NASA Technical Paper 1552 

Aerodynamic Characteris tics of a 

11 

NASA \ , 

Tp 

1552 ,,- 

c .1 

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 

Jim A. Penland, James B. Hallissy, 

and James L. Dillon 

.. 

" 

DECEMBER 1979 
.. ... . ." 

NASA Technical Paper 1552 

TECH LIBRARY KAFB, NM 

OL34bOb 

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 

Jim A. Penland, James B. Hallissy, 

and James L. Dillon 

Langley Research Center 

Hmnptorz, Virginia 

National Aeronautics 

and Space Administration 

Scientific and Technical 

Information Branch 

1979 
SUMMARY 

The data of the present report are divided into three areas. The first area includes 

the results of a wind-tunnel investigation of the static longitudinal, lateral, and directional stability characteristics of a hypersonic research airplane concept having a 70' 

swept double-delta wing. The force tests were conducted in the Langley 8-foot transonic 

pressure tunnel for Mach numbers from 0.80 to 1.20, for a Reynolds number (based on 

fuselage length) range of 6.30 x lo6 to 7.03 x lo6, at angles of attack from about -4' 

to 23O, and at angles of sideslip of 0' and 5'. The configuration variables included the 

wing planform, tip fins, the center vertical tail, and scramjet engine modules. The 

second area is a summary of the variations of the more important aerodynamic parameters with Mach number for Mach numbers from 0.20 to 6.0. The third area is a stateof -the-art example of theoretically predicting performance parameters and static longitudinal and directional stability over the Mach number range. 

The model with tip fins had more linear lift curves than models with the center 

vertical tail but showed more tendency to pitch up and had less lateral and directional 

stability at the higher angles of attack at low and transonic speeds. The forward delta 

wing decreased the longitudinal stability and maximum lift-drag ratio but enhanced the 

positive dihedral effect. The highest maximum lift-drag ratio was measured at the lowest 

subsonic speed, and the lowest lift-drag ratio at Mach number 6. The model scramjet 

engine had only small positive effects on the lateral and directional stability over the 

wide Mach number test range. Adequate preliminary predictions of performance and 

static longitudinal stability may be expected from presently available theories on models 

having body fineness ratios of about 7, but nonempirical predictions of directional stability 

are either unsatisfactory or are beyond the state of the art at subsonic and supersonic 

speeds and are overly optimistic at hypersonic speeds. 

INTRODUCTION 

Present jet airplanes are cruising at speeds of Mach 2 to 3 (refs. 1 to 3), and it 

appears that the Mach number limit for aircraft utilizing conventional petroleum-based 

fuels is about 5 (i.e., turbojets to about 3.5 and ramjets to about 5, ref. 4). Some unique 

problems associated with Mach numbers of 5 and above include the development of new 

propulsion systems which use nonpetroleum-derived fuels such as liquid hydrogen (ref. 5). 

Some of these new propulsion systems include cryogenic-fueled turbojets for low speeds, 

I 
ramjets (subsonic combustion) for moderate supersonic speeds, and scramjets (supersonic combustion ramjets) for high supersonic and hypersonic speeds. New structural 

concepts must be developed which can provide cooled airframes and engine surfaces for 

protection from high aerodynamic heating as well as insulated tankage for cryogenic fuels 

such as liquid hydrogen. 

One industry study (refs. 6 to 9) concluded that only through the use of both ground 

facilities and flight vehicles could these major required advancements in technology be 

made. These findings were in accord with previous NACA-NASA experience with the 

various research airplane projects from the X-1 to the X-15, each of which resulted in an 

extensive technology advancement at a minimum expenditure of cost and time (ref. 10). 

In addition to the unique problems previously outlined, correlation of ground test data 

(forces, moments, and aerodynamic heating all measured at relatively low Reynolds numbers in mixed laminar-turbulent boundary layers) with flight test data is required for 

future efficient design. A need thus exists for comprehensive flight research in the range 

of Mach numbers from 3 to 5 and for detailed exploration to Mach 8. 

The present configuration is one of several research airplane concepts under study 

at the NASA Langley Research Center (refs. 11 to 14) that meets the requirements 

envisioned as necessary to provide a technology base for future high-speed aircraft. 

This configuration was the result of a preliminary design study utilizing computerized 

aerodynamic, structural, and weight and balance programs. The requirements of largevolume fuel tanks for the low-density liquid hydrogen fuel and sufficient depth between the 

airstream surfaces and the tanks and primary structure for thermal protection insulation 

resulted in the low-fineness-ratio fuselage of about 7 and the 5- to 6-percent thick wing. 

Wing planform area was based on landing criteria, and vertical-tail design on hypersonic 

directional stability requirements. Such a research airplane would be air launched from 

a B-52 or C-5 and would have a length of 15.24 to 24.38 m (50 to 80 ft), a flight time of up 

to 800 sec with a nominal 40-sec cruise at a Mach number of about 7 on the scramjet 

engine, and a return to base for a dead-stick landing. In-flight tests would include powerless glides, rocket-boosted flights, and combined rocket-scramjet boost-cruise 

experiments. 

The data of the present report are divided into three areas. The first area includes 

the results of a study designed to experimentally investigate the longitudinal, lateral, and 

directional stability and control of a large-fuselage, 70' swept double-delta wing configuration at transonic speeds. Studies of this same design concept at subsonic, supersonic, 

and hypersonic speeds (refs. 15 to 21) have also been completed. This study was conducted at Mach numbers of 0.80 to 1.20 and a Reynolds number (based on fuselage length) 

range of 6.30 X lo6 to 7.03 X lo6 in the Langley 8-foot transonic pressure tunnel. The 

angle-of-attack range was from about -4' to 23O, with angles of sideslip of 0' and 5'. 

2 
I . . . 

The tests were parametric in nature and included configuration buildup and variations in 

, wing planform, longitudinal control, and roll control. The second area of this paper 

summarizes the variations of the more important aerodynamic parameters from previous tests (refs. 15 to 21) with Mach number for Mach numbers from 0.20 to 6.0. The 

third area presents a state-of-the-art example of theoretically predicting performance 

parameters and static longitudinal and directional stability over the Mach number test 

range . 

SYMBOLS 

The longitudinal characteristics are presented about the stability axes, and the 

lateral-directional characteristics are presented about the body axes. The body and 

stability axis systems are illustrated in figure 1. The moment reference point was at 

the design center-of-gravity location which was at a longitudinal station 64.5 percent 

of the fuselage length and at a vertical station 1.25 percent of the fuselage length below 

the vehicle reference line. Values are given in both SI and U.S. Customary Units. The 

measurements and calculations were made in U.S. Customary Units. 

A reference area, area of 70'delta wing including fuselage intercept, 

0.043 m2 (67.2 in2) 

b wing span, 0.217 m (8.542 in.) 

'A,b base axial-force coefficient, -- 

Base axial force 

q.2 

cD 

cL 

drag coefficient, - 

D 

qm-4 

lift coefficient, - L 

qooA 

C rate of change of CL with angle of attack, per degree La! 

MX 

q,Ab 

rolling-moment coefficient, - 

C rate of change of Cz with angle of sideslip, per degree 

3 
cz rolling-moment coefficient due to roll control, per degree 

6H 

MY 

Cm pitching-moment cofficient, - q,AQ 

cma! 

rate of change of Cm with angle of attack, per degree 

aCm/XL rate of change of Cm with lift coefficient, longitudinal stability parameter 

(aCnpCy) (b/~) rate of change of Cn with side-force coefficient, directional 

stability parameter 

Cn 

CnB 

Cn 

6H 

CY 

C 

yP 

C 

c.g. 

D 

FA 

FN 

FY 

4 

MZ 

q,Ab yawing-moment coefficient, - 

rate of change of Cn with angle of sideslip, per degree 

yawing-moment coefficient due to roll control, per degree 

FY 

%oA 

side-force coefficient, - 

rate of change of Cy with angle of sideslip, per degree 

side-force coefficient due to roll control, per degree 

design center of gravity, moment reference 

drag, FN sin CY + FA cos CY 

aial force along X-axis (positive direction is -X) 

normal force along Z-axis (positive direction is -Z) 

side force along Y-axis (positive direction is +Y) 
L lift, FN cos CY - FA sin CY 

L/D lift-drag ratio 

Q length of model fuselage, 0.508 m (20.0 in.), reference length 

M Mach mber 

MX' My, MZ moments about X-, Y-, and Z-axes 

free-stream dynamic pressure 

Reynolds number based on fuselage length 

reference axes 

angle of attack, degree 

angle of sideslip, degree 

elevon deflection angle, positive when trailing edge is down, degree 

differential elevon deflection, positive to provide positive roll, 

(6H,left - 'H,right)' degree 

Subscripts: 

max 

min 

0 

S 

t 

maximum 

minimum 

condition at zero lift 

stability axis system 

trim condition, Cm = 0 

5 
Model nomenclature: 

B body with high profile nose 

E model scramjet engine 

FD 

vC 

vT 

forward delta wing 

center vertical tail 

tip fins, vertical 

W wing 

MODEL 

A photograph of a model of the winged hypersonic airplane configuration is shown 

in figure 2(a). The test model was of modular design, as shown in figure 2(b), to allow 

the buildup of variations of the basic model from components consisting of the body, a 

forward delta wing, a 70' swept clipped-delta wing, vertical tip fins, a center vertical 

tail, and an engine. Geometric details of the model are shown in figure 3 and are given 

in table I. The model design rationale was primarily based on the stability and control 

requirements at the design hypersonic cruise Mach number range up to 8. The forward 

delta wing was included in the design to help decrease the rearward shift of the aerodynamic center with increasing Mach number. The tip fins (fig. 3(b)) were designed with a 

7.5' toe-in and were located outboard of the fuselage flow field to enhance the directional 

stability at hypersonic speeds. A center vertical tail (fig. 3(c)) having a wedge airfoil 

section and the same total planform area as the sum of the tip fins was tested to assess 

the difference in directional stability between it and the tip fins. Elevons having a combined area equal to 10.64 percent of the reference area could be deflected from -20' 

to 5' in 5' increments. A model scramjet engine (fig. 3(d)) was used to complete the 

configuration buildup. The engine consisted of six clustered modules of the concept 

described in reference 22, each having outside dimensions, angles, and areas to scale, 

but with the inside fuel struts and compression surfaces designed to provide an internal 

contraction ratio of approximately 2 (compared with about 4 for the flight engine). This 

increased internal area was provided to partly account for the relatively low Reynolds 

number and the resulting thick boundary layer of this test. Details of these model engine 

inlet flow fields and boundary layers may be found in references 23 and 24. 

6 
APPARATUS AND TESTS 

Tunnel 

The investigation was conducted in the Langley 8-foot transonic pressure tunnel. 

The test section of this tunnel is 2.164 m (7.1 ft) square in cross section, with longitudinal slots in the upper and lower walls to permit changing the test section Mach number 

continuously from 0 to over 1.20, with negligible effects of choking and blockage. The 

span of the present small model was only 1/11 the tunnel diameter, and no wall corrections were required. The sting diameter was 0.0254 m (1.00 in.). 

Test Conditions 

Tests were made at Mach numbers from 0.80 to 1.20, at a stagnation pressure of 

1 atm (1 atm = 101.3 kPa), and at a stagnation temperature of 322 K. The Reynolds number and dynamic pressures of the investigation are presented in the following table: 

0.80 

.90 

.95 

.98 

1.10 

1.20 

R I 

6.30 X lo6 

6.63 

6.77 

6.83 

7.00 

7.03 

N/cm2 

2.99 

7.16 

3.59 

3.69 

4.02 

4.22 

psia 

4.33 

4.94 

5.21 

5.35 

5.83 

6.12 

The tunnel air was dried sufficiently to avoid condensation effects. The angle-of -attack 

range was from about -4' to 23' for angles of sideslip of 0' and 5'. In order to insure a 

turbulent boundary layer, 0.159-cm (0.063-in.) wide strips of No. 150 carborundum 

grains were applied 1.0 cm (0.394 in.) streamwise aft of the leading edges of the wings 

and the tail surfaces; a 0.159-cm (0.063-in.) wide strip of No. 150 carborundum grains 

was applied at the nose at a station 3.0 cm (1.181 in.) streamwise from the apex 

(refs. 25 and 26). 

Measurements and Corrections 

The aerodynamic forces and moments were measured by means of a six-component 

strain-gage balance which was housed within the model fuselage. Balance-chamber pressure was measured with pressure tubes located in the vicinity ofthe balance. 

7 
Angles of attack and sideslip have been corrected for the deflection of the balance 

and sting due to aerodynamic loads. The angle of attack was corrected for tunnel flow 

angularity. The drag coefficients were corrected to the condition of free-stream static 

pressure in the balance chamber. The base perimeter of the model was internally 

feathered to equalize the pressure over the base. Typical variations of base axial-force 

coefficient measured during the tests are presented in figure 4 for the BWVTFDE and 

BWVTFD models at the various Mach numbers tested. No correction was made to the 

drag data for flow through the model scramjet engine. 

PRESENTATION OF RESULTS 

The results of the wind-tunnel investigation in the Langley 8-foot transonic pressure tunnel are presented in the following figures: 

Figure 

Base axial-force coefficients at M = 0.80 to 1.20 for BWVTFDE 

and BWV F configurations ........................... TD 4 

Longitudinal stability characteristics at M = 0.80 to 1.20 for - 

Component buildup of tip-fin configurations .................. 5 

Component buildup of center-vertical-tail configurations ........... 6 

Comparison of tip-fin and center-vertical-tail configurations ......... 7 

Elevon effect on longitudinal characteristics at M = 0.80 to 1.20 for - 

BWVTFDE configuration ............................. 8 

BWVcFD configuration ............................. 9 

Comparison of longitudinal characteristics at trim of BWVTFDE and 

BWV F configurations at M = 0.80 to 1.20 .................. 10 CD 

Lateral-directional stability characteristics at M = 0.80 to 1.20 for - 

Component buildup of tip-fin configurations ................... 11 

Component buildup of center-vertical-tail configurations ........... 12 

Comparison of tip-fin and center-vertical-tail configurations ......... 13 

Roll control for BWVTFDE configuration at M = 0.80 to 1.20 .......... 14 

A summary of the variation of the more important aerodynamic parameters with 

Mach numbers is presented in the following figures for M = 0.20 to 6.0: 

Test Reynolds numbers and angles of attack/lift coefficients ........... 15 

Component buildup of tip-fin configurations: 

Minimum drag coefficient and maximum lift-drag ratio ............ 16 

Pitching-moment coefficient and lift-curve slope at CL = 0 ......... 17 

8 
Figure 

Longitudinal stability at CI, = 0 and 0.2 ..................... 18 

Lateral-directional stability at CL = 0 and 0.2 ................. 19 

Comparison of complete configuration with and without engine: 

Minimum drag coefficient and maximum lift-drag ratio ............ 20 

Pitching-moment coefficient and lift-curve slope at CL = 0 .......... 21 

Longitudinal stability at CL = 0 and 0.2 ..................... 22 

Lateral-directional stability at CL = 0 and 0.2 ................. 23 

Component buildup of center-vertical-tail configurations: 

Minimum drag and maximum lift-drag ratio ................... 24 

Pitching-moment coefficient and lift-curve slope at CL = 0 .......... 25 

Longitudinal stability at CL = 0 and 0.2 ...................... 26 

Lateral-directional stability at CL = 0 and 0.2 ................. 27 

A comparison of theoretical estimates and various experimental longitudinal and 

directional parameters on the WBVT configuration at M = 0.20 to 6.0 is presented in 

the following figures: 

Minimum drag coefficient and maximum lift-drag ratio .............. 28 

Lift coefficient and lift-curve slope at a = 0' ................... 29 

Longitudinal and directional stability ........................ 30 

RESULTS AND DISCUSSION 

All the basic data presented in this paper were machine plotted from the computational tape, and the data points were machine faired using the cubic spline method. 

Because of the preliminary nature of this study, individual curves such as L/D plotted 

against CL were not given detailed cross-checks against polar plots of CL plotted 

against CD. 

Static Longitudinal Aerodynamics 

__ Component buildup of model with tip fins. - Variations of the longitudinal characteristics with component buildup are presented in figure 5 for the six transonic Mach numbers tested. A comparative study of these data shows a trend of increasing Cm from 

a negative value at M = 0.80 to a positive value at M = 0.95 and above for all variations 

of the configurations tested, from the basic body-wing to the complete model with engine. 

This positive increase in was accompanied by a positive shift in the angle for zero 

lift. All configurations exhibited a small increase in longitudinal stability with increasing 

subsonic Mach number. 

0 

cmO 

9 
As expected, the drag increases with subsonic Mach number, reaches a maximum 

for the present data at M = 0.98, and then decreases with supersonic Mach number. The 

lift-drag ratio is highest at the lowest test Mach number of 0.80 and inversely follows the 

trends of the drag due in part to the nearly constant lift-curve slope. Whereas the addition of the forward delta wing FD to the nonfinned configurations made relatively small 

changes to the lift and drag, the addition of the tip fins made significant increases in both 

lift and drag; thus, about a 20-percent loss in lift-drag ratio resulted. The gain in lift 

may be attributed to an alteration of the wing-tip flow field by the tip fins acting as end 

plates and thereby reducing the tip losses. The increase in drag was due to skin friction 

and to the drag due to the 7.5' toe-in of the tip fins. 

Component buildup on model with center vertical tail.- The use of a center vertical ". ~" ~ . "~ - . . - 

tail (fig. 6) in lieu of twin tip fins produces some trends similar to those produced with 

the tip fins i.e., the Cm moves from negative to positive, a. increases positively, 

and the configurations become slightly more stable with increasing Mach number . One 

noticeable difference was a slight loss of lift with the addition of the center tail to the 

body-wing configuration, whereas there was a gain in lift with the addition of the tip fins, 

as previously discussed. The increase in drag with the addition of the center tail was 

considerably less than with the addition of the tip fins. This small increase in drag and 

slight loss in lift combined to produce a decrease of approximately one-half unit in maximum lift-drag ratio, as compared with a loss of about one unit for the addition of the 

tip fins. 

( 0 

) 

Comparison of tip fins and ". . center . vertical tail.- A comparison of models equipped 

with tip fins and a center vertical tail is presented in figure 7 both with and without the 

forward delta wing. At all test Mach numbers, the models with tip fins exhibit higher lift 

coefficients and more linear lift curves with angle of attack than the models with the tenter vertical tail. The addition of the forward delta wing to either fin configuration made 

only small positive increases to lift as a function of angle of attack. The pitching-moment 

curves plotted as a function of lift coefficient for the center-tail models were slightly 

more linear and showed less tendency to become neutrally stable or to pitch up at the 

higher angles of attack. The addition of the forward delta wing increased the drag through 

the range of medium lift coefficients at all test Mach numbers and contributed to the consistent trend of lower maximum lift-drag ratios for those models equipped with the forward delta wing. The drag of the tip-fin models was substantially higher than the drag 

of the center-tail models, thus overcoming the higher lift and resulting in consistently 

lower maximum lift-drag ratios for the tip-fin models at all Mach numbers. 

- 

10 
Trim Characteristics 

The effect of elevon deflection on the longitudinal aerodynamic characteristics of 

the complete tip-fin configuration BWV F E with scramjet engine is presented in ( TD) figure 8 for all test Mach numbers and in figure 9 for the center-vertical-tail model 

( BWVCFD) without engine. Data were obtained at elevon deflections of Oo, -loo, 

and -20° for both configurations. 

The longitudinal characteristics at trim are presented in figure 10 for both test configurations. The trends of the performance parameters at trim are similar for the two 

configurations through the transonic Mach number test range from M = 0.80 to 1.20, with 

some exceptions, one of which is the 0.5 to 1.5 loss in maximum lift-drag ratio due to the 

combined scramjet engine drag and trim drag required to trim the incurred incremental 

pitching moments. The trim angle-of-attack range is also up to 3' greater at Mach numbers of 0.98 and higher for the complete configuration than for the center-vertical-tail 

model without engine. The lift-curve slopes were similar in trend and magnitude for the 

two tests, whereas the angle of attack for zero lift was consistently lower by about 0.5' 

to lo for the complete tip-fin model with engine. 

Both configurations were statically longitudinally stable throughout the test angleof-attack range at the design center of gravity, with the static margin varying from about 

2 percent at M = 0.80 to as much as 8 percent at M = 1.20. This high degree of stability contributed to the lower trim angle-of-attack range at the higher transonic test 

Mach numbers for the center-vertical-tail model without engine. 

Static Lateral-Directional Stability 

The variations of the static lateral-directional stability with component buildup of 

the tip-fin configurations are presented in figure 11. The models tested without tip fins 

were unstable directionally but exhibited positive dihedral effect (-clP) at angles 

of attack above about 3' for all Mach numbers tested. The installation of the tip fins 

provided some improvement of the dihedral effect at low angles of attack and altered the 

directional stability in a nonsystematic manner with both angle of attack and Mach number. The directional stability was enhanced to positive values at low and high angles of 

attack for all Mach numbers but was variously altered with Mach number in the range of 

a! = 8O to 12O, from only slightly at M = 0.80 to neutral stability at M = 1.20. A 

decrease in directional stability at a! = 8' to 12' was observed at M = 0.20 (ref. 19) 

on this same configuration, but instability did not occur. The exact cause of this region 

of tip-fin ineffectiveness is not known, but it is probably a function of the vorticular flow 

emanating from the fuselage nose section and the wing leading edge. Before these tip fins 

can be used with confidence, this region of reduced effectiveness must be explored and the 

(-%) 

11 
design altered to eliminate the deficiency. The addition of the model scramjet engine to 

the BWVTFD model improved both the directional stability and positive dihedral effect. 

The variations of the static lateral-directional stability with component buildup of 

the center-tail configurations are presented in figure 12. The installation of the center 

tail greatly improved both the positive dihedral effect and the directional stability at all 

angles of attack and at all test Mach numbers. The positive dihedral effect increased 

with angle of attack for all Mach numbers. The directional stability remained positive 

for all Mach numbers but decreased with increasing angle of attack for the higher Mach 

numbers . 

A comparison of the lateral-directional stability characteristics of the tip-fin and 

center-tail configurations is presented in figure 13. The center tail not only provides a 

greater level of positive dihedral effect but a higher degree of directional stability than 

the tip fins at all angles of attack and at all test Mach numbers. The addition of the 

forward delta wing enhanced both stability derivatives. 

Roll Control 

The results of tests to determine the roll-control characteristics with 5O of differential aileron deflection are presented in figure 14 for the complete tip-fin configuration 

with engine. The data indicate that the configuration had positive roll control at all test 

conditions. However, some adverse yaw or cross coupling of yaw due to roll control 

occurred at the higher angles of attack for M = 0.80 and 0.90 at all angles of attack for 

M = 0.95 to 1.20. 

SUMMARY OF DATA WITH MACH NUMBER FROM 0.20 TO 6.0 

The present paper is the last of a series of reports that give details of wind-tunnel 

tests made on the 70' swept double-delta wing research airplane concept over the Mach 

number range of 0.20 to 6.0. These reports include subsonic studies (refs. 15 and 19), 

transonic studies (present report), supersonic studies (ref. 16), hypersonic studies 

(ref. 17), and hypersonic heat-transfer studies (refs. 20 and 21). A summary of the 

more important aerodynamic parameters as they vary with Mach number will be presented in this section. Summaries of data variation with Mach number for other research 

airplane configurations may be found in reference 18. 

Tests 

Reynolds number.- The variation of test Reynolds number with the test Mach number as used in the various wind tunnels is presented in figure 15. The widest range of 

test Reynolds numbers (based on fuselage length), 2.2 x lo6 to 19.7 x lo6, was made in 

12 
the Langley low-turbulence pressure tunnel at M = 2.0 with free transition. The 

majority of tests at M = 0.20 were made at a constant Reynolds number of 10 X lo6 

with fixed transition (ref. 19). Transonic tests were made with fixed transition at a constant stagnation pressure in the Langley 8-foot transonic pressure tunnel, thus accounting 

for variations in Reynolds number with increasing Mach number, as discussed in the section entitled "Test Conditions." Supersonic tests were made with fixed transition at 

M = 1.50 to 2.86 in the Langley Unitary Plan wind tunnel at a constant Reynolds number 

of 3.3 x lo6 (ref. 16). Hypersonic tests were made at M = 6 with free transition in the 

Langley 20-inch Mach 6 tunnel at a constant Reynolds number of 10.5 x 106 (ref. 17). The 

summary of data presented herein is plotted as received from the tunnels, with no corrections made for variations in Reynolds number or for grit drag for the fixed transition 

tests. 

Angles of attack for lift coefficient of 0 and 0.2. - The variation of angle of attack 

with Mach number at lift coefficients of 0 and 0.2 is presented in figure 15. These data 

are shown for two typical configurations to give a better understanding of subsequent 

figures of longitudinal and lateral-directional stability presented at CL = 0 and 0.2. The 

value of CL = 0.2 was selected because it was the highest value of lift coefficient available for all the Mach numbers utilized in the test program. It should be noted that the 

angles of attack for CL = 0 varied by about *2O from M = 0.20 to 6.0 and that angle of 

attack for CL = 0.2 varied from about 8O for the body-wing (BW) model to 8.5' for the 

complete BWV F E model. This relatively wide spread in angle of attack to obtain 

CL = 0.2 is an illustration of the decrease in lifting efficiency that occurs with increasing 

Mach number. 

( TD) 

Buildup of Tip-Fin Models 

Minimum drag.- The minimum drag coefficients of the various configurations used 

in the buildup of the tip-fin model are presented in figure 16 for the various test Mach 

numbers. It should be noted that all of the faired data have been corrected for base 

pressure and that only the flagged symbols at M = 0.20 are uncorrected. The minimum 

drag coefficients at the low Mach number of 0.20 increased to a maximum at M = 1 and 

are about the same as those at M = 6. The addition of the forward delta wing made only 

minor drag variations, but the tip fins increased the minimum drag as much as 78 percent 

at low speeds. The values of minimum drag coefficient shown without base pressure 

correction are higher by as much as 83 percent and, when corrected for Reynolds number, are considered as good engineering estimates of full-scale flight values at low subsonic Mach numbers. This large base-drag contribution illustrates the purpose of 

reducing base area to a minimum, particularly on aircraft that operate subsonically. 

13 
Maximum lift-drag ~" ~~~ ratio.- The maximum lift-drag ratio for the various tip-fin 

buildup models is presented in figure 16 over the Mach number test range. The maximum 

lift-drag ratio decreases rapidly with higher subsonic and transonic speeds but, once 

supersonic, appears to level out and remain above 3 up to M = 6. Each addition to the 

basic (BW) configuration decreased the lift-drag ratio substantially at subsonic speed, 

but the forward delta wing has considerably less effect than the tip fins at supersonic and 

hypersonic speeds. 

Pitching moment at zero lift.- The variation of pitching-moment coefficient with 

Mach number at zero lift is presented in figure 17 for the tip-fin model buildup. The 

c,O 

of all configurations increase up to M = 1 and exhibit the undesirable characteristics of negative at low speeds, with the tip-fin models having the most negative 

values. The tip fins tend to act as end plates and increase the local lift coefficients near 

the wing trailing edge, thus contributing to the negative pitching moments. The Cmo 

values for all models at all speeds above about M = 0.95 exhibit positive Cmo as 

desired. The Cmo is adversely affected by the addition of the tip fins at low speeds but 

is favorably affected at all supersonic speeds. The addition of the forward delta wing 

produces a desirable positive increment in at all test speeds. 

cmO 

cmO 

Lift-curve slope.- The variation of the lift-curve slope with Mach number for the 

tip-fin model buildup is presented in figure 17. There was a general decrease in 

with Mach number for all configurations. The addition of the tip fins had a marked 

positive effect on the C at low speeds but little effect at supersonic and hypersonic 

speed. The addition of the forward delta wing in conjunction with the tip fins produced 

the highest subsonic lift-curve slope. 

cL, 

La 

Longitudinal stability.- The variation of longitudinal stability with Mach number of 

the buildup of the tip-fin models is presented in figure 18 for lift coefficients of 0 and 0.2. 

A general trend for all configurations is a sharp increase in stability from low subsonic 

speeds to M = 1 and a more gradual decrease in stability with increasing Mach number. 

The stability may be expected to become neutral or negative at a sufficiently high Mach 

number for any given fixed configuration. The addition of the tip fins increased the longitudinal stability at all Mach numbers and lift coefficients. As expected, the addition of 

the forward delta wing deteriorated the longitudinal stability at all Mach numbers for all 

configurations tested. 

~ .. 

Lateral-directional stability. - The variation of lateral-directional stability with 

Mach number of the tip-fin buildup configurations is presented in figure 19. The configurations without tip fins were directionally unstable at all Mach numbers at both lift coefficients and exhibited negative dihedral effect at CL = 0 through a range of Mach numbers 

from 0 to 1.20 and from 2.86 to 6.0. The addition of tip fins provided positive directional 

~ . ~~~____ 

14 
stability for both configurations at Mach numbers up to at least 2.36 at CL = 0 and 0.2, 

and it also provided positive dihedral effect below M = 2.86 at CL = 0. The lateraldirectional stability was unaffected by the addition of the forward delta wing. 

Effect of Addition of Model Scramjet Engine 

Minimum drag.- The variation of the minimum drag with Mach number due to the 

addition of the model scramjet engine is presented in figure 20. At low speeds, the minimum drag was increased more than 100 percent, excluding base pressure, and increased 

about 64 percent when the base pressure is included. The increment of minimum drag 

coefficient was relatively constant for speeds up to about M = 1 (as discussed in 

ref. 18 for other research airplane configurations), but gradually decreases thereafter as Mach number increases. 

Maximum - lift-drag ratio.- The losses in maximum lift-drag ratio due to the installation of the scramjet engine are presented in figure 20. The largest loss occurred at low 

subsonic speeds and where decreases in L/D could be least tolerated (ref. 18). The 

incremental reduction in (L/D)max with Mach number due to engine addition decreases 

to only about 0.3 at cruising speed. It may be seen that base pressure further contributes 

to the low subsonic (L/D)max ratios. 

- "- ~~ 

Pitching moment at zero lift.- The effect of the addition of the model scramjet 

engine on the pitching moment at zero lift is presented in figure 21. There is a beneficial 

effect up to M = 2.36 and only a small undesirable decrease in positive at Mach 

numbers up to 6.0. 

.~ ~-. 

cmO 

Lift-curve slope.- The variation of the slope of the lift curve at zero lift with the 

installation of the model scramjet engine is presented in figure 21. A loss in C was 

experienced at subsonic speeds, but a gain was measured at the design cruise Mach number of 6.0. 

LCY 

Longitudinal stability.- The effects of the addition of the model scramjet engine on 

the longitudinal stability variation with Mach number are presented in figure 22 at lift 

coefficients of 0 and 0.2. No variations may be seen that exceed 0.01 except at transonic 

speeds, and there the decrease in stability was desirable. 

Lateral-directional stability.- The results of tests to determine the variations in 

lateral-directional stability with the installation of the model scramjet engine are presented in figure 23. The addition of the model engine was either favorable or nondetrimental on the lateral-directional stability at CL = 0 and 0.2 except at M = 6.0 for 

CL = 0. 

15 
Buildup of Center-Vertical-Tail Models 

Minimum drag.- The variation of the minimum drag with model buildup of the 

center-vertical-tail configurations is shown in figure 24. The positive increment in 

drag due to the center-vertical-tail addition was nearly constant with Mach number, with 

the base pressure corrected to free-stream static pressure. It should be noted that the 

base area of the center vertical tail was not included in this correction. The addition 

of the forward delta wing had only a very small effect on the minimum drag coefficients. 

Maximum lift-drag ~- ratio.- The largest loss of maximum L/D was at low subsonic 

speeds with the addition of the center vertical tail (fig. 24). The wedge airfoil section 

contributed to the large losses in (L/D)max and would not normally be used at subsonic 

or low supersonic speeds. The addition of the forward delta wing made a favorable contribution to (L/D)max only at M > 3. 

Pitching moment at zero lift.- Both the center vertical tail and the forward delta wing 

contributed to the positive increases of 

negative for the complete configuration and may compromise the pitch control and 

~ ~ .. 

cmO (fig. 25). The low-speed C m0 remained 

increase the trim drag. 

Lift-curve slope.- The variation of the lift-curve slope with Mach number at zero 

lift for the buildup of the center-vertical-tail model is shown in figure 25. The addition 

of the center vertical tail and the forward delta wing made only small nonsystematic 

changes in the variation of lift-curve slope with Mach number. 

Longitudinal stability.- The variation of longitudinal stability with Mach number for 

the configuration buildup on the center-vertical-tail configuration is presented in figure 26. With the exception of the increase in stability at M = 0.20 for CL = 0, the 

addition of the center vertical tail either made little change in the longitudinal stability or decreased it, as expected. The installation of the forward delta wing decreased 

the stability as expected at both lift coefficients throughout the Mach number test range. 

Lateral-directional stability.- The addition of the center vertical tail had a major 

positive effect on both the directional stability and the dihedral effect (fig. 27). The 

directional stability parameter was increased from unstable negative values to cnL3 

stable positive values at all Mach numbers and lift coefficients except for M = 6 at 

CL = 0.2. The positive dihedral effect was increased at all conditions. The 

installation of the forward delta wing improved the directional stability at the higher 

speeds for the higher lift coefficient, thereby showing more effect in conjunction with the 

center vertical tail Vc than with the tip fins VT. 

(-“d 

16 
Theoretical Estimates of Configuration Aerodynamics 

Estimates are presented of currently available computerized theories of performance and static stability parameters on the wing-body-tip-fin WBVT configuration 

through the Mach number test range from M = 0.20 to 6.0. Estimates were calculated 

at M = 0.20 by the vortex-lattice method described in references 27 and 28, with the 

base-pressure drag coefficient being made by methods outlined by Hoerner in reference 29. Estimates at M = 1.20 to 3.0 were made by the linear-theory methods presented in references 30 and 31, and at M = 3.0 to 6.0 by the Gentry Hypersonic Arbitrary 

Body Program (GHABP) of reference 32. Skin friction was calculated by the Spalding 

and Chi method of reference 33 with the test Reynolds numbers of figure 15, and boundary 

layers were assumed all turbulent. Other examples of the application of these theories on 

similar configurations may be seen in references 12, 18, and 34 to 38. 

0 

Estimates of the minimum drag coefficient and the maximum lift-drag ratio on the 

WBVT model are presented in figure 28 for the test Mach number range. The semiempirical values of the calculated minimum drag coefficient presented at M = 0.20 

underestimate the experiment data by a greater percentage than those calculated for 

supersonic and hypersonic speeds by high-speed computer programs. The estimated 

drag coefficients at M = 0.20 combined the turbulent flat-plate skin-friction coefficients 

modified by empirical form factors for body fineness ratios and wing thickness ratios 

with the base pressure coefficients determined from the empirical correlations with the 

body skin friction of reference 29. Possible sources of error were that no component 

interference effects were included and that no estimate of possible separated flows was 

made over the lower model afterbody. 

The viscid estimates of minimum drag coefficient are satisfactory for Mach numbers of 2.0 and above, but those at transonic speed are low by about 20 percent. The discontinuity in the viscid curve for M = 1 to 1.20 is due to the sharp increase in test 

Reynolds number (fig. 15). 

A comparison of estimates of the maximum lift-drag ratio on the WBVT model with 

experimental data in figure 28 shows adequate predictions except at high subsonic Mach 

numbers. The estimates at Mach numbers up to 3.0 were made by using linear theory 

with the predicted values of minimum drag coefficient and lift-curve slope. For Mach 

numbers from 3.0 to 6.0, the GHABP was used. 

A comparison of measured lift coefficients and lift-curve slopes with calculated 

estimates is presented in figure 29 for a = 0'. The theoretical model for the low-speed 

vortex-lattice calculations and the supersonic linear-theory calculations was an uncambered flat surface; thus, a CL value of zero resulted at Q = 0'. The slope of the lift 

curve, however, was well predicted with this simplified theoretical model. The GHABP, 

17 
utilizing a three-dimensional input model, predicted the lift coefficient poorly for a! = 0 

at M = 3, but predicted well at M = 6. The lift-curve slope was well predicted by the 

GHABP. 

0 

The calculated static longitudinal and directional stability parameters at zero lift 

are compared in figure 30 with experimental data for the WBV, configuration. The predictions of the longitudinal stability by the vortex-lattice method and the GHABP are good 

at low subsonic and hypersonic speeds, respectively, but are less than desirable by linear 

theory at supersonic speeds. The linear theory conservatively predicted the approximate 

level of positive longitudinal stability but showed a fortuitous trend with Mach number. 

This trend is partly due to the effects of the tip fins and the low body fineness ratio, 

neither of which was accounted for by the linear theory. 

A comparison of the static directional stability with the GHABP is presented in figure 30. For this plot, the directional stability parameter aCn/aCy was referenced to 

body length by the factor b/P to make it parallel to the previous discussion of longitudinal data. No present vortex-lattice or linear-theory method addresses the prediction 

of directional stability; thus, only the prediction from the GHABP is shown. It may be 

seen that the hypersonic theory overestimates the experimental data by a substantial 

amount, due in part to the inability of the program to take into account local variations 

in dynamic pressure and flow direction. 

It should be noted that none of the prediction methods discussed herein give completely satisfactory estimates of the aerodynamic characteristics of configurations with 

the model scramjet engine installed. These inadequacies are due in part to the inability 

of the programs to handle internal flows with the converging walls, the external spillage of 

flow not swallowed by the inlet, and the resulting interference effects. 

It may be concluded that good preliminary predictions of performance and static 

longitudinal stability can be expected across the speed range with the vortex-lattice, 

linear, and hypersonic theories on configurations with body fineness ratios of about 7. 

Predictions of directional stability are beyond the scope of these methods at subsonic and 

supersonic speeds and are grossly optimistic at hypersonic speeds. 

CONCLUSIONS 

An analysis of the experimental data for a hypersonic research airplane concept 

having a 70' double-delta wing at Mach numbers from 0.20 to 6.0 for a range of Reynolds 

numbers (based on fuselage length) from about 3.3 X lo6 to 10.5 X lo6 leads to the following 

conclusions: 

18 

. ". 
1. Models equipped with tip fins had more linear lift curves than models with the 

center vertical tail. 

2. Models equipped with the center vertical tail showed less tendency to pitchup 

and had a higher degree of directional stability and positive dihedral effect at higher 

angles of attack than models with tip fins. 

3. At transonic speeds, a region of directional instability that was not present for 

the center-tail models was observed at angles of attack from 8' to 12' for the models 

equipped with tip fins. 

4. The addition of the forward delta wing made only small changes in lift and drag 

but decreased maximum lift-drag ratio. 

5. The forward delta wing decreased longitudinal stability as expected but enhanced 

the positive dihedral effect. 

6. A positive increase in minimum drag coefficient, zero-lift pitching-moment 

coefficient, and zero-lift longitudinal stability was observed for all configurations with 

Mach number up to about 1. 

7. The tip fins made signficant increases in lift and drag to Mach 1.20 but resulted 

in a net loss in maximum lift-drag ratio at all speeds, due in part to the large fixed toe-in 

angle. 

8. Excluding base drag, the highest maximum lift-drag ratio for any one configuration 

occurred at the minimum subsonic test Mach number of 0.20 and the lowest lift-drag ratio 

at the hypersonic Mach number of 6.0. 

9. The installation of the model scramjet engine either had little effect or increased 

directional stability and positive dihedral effect through the Mach number test range. 

10. Good preliminary predictions of performance and static longitudinal stability can 

be expected across the speed range by use of the vortex-lattice, linear, and hypersonic 

theories on configurations with body fineness ratios of about 7. 

11. Predictions of directional stability are beyond the scope of these methods at subsonic and supersonic speeds and are grossly optimistic at hypersonic speed. 

Langley Research Center 

National Aeronautics and Space Administration 

Hampton, VA 23665 

September 20, 1979 

19 
REFERENCES 

1. Wetmore, Warren C.: Concorde Impressive in Flights. Aviat. Week & Space 

Technol., vol. 100, no. 25, June 24, 1974, pp. 25-28. 

2. SR-71 Sets Record. Aviat. Week & Space Technol., vol. 101, no. 10, Sept. 9, 1974, 

p. 17. 

3. SR-71 Sets Mark in New Speed Category: London-Los Angeles. Aviat. Week & Space 

Technol., vol. 101, no. 12, Sept. 23, 1974, p. 28. 

4. Kirkham, Frank S.; Jackson, L. Robert; and Weidner, John P.: The Case for a HighSpeed Research Airplane - Results From an In-House Study. AIAA Paper 

NO. 74-988, Aw. 1974. 

5. Small, W. J.; Fetterman, D. E.; and Bonner, T. F., Jr.: Potential of Hydrogen Fuel 

for Future Air Transportation Systems. [Preprint] 73-ICT-104, American SOC. 

Mech. Eng., Sept. 1973. 

6. Hypersonic Research Facilities Study. Volume 11, Part 1, Phase I: Preliminary 

Studies - Research Requirements and Ground Facility Synthesis. NASA 

CR-114323, 1970. 

7. Hypersonic Research Facilities Study. Volume 11, Part 2, Phase I: Preliminary 

Studies - Flight Vehicle Synthesis. NASA CR-114324, 1970. 

8. Hypersonic Research Facilities Study. Volume 111, Part 2, Phase 11: Parametric 

Studies - Flight Vehicle Synthesis. NASA CR-114326, 1970. 

9. Hypersonic Research Facilities Study. Volume IV, Part 1, Phase III: Final Studies - 

Flight Research Facilities. NASA CR-114327, 1970. 

10. Becker, John V.: The X-15 Program in Retrospect. Raumfahrlforschung, Bd. XIII, 

Heft 2, Mar./Apr. 1969, pp. 45-53. 

11, Clark, Louis E.: Hypersonic Aerodynamic Characteristics of an All-Body Research 

Aircraft Configuration. NASA TN D-7358, 1973. 

12. Penland, Jim A.; Creel, Theodore R., Jr.; and Howard, Floyd G.: Experimental LowSpeed and Calculated High-speed Aerodynamic Characteristics of a Hypersonic 

Research Airplane Concept Having a 65' Swept Delta Wing. NASA TN D-7633, 

1974. 

13. Penland, Jim A.; and Creel, Theodore, R., Jr.: Low-Speed Aerodynamic Characteristics of a Lifting-Body Hypersonic Research Aircraft Configuration. NASA 

TN D-7851, 1975. 

20 

I II 
14. Lawing, Pierce L.; and Hunt, James L.: Aerodynamic Heat Transfer to a Hypersonic 

Research Aircraft Model (X-24C) at Mach 6. J. Aircr., vol. 13, no. 12, Dec. 1976, 

pp. 1018-1020. 

15. Creel, Theodore R., Jr.; and Penland, Jim A.: Low-Speed Aerodynamic Characteristics of a Hypersonic Research Airplane Concept Having a 70' Swept Delta Wing. 

NASA TM X-71974, 1974. 

16. Penland, Jim A.; Fournier, Roger H.; and Marcum, Don C., Jr.: Aerodynamic Characteristics of a Hypersonic Research Airplane Concept Having a 70' Swept DoubleDelta Wing at Mach Numbers From 1.50 to 2.86. NASA TN D-8065, 1975. 

17. Clark, Louis E.; and Richie, Christine B.: Aerodynamic Characteristics at Mach 6 

of a Hypersonic Research Airplane Concept Having a 70' Swept Delta Wing. NASA 

TM X-3475, 1977. 

18. Penland, Jim A.; Dillon, James L.; and Pittman, Jimmy L.: An Aerodynamic Analysis 

of Several Hypersonic Research Airplane Concepts From M = 0.2 to 6.0. J. Aircr., 

vol. 15, no. 11, Nov. 1978, pp. 716-723. 

19. Penland, Jim A.; Creel, Theodore R., Jr.; and Dillon, James L.: Aerodynamic Characteristics of a Hypersonic Research Airplane Concept Having a 70' Swept DoubleDelta Wing at Mach Number 0.2. NASA TP-1252, 1978. 

20. Johnson, Charles B.; Taylor, Allan H.; and Weinstein, Irving: Heat-Transfer and 

Pressure Measurements on a Simulated Elevon Deflected 30' Near Flight Conditions 

at Mach 7. NASA TM X-3563, 1977. 

21. Lawing, Pierce L.: Configuration Heating for a Hypersonic Research Airplane Concept Having a 70' Swept Double-Delta Wing. NASA TP-1143, 1978. 

22. Henry, John R.; and Anderson, Griffin Y.: Design Considerations for the AirframeIntegrated Scramjet. NASA TM X-2895, 1973. 

23. Lawing, Pierce L.; and Johnson, Charles B.: Inlet Boundary-Layer Shapes on Four 

Aircraft Forebodies at Mach 6. J. Aircr., vol. 15, no. 1, Jan. 1978, pp. 62-63. 

24. Johnson, Charles B.; and Lawing, Pierce L.: Mach 6 Flowfield Survey at the Engine 

Inlet of a Research Airplane. J. Aircr., vol. 14, no. 4, Apr. 1977, pp. 412-414. 

25. Braslow, Albert L.; and Knox, Eugene C.: Simplified Method for Determination of 

Critical Height of Distributed Roughness Particles for Boundary-Layer Transition 

at Mach Numbers From 0 to 5. NACA TN 4363, 1958. 

26. Blackwell, James A., Jr.: Preliminary Study of Effects of Reynolds Number and 

Boundary-Layer Transition Location on Shock-Induced Separation. NASA 

TN D-5003, 1969. 

21 
27. Lamar, John E.; and Gloss, Blair B.: Subsonic Aerodynamic Characteristics of 

Interacting Lifting Surfaces With Separated Flow Around Sharp Edges Predicted 

by a Vortex-Lattice Method. NASA TN D-7921, 1975. 

28. Polhamus, Edward C.: A Concept of the Vortex Lift of Sharp-Edge Delta Wings 

Based on a Leading-Edge-Suction Analogy. NASA TN D-3767, 1966. 

29. Hoerner, Sighard F.: Fluid-Dynamic Drag. Publ. by the author (148 Busteed Drive, 

Midland Park, New Jersey 07432), 1965. 

30. Middleton, Wilbur D.; and Carlson, Harry W.: Numerical Method of Estimating and 

Optimizing Supersonic Aerodynamic Characteristics of Arbitrary Planform Wings. 

J. Aircr., vol. 2, no. 4, July-Aug. 1965. 

31. Harris, Roy V., Jr.: An Analysis and Correlation of Aircraft Wave Drag. NASA 

TM X-947, 1964. 

32. Gentry, Arvel E.: Hypersonic Arbitrary-Body Aerodynamic Computer Program 

(Mark 111 Version). Vol. I - User's Manual. Rep. DAC 61552, Vol. I (Air Force 

Contract Nos. F33615 67 C 1008 and F33615 67 C 1602), McDonnell Douglas Corp., 

Apr. 1968. (Available from DDC as AD 851 811.) 

33. Spalding, D. B.; and Chi, S. W.: The Drag of a Compressible Turbulent Boundary 

Layer on a Smooth Flat Plate With and Without Heat Transfer. J. Fluid Mech., 

vol. 18, pt. 1, Jan. 1964, pp. 117-143. 

34. Pittman, Jimmy L.; and Dillon, James L.: Vortex Lattice Prediction of Subsonic 

Aerodynamics of Hypersonic Vehicle Concepts. J. Aircr., vol. 14, no. 10, Oct. 

1977, pp. 1017-1018. 

35. Pittman, Jimmy L. (appendix by C. L. W. Edwards): Application of Supersonic 

Linear Theory and Hypersonic Impact Methods to Three Nonslender Hypersonic 

Airplane Concepts at Mach Numbers From 1.10 to 2.86. NASA TP-1539, 1979. 

36. Dillon, James L.; and Pittman, Jimmy 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. 

37. Dillon, James L.; and Creel, Theodore R., Jr.: Aerodynamic Characteristics at Mach 

Number 0.2 of a Wing-Body Concept for a Hypersonic Research Airplane. NASA 

TP-1189, 1978. 

38. Dillon, James L.; and Pittman, Jimmy L.: Aerodynamic Characteristics at Mach 6 

of a Wing-Body Concept for a Hypersonic Research Airplane. NASA TP-1249, 

1978. 

22 
TABLE I.- GEOMETRIC CHARACTERISTICS OF MODEL 

Wing: 

Reference area (includes area projected to fuselage center line). 

m2 (in2) ............................... 0.043 (67.200) 

Exposed area. m2 (in2) ......................... 0.023 (36.121) 

Wetted area. m2 (in2) .......................... 0.047 (72.242) 

Span. m (in.) .............................. 0.217 (8.542) 

Aspect ratio .................................... 1.086 

Root chord (on fuselage center line). m (in.) ............. 0.353 (13.896) 

Tip chord. m (in.) ........................... 0.085 (3.355) 

Taper ratio .................................... 0.241 

Mean aerodynamic chord. m (in.) ................... 0.248 (9.779) 

Sweepback angles. deg: 

Leading edge .................................. 70 

25-percent chord line .............................. 64 

Trailing edge................................. 0 

Dihedral angle (airfoil mean line). deg ...................... -3.64 

Incidence angle. deg ................................ 0 

Airfoil section (see fig . 3(a)): 

Thickness ratio - 

Exposed root ................................. 0.05 

Tip ...................................... 0.06 

Leading-edge radius. m (in.): 

Fuselage center-line chord ..................... 0.000508 (0.020) 

Tip ................................. 0.000508 (0.020) 

Elevon area (both). m2 (in2) ....................... 0.005 (7.161) 

Forward delta wing: 

Area exposed (outside of fuselage. forward of wing leading edge). 

m2 (in21 ................................ 0.002 (3.394) 

Leading-edge sweep. deg ............................. 80 

Wetted area (both). m2 (in2) ....................... 0.0044 (6.788) 

Tip fin (vertical tails): 

Area (each). m2 (in2) ........................... 0.0038 (5.848) 

Span. m (in.) ............................... 0.069 (2.730) 

Aspect ratio .................................... 1.274 

Root chord. m (in.) ............................ 0.086 (3.354) 

Tip chord. m (in.) ............................ 0.029 (1.135) 

Taper ratio .................................... 0.336 

Mean aerodynamic hord. m (in.) .................... 0.062 (2.445) 

Sweepback angles. deg: 

Leading edge. top ................................ 55.0 

Leading edge. bottom .............................. 70.1 

Trailing edge. top ................................ 21.3 

23 
TABLE I.- Concluded 

Toe-in angle. deg ................................. 7.5 

Airfoil section: 

Thickness ratio ................................ 0.067 

Leading-edge radius. m (in.) ................... 0.000508 (0.020) 

Area (exposed). m2 (in2) ........................ 0.007 (11.492) 

Span (exposed). m (in.) ......................... 0.086 (3.380) 

Aspect ratio of exposed area ........................... 0.994 

Root chord (fuselage surface line). m (in.) ............... 0.128 (5.040) 

Tip chord. m (in.) ............................ 0.045 (1.760) 

Mean aerodynamic chord of exposed area. m (in.) ........... 0.093 (3.264) 

Leading edge .................................. 55.0 

Trailing edge .................................. 24.6 

Center fin (vertical tail): 

Taper ratio .................................... 0.349 

Sweepback angles. deg: 

Airfoil section: 

Thickness ratio - 

Tip ..................................... 0.106 

Root ..................................... 0.106 

Leading-edge radius. m (in.) .................... 0.000508 (0.020) 

Length. m (in.) ............................ 0.508 (20.000) 

Maximum height. m (in.) ....................... 0.071 (2.782) 

Maximum width. m (in.) ....................... 0.073 (2.866) 

Fineness ratio of equivalent round body ..................... 6.822 

Planform area. m2 (in2) ....................... 0.026 (40.445) 

Wetted area. m2 (in2) ......................... 0.083 (128.460) 

Wetted area (with wing on). m2 (in2) ................. 0.078 (120.695) 

Wetted area (with wing and forward delta on). m2 (in2) ........ 0.077 (118.747) 

Base area. m2 (in2) .......................... 0.002 (3.726) 

Planform area. m2 (in2) ....................... 0.052 (79.960) 

Aspect ratio of planform ............................. 0.913 

Planform area (without forward delta). m2 (in2) ........... 0.049 (76.566) 

Aspect ratio (without forward delta) ....................... 0.953 

Frontal area. m2 (in2) ........................ 0.00084 (1.308) 

Width-to-height ratio .............................. 5.23 

Fuselage: 

Complete model (wing and forward delta): 

Model scramjet engine: 

24 
X 

Figure 1.- Systems of reference axes; arrows indicate positive direction. 
(a) Configuration model. 

Figure 2.- Configuration and sketch showing interchangeable parts. 
(b) Sketch showing interchangeable parts. 

Figure 2. - Concluded. 

27 
Airfoil section at wing tip 

I 

- 

Vertical tip-fin 

airfoil section 

(a) Base-line configuration. 

Figure 3.- Model general dimensions. All dimensions have been normalized by the body length (e = 50.8 cm). 
Station .839 

(b) Tip fin. 

I 

- 

I .0266 - 

Statim. 748 t 

(c) Center vertical tail. 

Figure 3. - Continued. 

29 
E Section A-A 

GI1016 

Inside splitter plate 

?IL Station. 625 -x 

Bottom view 01 scramjet engine with cowl on 

I 

,0278 

15625 Bonom view 01 scramjet engine with cowl removed 

Len Outside plate 

(d) Scramjet engine. 

Figure 3.- Concluded. 

30 
,040 

-035 

.030 

.025 

.015 

. 010 

.005 

0 

-4 0 4 

0 M = 0.80 

0 0.90 

0 0.95 

0.98 

@. 1.10 

Is 1.20 

8 

(a) BWVTFDE. 

n 

24 

Figure 4.- Variation of base axial-force coefficient with angle of attack 

for various test Mach numbers for complete configuration with and 

without engine. 

31 
.04c 

.035 

.03C 

.025 

.015 

.010 

.005 

fl 

u B"""~F~ 

0 M = 0.80 

0 0.90 

0 0.95 

A 0.98 

b 1.10 

~ 1.20 

12 2 

Figure 4.- Concluded. 

32 
24 

22 

20 

18 

16 

14 

IO 

8 

6 

4 

2 

0 

-2 

cL 

(a) M = 0.80. 

Figure 5. - Longitudinal aerodynamic characteristics of body-wing configuration 

alone and with various forward delta, tip-fin, and engine components. 

33 
uo 

CL 

(a) Concluded. 

Figure 5.- Continued. 

34 
4 

CI 

(b) M = 0.90. 

Figure 5.- Continued. 

35 
(b) Concluded. 

Figure 5.- Continued. 

36 
.M 

e 

cm 

-. M 

-. 08 

2 

1 

1 

1 

1 

1 

u. deg 

0 

a6 

4 

2 

0 

8 

6 

4 

2 

-. 1 .1 

0 BW 

Figure 5. - Continued. 

37 
L 10 

CL 

(c) Concluded. 

Figure 5.- Continued. 

38 
a. deg 

0 BW 

0 BWFD 

0 BWVT 

BWVTFD 

BWVTFDE 

.5 .r .a 

(d) M = 0.98. 

Figure 5. - Continued. 

39 
I111111 I 

0 

Tr 

+ 

! 

I 

I 

0 BW 

.6 L 

(d) Concluded. 

Figure 5.- Continued. 

40 
-. 2 -. 1 0 .1 .3 .4 .5 

cL 

.6 .7 

(e) M = 1.10. 

Figure 5.- Continued. 

41 
111 11 II I I 

. ZLI 

.24 

.M 

.16 

cD 

.I2 

.08 

.&I 

! 

42 

(e) Concluded. 

Figure 5.- Continued. 
21 

12 

16 

14 

li 

1t 

0. dql E 

t 

4 

i 

C 

-2 

-4 -. 1 0 .1 .2 .3 

cL 

.I 

(f) M = 1.20. 

Figure 5.- Continued. 

43 
LID 

.24 

.M 

.16 

'D .12 

.oE 

.w 

0 -. 1 0 .2 .3 .4 

'L 

I 

t 

f 

i 

I 

(f) Concluded. 

Figure 5.- Concluded. 

44 
2, 

2: 

21 

It 

If 

14 

4 *q 12 

IO 

a 

6 

4 

2 

0 

-2 

-4 

0 BW 

0 BWFD 

0 BWc 

A BWCFD 

(a) M = 0.80. 

Figure 6.- Longitudinal aerodynamic characteristics of body-wing configuration 

alone and with various forward delta and center-tail components. 

45 
Ct 

(a) Concluded. 

Figure 6. - Continued. 

46 
2; 

21 

IS 

16 

14 

12 

4 de5 

IO 

8 

6 

4 

2 

0 

-2 

-4 

cL 

(b) M = 0.90. 

Figure 6.- Continued. 

47 
L 

(b) Concluded. 

Figure 6. - Continued. 

48 
. . . 

.w- - -- - 

cL 

(c) M = 0.95. 

Figure 6.- Continued. 

49 
50 
21 

It 

It 

14 

12 

4 dql IO 

8 

6 

4 

2 

0 

-2 

-4. -. I 0 

0 BW 

0 BWFD 

0 BWVc 

A BWVCFo 

.L 

(d) M = 0.98. 

Figure 6. - Continued. 

.I 

51 
-. 2 -. 1 0 .1 .3 .4 .5 .6 .7 

cL 

(d) Concluded. 

Figure 6. - Continued. 

52 
-. 2 -. I 0 .I 2 .4 .5 

cL 

, .7 

(e) M = 1.10. 

Figure 6. - Continued. 

53 
, .. . . . __ ." . .. ~. 

0 BW 

0 BWFD 

0 BWVc 

A BWVcFD 

.2 .3 .6 7 

cL 

(e) Concluded. 

Figure 6. - Continued. 

54 
18 

16 

14 

12 

4 deg IO 

8 

6 

4 

2 

0 

-2 

-4 -. I 0 , .I .2 .3 " 

cL 

(f) M = 1.20. 

Figure 6.- Continued. 

55 
LID 

4 

3 

2 

1 

0 

-1 

-2 

-3 

-4 

.21 

.2r 

.2( 

. 16 

cD 

.12 

.08 

.D4 

0 

3- 

I1- 

I- 

- 

2 

56 
a 

(a) M = 0.80. 

Figure 7.- Longitudinal aerodynamic characteristics of tip-fin and center-tail 

configurations with and without forward delta. 

57 
(a) Concluded. 

Figure 7.- Continued. 

58 
22 

20 

18 

16 

14 

12 

4 deg 

10 

8 

6 

4 

2 

0 

-2 

4 -. I 

cL 

(b) M = 0.90. 

Figure 7.- Continued. 

59 
60 
r 

.M 

cm O 

-. M 

-. 08 

.I .2 .4 .5 

cL 

(c) M = 0.95. 

Figure 7.- Continued. 

61 
-. 2 -. 1 

CL 

(c) Concluded. 

Figure 7.- Continued. 

62 
.M 

cm 0 

-. M 

-. 08 

20 

18 

16 

14 

I2 

4 deg IO 

8 

6 

4 

2 

0 

-2 

-4 -. I 0 .I .L .4 .7 

(d) M = 0.98. 

Figure 7.- Continued. 

63 
4 

? 

i 

1 

L/D 

[ 

-1 

7 

-9 

.24 

.2c 

. 16 

'D .12 

.of 

.D4 

c 

T 

T 

! 

c 

,,, 

!I! 

0 BWVT 

BWV~ 

0 

A 

i ? 

I... 

,... 

L 

U .6 7 .a -. 1 .1 

'L 

(d) Concluded. 

Figure 7.- Continued. 

64 
.w 

cm 0 

-. 04 

-. 08 

.21 

18 

16 

14 

12 

4 deg 10 

8 

6 

4 

2 

0 

-2 

-4 

2 -. I 

(e) M = 1.10. 

Figure 7.- Continued. 

65 
C/ D 

(e) Concluded. 

Figure 7. - Continued. 

66 

" 
t 20 

18 

16 

14 

12 

4 deg IO 

8 

6 

4 

2 

0 

-2 

-A 

-. I 0 

0 BWV, 

(f) M = 1.20. 

Figure 7.- Continued. 

.5 

67 
4 

? 

i 

1 

-1 

- 

t + 

I 

! 

t 

I 

t 

! 

I 

c 

+ 

t 

i 

i 

.7 

BwT 

BwC 

BwCFD 

Bw~F~ 

I 

.6 - 0 1 c 

LL 

(f) Concluded. 

Figure 7. - Concluded. 

68 
.7 .0 .9 

CL 

(a) M = 0.80. 

Figure 8. - Elevon effects on longitudinal aerodynamic characteristics 

of BWVTFDE configuration. 

69 
LID 

CL 

(a) Concluded. 

Figure 8.- Continued. 

70 
a. 

*-. 

-.4 -.3 -.2 :1 0 .2 .3 .4 .5 .6 .7 .8 

CL 

(b) M = 0.90. 

Figure 8. - Continued. 

71 
LID 

(b) Concluded. 

Figure 8.- Continued. 

72 
a. dql 

(c) M = 0.95. 

Figure 8.- Continued. 

73 
LID 

cL 

(c) Concluded. 

Figure 8. - Continued. 

74 

.a 
a. 

CL 

(d) M = 0.98. 

Figure 8.- Continued. 

75 
LID 

76 

CL 

(d) Concluded. 

Figure 8. - Continued. 
a. 

CL 

(e) M = 1.10. 

Figure 8.- Continued. 

77 
[ 

LID 

-1 

-2 

-2 

-4 

.32 

.2f 

.24 

.M 

CD .1c 

.12 

.Pa 

i ~ I.M 

i 

0' -. 2 -.I 

(e) Concluded. 

Figure 8. - Continued. 

78 
(f) M = 1.20. 

Figure 8.- Continued. 

79 
3 

2 

1 

LID 0 

-1 

-. 2 -. 1 

Figure 8.- Concluded. 

80 
CL 

(a) M = 0.80. 

Figure 9.- Elevon effects on longitudinal aerodynamic characteristics 

of BWVCFD configuration. 

81 
LID 

(a) Concluded. 

Figure 9.- Continued. 

82 
CL 

(b) M = 0.90. 

Figure 9. - Continued. 

83 
LID 

84 
.. .. - -. . ... . . . . . . . . . ... . . 

I 

85 
86 

CL 

(c) Concluded. 

Figure 9.- Continued. 
-.4 -.3 -2 -.I 0 .2 .3 .4 .5 .6 .7 .8 

CL 

(d) M = 0.98. 

Figure 9. - Continued. 

87 
1 .ID 

CL 

(d) Concluded. 

Figure 9. - Continued. 

88 
CL 

(e) M = 1.10. 

Figure 9.- Continued. 
90 

LID 

CL 

(e) Concluded. 

Figure 9. - Continued. 
-. 1 0 .1 .2 .3 .4 .5 .6 

CL 

(f) M = 1.20. 

Figure 9.- Continued. 

91 
L ID 

.2 

cL 

(f) Concluded. 

Figure 9.- Concluded. 

92 

.3 .4 .5 .6 .7 
0 .1 L .3 .4 .5 

C 4 

(a) M = 0.80. 

-20 

-10 

0 

c 

'r 

Figure 10.- Longitudinal aerodynamic characteristics at trim of BWVTFDE 

and BWVCFD configurations. 
T! 

I 04 

02 

0 

t 

0 lli TI I 

!ll 

i li 

ft 

-. 004 I 

0 

02 

04 

06 

08 

-. 1 0 

I I 

i 

+ 

i 

i 

I .1 

t 

-t i 

t 4- 

c 

.2 .5 .6 .7 .8 

C 4 

(a) Concluded. 

Figure 10. - Continued. 

94 
.35 

.30 

.25 

20 

C 

Dt 

.15 

.10 

.05 1: 

0- 

0 .1 

BWVTFDE 

"_ BW 

.2 

T 

II TT 

.5 .6 

M = 0.90. 

.7 .8 

-- 0 

Figure 10.- Continued. 

95 

I. 
.06 

.04 

.02 

0 

1 

:I ft 

N I 

I 

r t t 

a i + 

.3 .4 

(b) Concluded. 

Figure 10. - Continued. 

96 
.35 

.30 

25 

.20 

C 

Dt 

.15 

.10 

.05 

/I 

'I 

I/ Dl .. 

.1 

ll 

1 I 

I 

i r 

I 

I 

I 

I 

13 

.2 .3 .4 .5 

(c) M = 0.95. 

F'igur e 10.- Continued. 

97 
.06 

.M 

0 

0 

-. 002 

-. 004 

t 

-. 006 

0 

(M 

04 

06 

08 

10 

0 .1 

(c) Concluded. 

Figure 10. - Continued. 

.2 .3 .4 .5 7 .8 

98 
3 

1 

0 

.3 .5 

(d) M = 0.98. 

Figure 10.- Continued. 

.4 

C 

Lt 

.6 .7 

99 
t 

0 

-. a2 

-. 04 

Mt-. 06 

-. 08 

-. 10 

0 .1 .2 

C 

Lt 

(d) Concluded. 

.6 .8 

Figure 10.- Continued. 

100 
-20 

-10 pg 

0 

.4c 

BWVTFDE 

BWVCFD .35 ;:! L 

.30 

.25 

I 

i .20 

Dt 

15 

051 

l11 

OO .I .4 .5 

C 

Lt 

.6 .? 

(e) M = 1.10. 

Figure 10.- Continued. 

101 
t 

-. 02 

-. 04, Ht -. 06 

-. 08 

-. 10 

0 .2 .3 .4 .5 .6 .7 

(e) Concluded. 

10.- Continued. 

102 
4 

3 

(LI Dlt 

2 

1 

0 

1 

i 

0 

.30 T 1 

25 

.20 

C 

Dt 

.15 

.10 

.05 

.4 

cLt 

M = 1.20. 

Figure 10.- Continued. 

103 
i r 

;,!!I" ;I 

0 ... 

.. 

.. 

,. 

....... ......... r ,, .-.., ........ 

,.I,,. . , . , , . . , ,. ,.. ...... 

........ 

........ 

........ ,,**#I, 

m: 

....... : ....... ......... .2*:: ........ 

::::::A: :::::::: . : , 

, I * .. , . , ........ ,..... ......... 1 .::" 

..... 

........ 

-. 006 

0 

-. 02 

jJ;i ........ 

I 

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

...... I. 

......... ,I 

........ 

....... 

...... 

........ 

........ .,,*.. I 

-. 04 .... 

. ,. .I 

..,. 

.. 

.. 

. .r 

.... Bt .... -. 06 

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

...... ..I. , 

...... 1 : ::x: 

:::A,:: 

z::.::: 

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

." 

.6 

...... ,*,, 

,, , I,,,.. 

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

...... ..... 

....... 

........ 

.... 

........ 

. , , . , , , , 

...... I,. I .. ....... 

:I I. .I. I::::: ..... ....... 

....... 

....... ....... :;I;:'':: ,, 

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

,, , I , 

......... 

......... 

. .,.I 

....... ....... -. 08 

......... 

......... 

,. ,,,, ,*, ......... 

.......... 

.. [ .... 

.L .3 .4 

C 

Lt 

(f) Concluded. 

.5 

Figure 10.- Concluded. 

104 
.02 

C 0 

-. 02 

cz P 

0 

-. 004 

-. 008 

0 BwT 

A BWVTFD 

h BWVTFDE 

-4 0 4 8 12 16 20 24 

a, deg 

(a) M = 0.80. 

Figure 11. - Lateral-directional stability characteristics of body-wing 

configuration alone and with various forward delta, tip-fin, and 

engine components. 

105 
Figure 11. - Continued. 

106 
a, deg 

(c) M = 0.95. 

Figure 11.- Continued. 

107 
108 

0 

Q 

-. 02 
C 

C 

I 

109 
0 

C 

Y P 

-. 02 

a, deg 

(f) M = 1.20. 

Figure 11.- Concluded. 

110 
4 

cz P 

-4 0 4 8 12 16 20 

4 deg 

(a) M = 0.80. 

Figure 12. - Lateral-directional stability characteristics of the body-wing 

configuration alone and with various forward delta wing and centertail components. 

111 
0 

cyP 

-. 02 

-. 04 

3 BW 

a, deg 

(b) M = 0.90. 

Figure 12.- Continued. 

112 
-4 0 4 8 12 16 20 

4 deg 

(c) M = 0.95. 

Figure 12. - Continued. 
0 

-. 02 P .. 

-. 04 

,004 

0 

004 

a, deg 

(d) M = 0.98. 

Figure 12.- Continued. 

114 
a, deg 

(e) M = 1.10. 

Figure 12. - Continued. 

115 
0 

C 

yp -. 02 

-. 04 

116 
.004 

0 

-. 004 

'! 

-. 008 

C 

"P 

9 deg 

(a) M = 0.80. 

Figure 13.- Comparison of the lateral-directional stability characteristics 

of tip-fin and center-tail configurations. 

117 
-4 0 4 8 12 16 20 24 

a, deg 

(b) M = 0.90. 

Figure 13. - Continued. 

118 
I 

0 

c Yp -.a 

-. 04 

-. 004 

0 

P 

-. 004 

" 008 

4 deg 

(c) M = 0.95. 

Figure 13.- Continued. 

! 119 
C 

-. 004 

0 

B 

-. 004 

-. 008 

. .. -4 0 4 8 12 16 20 

a, deg 

(d) M = 0.98. 

Figure 13. - Continued. 

120 
-4 0 4 8 12 16 LO 

a, deg 

(e) M = 1.10. 

Figure 13.- Continued. 

121 
0 

cyp -. 02 

__ 

-. 04 

0 

cz B 

-. 004 

-. 008 

-4 0 4 8 12 16 20 

a, deg 

(f) M = 1.20. 

Figure 13.- Concluded. 

122 
6H 

I/ I/ 

4 deg 

(a) M = 0.80. 

Figure 14. - Roll control of configuration BWVTFDE. 

123 
cy6 H 

C 

"% 

.002 

001 

C 

(b) M = 0.90. 

Figure 14.- Continued. 

124 
. 002' 

.Ool 0. r . I 1 . 002i. 

-. 002- -4 

C I - 

8 12 24 0 4 

(c) M = 0.95. 

Figure 14. - Continued. 

125 
C 

ygH 

6H 

C 

n 

6H 

.002 

.001 

0 I 

4 

4 

Tlii 

IIII 

I 

I 

12 

II 

1 Ili" Ill 

IIIII 

4 I1 1111 

I/ 1/11 

cx, deg 

(d) M = 0.98. 

Figure 14. - Continued. 

126 
4 

a, deg 

(e) M = 1.10. 

Y T 

8 12 16 

Figure 14.- Continued. 

127 
L002 

cyb H 0 

-. 02 

002 

001 

0 

C 

n 

dH 

- 

8 12 - 

a, deg 

(f) M = 1.20. 

Figure.14.- Concluded. 

16 i 

L!! i 

I 

n -- 24 

128 
20 x lo6 

16 

12 

R 

8 

4 

0 

16 

12 

8 

4 deg 

4 

c 

4 

0 Fixed transition 

0 Free transition 

I 

CL*. 2 

- 

0 

0 

CL .(I 

A 

M 

/ 

;::: BW 

BWVTFDE 

I 

3 4 5 6 

M 

Figure 15.- Variation of test Reynolds number and angle of attack 

with Mach number for lift coefficients of 0 and 0.2. 
/L 

M 

Figure 16.- Variation of minimum drag and maximum lift-drag ratio 

with Mach number on component buildup of tip-fin models. 

130 
.02 

. 01 

cm 0 

0 

-. 01 

-. 02 

.05 

.04 

.03 

cL 

a 

.02 

. 01 

0 1 

Figure 17.- Variation of values of pitching-moment coefficient with 

Mach number at zero lift and slope of lift curve at zero lift on 

component buildup of tip-fin models. 

131 
-. 12 

-. 08 

132 
-OM I 

-. ow 

.ow, 

-. 004 I 

. OM 

-. 004 

-. 004 

0 BW 

1 2 3 4 5 6 

M 

Figure 19.- Variation of lateral-directional stability with 

Mach number at lift coefficients of 0 and 0.2 on component buildup of tip-fin models. 

133 
M 

Figure 20.- Variation of minimum drag and maximum lift-drag 

ratio with Mach number on complete configuration with and 

without engine. 

134 

.... 
.04 

i; 

I 

.03 

C .02 

-01 

c 1 3 4 5 6 

M 

Figure 21.- Variation of values of pitching-moment coefficient with 

Mach number at zero lift and slope of lift curve at zero lift on 

complete configuration with and without engine. 

135 
-. 12 

-; 08 

k) CiQ2 "04 

0 

.04 

M 

Figure 22.- Variation of longitudinal stability with Mach number at lift coefficients of 

0 and 0.2 on complete configuration with and without engine. 

136 
1111 - 

i 

I 

I, 

-. 004 

i"lP) y0.2 

0 

M 

Figure 23.- Variation of lateral-directional stability with Mach number 

on complete configuration with and without engine. 

137 
8 

7 

6 

5 

4 

3 

2 

1 

0 

M 

6 

Figure 24.- Variation of minimum drag and maximum lift-drag ratio 

with Mach number on component buildup of center-tail models. 

138 
"_.._............ . , , .I ..I I I II 

.03 

C 02 

.01 

0 

-. 01 

.04 

.03 

C 

La .02 

. 01 

C 

1 2 3 4 5 6 

M 

Figure 25.- Variation of values of pitching-moment coefficient with Mach number 

at zero lift and slope of lift curve at zero lift on component buildup of centertail models. 

139 
-. 12 

-. 08 

k) 

0 

.04 

M 

Figure 26.- Variation of longitudinal stability with Mach number at lift coefficients 

of 0 and 0.2 on component buildup of center-tail models. 

140 
.M)8 

-. OM 1 3 4 

M 

Figure 27.- Variation of lateral-directional stability with 

Mach number on component buildup of center-tail 

models. 

I 

141 
.06 

.04 

'D, mi n 

.02 

0 

tGHABP 

Linear theory 

Vortex lattice 

"" 

"- ?# 'iii 

- I nviscid 

Flagged symbols include C 

'. 

A. b 

M 

Figure 28.- Comparison of theoretical estimates of minimum 

drag coefficient and maximum lift-drag ratio with experimental data at M = 0.20 to 6.0 for WBVT configuration. 

142 
04 

04 

08 

"" 

i 

M 

Figure 29.- Comparison of theoretical estimates of lift coefficient and 

lift-curve slope at CY = Oo with experimental data at M = 0.20 to 6.0 

for WBVT configuration. 

143 

n. . .. . . . .. . ." ." 
-. 12 

-. 08 k) ciao 

-. 04 

0 

-. 16 

-. 12 

-. 08 

or00 

-. 04 

0 

.04 

M 

Figure 30.- Comparison of theoretical estimates of static longitudinal and 

directional stability at M = 0.20 to 6.0 for WBVT configuration. 

144 
NASA TP-1552 " 1 

4. Title and Subtitle 

." " 

AERODYNAMIC CHARACTERISTICS OF A HYPERSONIC 

RESEARCH AIRPLANE CONCEPT HAVING A 70' SWEPT 

TO 1.20, WITH SUMMARY OF DATA FRO" 0.20 TO.6.0 - 

DOUBLE-DELTA WING AT MACH NUMBERS FROM 0.80 

7. Author(s) 

Jim A. Penland, James B. Hallissy, and James L. Dillon 

9. Performing Organization Name and Address 

" . -. - - 

NASA Langley Research Center 

Hampton, VA 23665 

12. Sponsoring Agency Name and Address 

-~ . " . ~ ___ 

National Aeronautics and Space Administration 

Washington, DC 20546 

15. Supplementary Notes 

3. Recipient's Catalog No. 

. - . - . - . - . .. . 

5. Report Date 

December 1979 

6. Performing Organization Code 

"" - - - -. . . 

" . - . "~ 

8. Performing Organization Report No. 

L-13158 

10. Work Unit No. 

505-11-33-01 

It. Contract or Grant No. 

13. Type of Report and Period Covered 

- 

Technical Paper 

14. Sponsoring Agency Code 

16. Abstract 

The data of the present report are divided into three areas. The first area includes the 

results of a wind-tunnel investigation of the static longitudinal, lateral, and directional 

stability characteristics of a hypersonic research airplane concept having a 70' swept 

double-delta wing. The force tests were conducted in the Langley 8-foot transonic pressure 

tunnel for Mach numbers from 0.80 to 1.20, for a Reynolds number (based on fuselage length) 

range of 6.30 X lo6 to 7.03 X lo6, at angles of attack from about -4' to 23O, and at angles of 

sideslip of 0' and 5'. The configuration variables included the wing planform, tip fins, the 

center vertical tail, and scramjet engine modules. The second area is a summary of the 

variations of the more important aerodynamic parameters with Mach number for Mach 

numbers from 0.20 to 6.0. The third area is a state-of-the-art example of theoretically 

predicting performance parameters and static longitudinal and directional stability over 

the Mach number range. 

7. Key Words (Suggested by Author(s)) 

~~ I 18. Distribution Statement 

" - . 

Hypersonic aircraft 

Transonic stability and control 

Unclassified - Unlimited 

Lift I 

Computational aerodynamics Subject Category 02 

9. Security Classif. (of this report) ' 20. Security Classif. (of this page) 21. No. of Pages 22. Price' 

. ~~ " - " ~ 

Unclassified Unclassified 144 $7.25 

.~ - "" 

1. Report No. 

". . i 2. Government Accession NO. 

~~ . . 

I 

- 

- 

1 

- 

1 

L 

For Sale by the National Technical Information Service, Springfield, Virginia 22161 

NASA-Langley, 1979 

I 

C 
National Aeronautics and . THIRD-CLASS BULK RATE 

Space Administration 

Postage and Fees Paid 

National Aeronautics and 

Space Administration 

NASA451 

,/Washington, D.C. 

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Official Business 

Penalty for Private Use, $300 

2 1 111.11. lllQ79 S00903DS 

DEPT 3F THE AIR F3RCE 

AF WERPOHS LABORIWORY 

ATTA: TXCBNICAL LIBBAR? (SUL) 

RIRTLAND AFB NM 37117 

NASA POSTMASTER: If Undeliverable (Section 158 

Postal Manual) Do Not Return