacelle (BWN) configurations, the influence

of a strong leading-edge vortex can be seen in both the

lift and pitching-moment coefficients. In lift, this influence can be seen as a continuous increase in the lift-curve

slope with increasing angle of attack above t_ = 4 °. The

effect on pitching moment is seen as a mild pitch-up that

begins near the same angle of attack. Below t_ = 8°, both

the tail-off (BW and BWN) and tail-on (BWV and

BWNV) configurations have similar longitudinal characteristics. It is clear from both the force and moment data

in figure 3 that the vertical tails interact with the leadingedge vortex system. This interaction is evidenced by the

decrease in the lift-curve slope seen for the tail-on configurations at angles of attack above 8° and by distinct

changes in pitch stability between o_= 11 ° and ot = 15 °.

Flow visualization studies also showed that the path of

the leading-edge vortex was very close to the baseline

vertical tail location. While the decrease in lift-curve

slope was expected, the behavior in pitch was not typical

for wing-mounted vertical tails on this type of conflguration. Normally, wing-mounted vertical tails will cause

the leading-edge vortex to burst prematurely, resulting in

a severe pitch-up. For this configuration, however, addition of the vertical tails results in an increase in pitch stability before the unstable pitch-up. The reason for this

behavior is not fully understood, but is believed to result

from the load induced on the wing that is produced from

the interaction of the wing leading-edge vortices with the

vertical tails. Based on the span loading data presented in

reference 6, and illustrated in sketch A, the following

argument can be made: side loads on the vertical tails

that are produced by the wing leading-edge vortices

induce an upward load on the wing outboard of the vertical tails and a downward load inboard. Normally the

downward load and the upward load would balance out;

however, because of the presence of the fuselage, part of

the downward load is directed laterally. The result is a

net upload which produces the pitch-down moment seen

in figure 3. Above o_= 15 °, the side load produced by the

leading-edge vortices decreases as the vortices move farther from the vertical tails, and the pitching-moment

curve again becomes unstable.

The effectiveness of the tiperons for providing pitch

control is presented in figure 4. The tiperons were sized

to provide enough pitch control to trim the vehicle to

tt= 16 ° with no more than 10 ° of surface deflection;

however, there still was some concern about achieving

this capability. As can be seen from the data, this goal

was nearly achieved, as 12 ° of tiperon deflection is sufficient to trim the configuration to the desired angle of

attack. Tiperon effectiveness does decrease for deflections above 12°, however, and this reduced effectiveness

may be a concern in generating sufficient nose-down

pitch rates for recovery from angles of attack beyond

15 °. An analysis using representative weights and inertias for this type of vehicle (using approach-to-landing

conditions) was therefore made to determine the

Tail load caused by

sideslip, rudder

deflection, or vortices

Sketch A

Induced loading

on wing-body
nosedownpitch-control requirements. Inertiavalues for

the Space Shuttle orbiter (ref. 7) were used for this and

other analyses as the orbiter's weight and size are reasonably close to the LoFLYTE TM configuration's landing

configuration. Unless otherwise noted, the analyses presented in this report are based on the values listed in

table II.

The pitch-control analysis used the criteria discussed

in references 8 and 9 for relaxed pitch-stability configurations to determine the required level of nose-down

pitching moment for satisfactory recovery response. The

level of nose-down pitching moment required C m at a

given angle of attack is

, Iyq

Cm = zgs-'--I

where the value of q is -0.08 rad/sec 2. While this criterion was initially developed for stall and poststall angles

of attack, the pitch-rate acceleration requirement for MilSpec class [] vehicles (non-high-performance aircraft) is

also applicable to noncritical pitch maneuvers at all

angles of attack. The analysis was made for a trimmed

condition at cx = 15 ° (C L = 0.6). The available nose-down

ACm = -0.022 (fig. 4). Based on the equation above,

Cra = -0.011 for landing weight. This result indicates

that the present configuration would have satisfactory

nose-down recovery characteristics in terms of handing

qualities.

Figure 5 shows the trimmed values of C L and L/D as

well as the tiperon deflection required for trim that is

based on the results shown in figure 4. Because of the

pitching-moment characteristics of this configuration

(neutral to unstable static margin and positive Cm,o), the

vehicle has a reasonably high level of unpowered

trimmed CL at nominal takeoff and approach-to-landing

conditions (cx = 10°). As a result, takeoff speeds of

263 knots are possible at full gross weight (551052 lb). If

the lift component of thrust and the expected nose-up

pitching moment caused by thrust were to be accounted

for, the takeoff speed would be significantly less than

263 knots. For approach-to-landing conditions, the

approach speed for a nominal weight of 183 000 lb would

be 151 knots. While this speed is somewhat high when

compared with some conventional aircraft, it is a very

reasonable speed for this vehicle class. (Shuttle orbiter

landing speeds are in excess of 200 knots.)

Lateral-Directional Characteristics

The effect of the configuration components on

lateral-directional aerodynamic characteristics is presented in figure 6. As would be expected for a configuration with such a highly swept planform, the level of

lateral stability increases with angle of attack to quite

4

high values. These high levels of lateral stability may

have an adverse impact on landing operations in crosswind conditions. As can be seen at angles of attack above

¢x= 14°, adding the vertical tails reduced lateral stability.

Flow visualization studies indicated that at angles of

attack near 14 ° the windward leading-edge vortex has

moved very close to the outboard side of the windward

vertical tail. At the same time, the leeward vortex has

moved farther outboard but is still in proximity to the

wing upper surface. The side load generated on the windward vertical tail by the windward vortex, acting above

the center of gravity of the configuration, and the upload

on the wing outer panel generated by the leeward vortex

would tend to reduce lateral stability. As angle of attack

is increased, the vertical tails tend to force the leadingedge vortices to burst over the wing, which reduces lateral stability even further.

Without the vertical tails, the configuration, as

expected, was unstable directionally up to about cx = 10%

Above these angles of attack, the wing-body exhibits stable values of Cn_ that appear to be caused by forces aft

of the center of gravity (Cy_ is increasingly negative).

This type of behavior is usually associated with forebody

vortical flows, although these usually produce forces on

the forebody. Addition of the vertical tails generates a

stable increment that results in stable directional characteristics up to ¢x= 10 °. Between cx= 10° and ct= 18°, the

vertical tails decrease the stability of the configuration

below that of the wing-body alone. It is not unusual for

the interaction of the wing leading-edge vortex and the

vertical tails to produce these directional characteristics.

Beyond ¢x = 18 °, it appears that the wing-body

directional-stability characteristics dominate the

aerodynamics.

Figure 7 presents the effectiveness of the tiperons for

providing roll control. The tiperons, deflected asymmetrically, do generate significant rolling moment, and the

control effectiveness is fairly linear (that is, the change in

moment versus control deflection is a linear function).

Up to ct = 10 °, the levels of adverse yawing moment generated by the tiperons are independent of control deflections above +12 °. Beyond cx = 10 °, only the largest

deflections increase adverse yawing moments significantly. These relatively large values for yawing moment

caused by roll control are believed to be a result of an

induced sidewash at the vertical tails. As theorized, this

sidewash results from flow curling around the inboard

portion of the wing along the streamwise cut between the

wing and the tiperons and generates a positive sideslip at

the vertical tails for a right roll-control input.

In general, the levels of roll control would appear to

be adequate; however, because of the high levels of static
lateralstability,a crosswind analysis, to be discussed

later in this section, must be performed. This analysis

also will require the rudder effectiveness data presented

in figure 8. Like the tiperons, rudder effectiveness

appears to be linear with control deflection. As would be

expected, rudder power does decrease as angle of attack

is increased beyond about 8°. This reduction may be a

result of the large wing-body planform shielding the vertical tails and effectively reducing the dynamic pressure

at the rudders. It is more likely, however, the result of a

more complex interaction of the wing leading-edge vortices with the vertical tails, which are heavily loaded

because of rudder deflection. In this situation the

leading-edge vortices are forced into an asymmetric position relative to the vertical tails, with the port side

leading-edge vortex having the stronger influence for

positive rudder deflections. This vortex arrangement produces an incremental load on the vertical tails that

opposes that caused by rudder deflection. Unlike the

tiperons, however, deflection of the twin rudders produces a favorable cross derivative (rolling moment

caused by rudder deflection) for angles of attack below

tx = 13 °. The flow physics here are similar to those discussed previously in the Longitudinal Characteristics

section in relation to the stabilizing effect the vertical

tails had on pitching moment• (See sketch A.) In this case

the side loads on the vertical tail that are caused by

positive rudder deflection result in an induced wing loading that produces a negative increment to rolling

moment• Because the wing is so large, this increment is

greater than the positive rolling moment produced by

forces on the vertical tail. The net result is a configuration with proverse rolling moment caused by rudder

deflection. Consequently, rudder inputs will tend to be

self-coordinating and will require less asymmetric

tiperon deflection to make a coordinated turn. As will be

discussed later in this section, this self-coordination will

have a positive impact on the crosswind capabilities of

this configuration because of the high levels of adverse

yaw generated by the tiperons.

Because approach to landing probably will be the

most demanding low-speed flight phase for this vehicle,

the conditions chosen for the crosswind and coordinated

roll analyses were as follows: tx = 10 °, V = 151 knots,

W = 183 000 lb. The crosswind analysis involves solving

the following set of simultaneous algebraic equations for

St and _r:

C% _ + C n ,5t _t + On,5 r _r = 0

Ct_ 13+ Ct_' gt + Cts _r = 0

where 13= tan -1 Vxw

Vapp

The solution for these equations is shown in figure 9

for et = 10°. In general, this type of vehicle must demonstrate the ability to land in a 30-knot crosswind. To trim

out the sideslip generated by this crosswind and to align

the vehicle with the runway centerline would require 22 °

of asymmetric tiperon and 13 ° of rudder deflection (by

using control derivatives based on full-control deflection). Because of the high levels of adverse yaw generated by the tiperons, both the rudders and the tiperons are

deflected positively for negative sideslip angles (trimming at sideslip usually requires the vehicle to be cross

controlled). Obviously, the rudder requirement is reasonable, as only half the available control authority is used.

The tiperon requirement, however, is more severe. When

combined with the necessary symmetric tiperon deflection for pitch trim (which reduces the available asymmetric deflection to 25°), one of the tiperons always will be

deflected nearly to its limits at these conditions. These

levels of control deflections will greatly reduce the nosedown pitch and roll-control margins available for maneuvering and stability augmentation. A reduction in lateral

stability or nonzero sideslip approaches are two potential

solutions to this problem.

While the ability to make a velocity vector roll in

this vehicle may not be mandatory, it is desirable and the

associated analysis does provide an indication of the relative balance between roll and yaw control. By using the

vehicle equations of motion, it is possible to develop a

relationship between roll and yaw control so that a rolling maneuver can be made without generating any sideslip. Like the nose-down pitch-control analysis presented

earlier, this analysis requires knowing the moments of

inertia of the vehicle• As noted before, inertia values

have been estimated by using the Shuttle orbiter data.

Starting with the kinematic relation between a stability

axis roll rate (with no yaw rate in the stability axes) and

the body axes roll and yaw rates,

P = Ps COS O_

r = Ps sin ot

and the relationship between rate acceleration and control moments,

Ctr_t_t_lSb

pI x

C nsr_ rilSb

i'-

Iz

an approximate relationship between rudder and asymmetric tiperon deflection for a velocity vector roll can be

developed when angle of attack is assumed to remain

constant. This relationship is

5
This simplified analysis indicates that for roll

maneuvers at o_= 10 °, 6.4 ° of rudder deflection would be

needed for every 1o of asymmetric tiperon deflection.

With a rudder deflection limit of 30 °, this would mean

that a velocity vector roll would be possible only for

asymmetric tiperon deflections of less than 5°. As a

result, if larger roll rates are desirable, a way to improve

rudder power must be found. It should be noted, however, that this analysis does not account for damping

caused by velocity vector roll rate. If the damping terms

are large, this analysis is only valid for the initiation of

the rolling maneuver.

Alternate Vertical Tail Configurations

In order to address the two directional stability and

control problem areas of poor rudder power and directional instability between (x = 10° and o_= 18°, alternate

locations for the vertical tails were investigated. Four

additional geometries were evaluated: (1) 8 in. outboard

of the baseline location, (2) 4 in. inboard of the baseline

location, (3) a single centerline vertical tail, and (4) 4 in.

inboard and 4 in. aft of the baseline location. The aerodynamic characteristics for the first three configurations are

compared to the baseline tails in figure 10. Alternate tail

locations had little impact on longitudinal characteristics

below (x = 10 ° (fig. 10(a)). Above this angle of attack,

changes in both lift and pitching moment are indicative

of changes in the way the vertical tail interacts with the

wing leading-edge vortex system. The outboard tail location apparently causes the leading-edge vortex to burst,

which results in a pitch-up and in a reduction in the liftcurve slope. Flow visualization studies showed no vortex

bursting for the inboard tail location although the behavior of the longitudinal data does show some pitch-up

when compared with the centerline tail.

As can be seen in both the lateral and directional data

(fig. 10(b)), all three alternate geometries eliminated the

adverse interaction between the wing leading-edge vortex and the vertical tails. While the outboard location

provided the largest stabilizing increment, the level of

directional stability associated with the inboard twin-tail

location is adequate for this vehicle. The inboard location

also allows for the tails to be moved farther aft and eliminates any potential physical interference problem

between the rudder and the tiperons, thus allowing the

notch in the baseline rudder to be filled. The effect of the

alternate tail locations on rudder power is presented in

figure 10(c). The results of the alternate twin-tail locations are mixed and indicate the complex nature of the

interaction between the wing leading-edge vortex system

and the vertical tails. Rudder power for the centerline

tail, while well behaved, is greatly reduced because of

the decrease in rudder volume for this configuration.

This reduced rudder power is not surprising because the

centerline tail was the same geometry as one of the baseline twin tails.

Figure 11(a) shows that moving the tails aft had only

a slight effect on lateral-directional stability, namely, by

providing a small increase in the level of directional stability. Note that the aft tail data were obtained with the

rudder notch filled. There was, however, a significant

increase in the effectiveness of the rudders for the aft tail

location. Figure 1l(b) compares the available rudder

power for the forward and aft tail positions. Moving the

tails aft and filling the rudder notch resulted in close to a

50-percent increase in rudder power. This increase would

allow velocity vector rolls with up to 8 ° of asymmetric

tiperon deflection. While this increase is small, it may

make the rolling performance of the LoFLYTE TM vehicle

more acceptable.

Concluding Remarks

In general, the low-speed characteristics of the

LoFLYTE TM configuration (Accurate Automation Corporation, Chattanooga, Tennessee) appear satisfactory.

Because of the pitching-moment characteristics of this

configuration (neutral to unstable static margin and positive Cm,o), the vehicle has a reasonably high level of

unpowered trimmed CL at nominal takeoff and approachto-landing conditions (ct = 10°). These levels of unpowered trimmed C L should allow for acceptable takeoff and

landing speeds for this vehicle. Locating the vertical tails

inboard of the baseline location improved a directionalstability problem between tx = 10 ° and tx = 18 °. Lateraldirectional stability and control characteristics are such

that crosswind and velocity vector roll criteria can be

met, although control saturation remains an issue in both

cases. Reduction in static lateral stability or a modified

approach-to-landing profile (i.e., a nonzero sideslip

approach) would alleviate the saturation problem during

crosswind landings. Higher levels of yaw control are

necessary to address the issue of velocity vector rolls.

Moving the vertical tails inboard and aft of the baseline

location does improve rudder power; however, the

increase in the roll-rate envelope is small.

NASA Langley Research Center

Hampton, VA 23681-0001

January 8, 1997
References

1. Pegg, Robert J.; Hunt, James L.; Petley, Dennis H.; Burkhardt,

Leo; Stevens, Daniel R.; Moses, Paul L.; Pinckney, S. Z.;

Kabis, Hanee Z.; Spoth, Kevin A.; and Dziedzic, William M.:

Design of a Hypersonic Waverider-Derived Airplane. AIAA93-0401, Jan. 1993.

2. Townend, L. H.: Research and Design for Lifting Reentry.

Prog. Aerosp. Sci., vol. 19, 1979, pp. 1-80.

3. Capriotti, Diego; Bowcutt, Kevin G.; and Anderson, John D.,

Jr.: Viscous Optimized Hypersonic Waveriders. AIAA-87-

0272, Jan. 1987.

4. Anderson, John D., Jr.; Lewis, Mark J.; and Corda, Stephen:

Several Families of Viscous Optimized Waveriders--A

Review of Waverider Research at the University of Maryland.

Proceedings of the 1st International Hypersonic Waverider

Symposium, Univ. of Maryland, Oct. 1990, p. 39.

5.

6°

7.

8°

9.

Pope, Alan; and Harper, John J.: Low-Speed Wind Tunnel Testing. John Wiley & Sons, 1966.

Queijo, M. J.; and Riley, Donald R.: Calculated Subsonic Span

Loads and Resulting Stability Derivatives of Unswept and 45 °

Sweptback Tail Surfaces in Sideslip and in Steady Roll. NACA

TN 3245, 1954.

Boyden, R. P.; and Freeman, D. C., Jr.: Subsonic and Transonic Dynamic Stability Characteristics of a Space Shuttle

Orbiter. NASA TN D-8042, 1975.

Nguyen, Luat T.; and Foster, John V.: Development of a Preliminary High-Angle-of-Attack Nose-Down Pitch Control

Requirement for High-Performance Aircraft. NASA

TM- 101684, 1990.

Ogburn, Marilyn E.; Foster, John V.; Pahle, Joseph W.;

Wilson, R. J.; and Lackey, James B.: Status of the Validation

of High-Angle-of-Attack Nose-Down Pitch Control Margin

Design Guidelines. AIAA-93-3623, Aug. 1993.

7
Table I. Geometric Characteristics of LoFLYTE TM Model

Overall length, ft ...................................................... 8.34

Wing-body:

Span, fi ............................................................ 5.19

Area, ft 2 ........................................................... 22.62

Reference length, fi .................................................. 8.34

Aspect ratio ......................................................... 1.19

Tip chord, ft ........................................................ 0.48

Inboard leading-edge sweep, deg ....................................... 75.13

Outboard leading-edge sweep, deg ...................................... 65.00

Trailing-edge sweep, deg ............................................. -5.00

• 2

Tlperon area (each), ft ............................................... 1.14

Baseline twin vertical tail:

Span, fi ........................................................... 0.59

Area (each), ft 2 ...................................................... 0.87

Root chord, ft ....................................................... 2.86

Aspect ratio ........................................................ 0.40

Leading-edge sweep, deg ............................................ 75.00

Rudder area (each), ft2 ............................................... 0.14

Twin vertical tail with extended rudder:

Span, ft ........................................................... 0.59

Area (each), ft 2 ...................................................... 1.03

Root chord, ft ....................................................... 2.86

Aspect ratio ........................................................ 0.34

Leading-edge sweep, deg ............................................. 75.00

Rudder area (each), ft2 ................................................ 0.30

Centerline vertical tall:

Span, ft ........................................................... 0.59

Area (each), fl2 ...................................................... 0.87

Root chord, ft ....................................................... 2.86

Aspect ratio ........................................................ 0.40

Leading-edge sweep, deg ............................................. 75.00

2

Rudder area (each), ft ................................................ 0.14
TableII. DataUsedforAnalyticalCalculations

[Datafortrimmedconditionsattx = 10°; inertias are for landing weight of 183 000 lb]

WTO, lb ................................................ 551 052

Wlanding, lb ............................................. 183000

Ix, slugs-ft 2 ............................................. 747 424

ly, slugs-ft 2 ............................................ 5682812

2

Iz, slugs-ft ............................................ 5 800 737

2

S, ft ..................................................... 5884

b, ft ...................................................... 83.7

l, ft ..................................................... 134.5

cL ....................................................... 0.38

LID ....................................................... 3.9

_it, deg ..................................................... 5.5

C%, per deg .............................................. 00015

Ct, per deg .............................................. -.003

C n , per deg ............................................. 00027

8 t

Ct6 t, per deg ............................................. -.0015

Cns r, per deg ........................................... -.00032

Cts r, per deg ........................................... -.00003
.JJ'T3

Baseline rudder

Extended rudder

5.48

(a) Three-view sketch of model.

i

(b) Photograph of LoFLYTE TM model.

Figure 1. Test configuration. All dimensions in feet unless otherwise noted.

10
Figure2. Photographof test setup in Langley 12-Foot Low-Speed Tunnel.

11
C m

.08

.06

.04

.02

[ Configuration

I o BW

I [] BWN

i <> BWV

A BWNV

CL

1.0

.8

.6

.4

.2

-.2

-.4 i , i ,

-5 0

I

: i

i

i !

Illllillll

15 20 25

!

I I I I I I I I I I I I

5 10

Oq deg

Figure 3. Longitudinal characteristics of test configuration.

12
.10

.08

.06

0

5 10 15

Or, deg

Figure 4. Pitch-control effectiveness.

20 25

13
C L,tr

(LID) tr 3

1

0

30

i

_it,tr, deg

25

20

15

10

5

0

-5

-5 0 5 10 15

c¢, deg

Figure 5. Trimmed longitudinal characteristics.

20 25

14
Clf_

Cnp

%

.002

0

-.002

i

: _ Configuration

i i o BW

i i rn BWN I

! <) BWV

Zx BWNV

-.004

-.006

.002

.001

-.001

-.002

.002

!

• I

0 O____o.__ ___O_____<_ _ _ !_

-.004 A_A_ _"'_&--&_ _ lX'_'_

-.006 .................

i

-.008 i

i

-.010 , , , i .................... i

-5 0 5 10 15 20 25

(X, deg

Figure 6. Lateral-directional characteristics of test configuration.

15
Ac/

.04 •

.02

-.02

-.04

O-----O------

.... 15t, deg L/R

! o -12/12

[] -19119

_ -30130

r

I

i

!

i

l

J

AC n

ACy

.02

I

I

I

i

T

[

i

i

I

]

I

I

i

i

i

i

0 10

or, deg

Figure 7. Roll-control effectiveness.

15 20 25

16
.OO6

.OO4

.002

r_deg L/R

-- i

i o 14/14

D 25/25

AC l

AC

n

AC Y

.0O4

I

.03 - - --

.02

.01

0

-.01 ....

-5

I I t I [ I I I I

10

_, deg

i i I i i i i i

15 20 25

Figure 8. Rudder effectiveness.

17
35

_tor8r,

deg

30

25

20

15

10

5

I I

Maximuml mdderl'-_

- Maximumasymmetrictperon

-(maintainingzeropitchingmoment)

f

J

St f

8r

5 10 15 20 25

f

0 30

Crosswind velocity, knots

Figure 9. Crosswind landing-control requirements, ct = 10 °, Vapp -- 151 knots.

18
C

m

.08

.06

.04

.02

0

-.02

Tail position

o Baseline

[] Centerline

O 4 in. inboard

A 8 in. outboard

q

0 5 10 15 20

0%deg

(a) Longitudinal characteristics.

Figure 10. Effect of vertical tail location.

25

19
ctp

.0O2

-.002

-.004

- - Tail position

o Baseline

[] Centerline

, i 0 4 in. inboard

I I tx 8 in. outboard

I

i i

!

I

i

i

Cnp

C

-.006

.002

.001

-.001

0

i !

I

5 10 15 20

ix, deg

25

(b) Lateral-directional characteristics.

Figure 10. Continued.

2O
,,cl

AC n

ACy

.006

.0O4

.002

Tail position

o Baseline

[] Centerline

(> 4 in. inboard

/x 8 in. outboard

I

3

5 10 15 20

Or, deg

(c) Rudder effectiveness. _r = 25°/25°"

25

Figure 10. Concluded.

21
Clp

C

np

Cyp

.OO2

-.002

-.004

-.006

.002

.001

-.001

-.002

.002

-.002

-.004

-.006

-.008

-.010

-5

i

Tail position

o Forward

o Aft

I

i

i l

I

i

; I

I

I

[3---q3---

0----0---

i

i

i

i

J

i i I .... i .....

0 5 10 15

tx, deg

i

i

i I

I

I

I I I I I

2O

I I

25

(a) Lateral-directional characteristics.

Figure 11. Effect of vertical tail longitudinal position.

22
AC l

AC

12

ACy

.006

.004

.002

-.002

-.004

-.006

.004

-.004

-.008

-.012

-.016

-.020

.04

::l i[]°p°sti

J

I

1

-- T r

i

I

I

I

i

I

5 10 15 20 25

a, deg

(b) Rudder effectiveness. _r = 25°/25°"

Figure 11. Concluded.

23


REPORT DOCUMENTATION PAGE _,,._o,o_

OMB No. 07040188

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1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED

May 1997 Technical Memorandum

4. TITLE AND SUBTITLE 5. FUNDING NUMBERS

Evaluation of the Low-Speed Stability and Control Characteristics of a

Mach 5.5 Waverider Concept WU 505-70-69-01

5. AUTHOR(S)

David E. Hahne

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

NASA Langley Research Center

Hampton, VA 23681-0001

9. SPONSORING/MONITORINGAENCYNAME(S)ANDADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001

8. PERFORMING ORGANIZATION

REPORT NUMBER

L-17581

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA TM-4756

11. SUPPLEMENTARY NOTES

1211. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified-Unlimited

Subject Category 08

Availability: NASA CASI (301) 621-0390

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)

Static force and moment tests of a 0.062-scale model of a hypersonic vehicle study concept known as the

LoFLYTE TM configuration were conducted in the Langley 12-Foot Low-Speed Tunnel. These tests looked primarily at the low-speed static stability and control characteristics of this configuration. Data were obtained over an

angle-of-attack range of-5 ° to 22 ° at sideslip angles that ranged between -10 ° and 10 °. The tiperons were sized to

provide enough pitch control to trim the vehicle up to (x = 16° with no more than 10 ° of surface deflection and data

obtained in this test showed that 10° of tiperon deflection was nearly sufficient to trim the configuration up to the

desired angle of attack. Because of the pitching-moment characteristics of the LoFLYTE ru configuration, there is a

reasonably high level of unpowered trimmed lift at nominal takeoff and approach to landing that should allow for

acceptable takeoff and landing speeds for this vehicle. Initial evaluation of the directional stability characteristics of

this configuration showed a significant instability between ct = 10° and about (x = 18°. This test determined that the

cause of this instability was the interaction of the wing leading-edge vortex with the vertical tails. Moving the vertical tails either inboard or outboard from the baseline location eliminated this unfavorable interaction.

14. SUBJECT TERMS

Low speed; Waverider; Stability and control; Vortical flows; LoFLYTE TM

17. SECURITY CLASSIFICATION

OF REPORT

Unclassified

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Unclassified

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A03

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