AN
D-7817 NASA TECHNICAL NOTE , NASA TN
N75-1818 ASA-TN-D-78 7 ) EFFECTS OF NOZZLE INTERFAIRING MODIFICATIONS ON LONGITUDINAL CHARACTERISTICS OF A TWIN JET, SAERODYNAMIC Unclas VARIABLE WING SWEEP FIGHTER MODEL (NASA) 128 p HC $5.75 CSCL 01A H1/02 12357
-,
EFFECTS OF NOZZLE INTERFAIRING MODIFICATIONS ON LONGITUDINAL AERODYNAMIC CHARACTERISTICS OF A TWIN-JET, VARIABLE-WING-SWEEP FIGHTER MODEL
David E. Reubush and Charles E. Mercer Langley Research Center Hampton, Va. 23665
D. C. * FEBRUARY 1975 NATIONAL AERONAUTICS AND SPACE ADMINISTRATION * WASHINGTON, 1. Report No. 2. Government Accession No. 3. Recipient's Catalog No. NASA TN D-7817 4. Title and Subtitle 5. Report Date EFFECTS OF NOZZLE INTERFAIRING MODIFICATIONS ON February 1975 LONGITUDINAL AERODYNAMIC CHARACTERISTICS OF A 6. Performing Organization Code TWIN-JET, VARIABLE-WING-SWEEP FIGHTER MODEL 7. Author(s) 8. Performing Organization Report No. David E. Reubush and Charles E. Mercer L-9802 10. Work Unit No. 9. Performing Organization Name and Address 760-17-01-11 NASA Langley Research Center 11. Contract or Grant No. Hampton, Va. 23665 13. Type of Report and Period Covered 12. Sponsoring Agency Name and Address Technical Note National Aeronautics and Space Administration 14. Sponsoring Agency Code Washington, D.C. 20546 15. Supplementary Notes
16. Abstract
A wind-tunnel investigation has been made to determine the effects of nozzle interfairing modifications on the longitudinal aerodynamic characteristics of a twin-jet, variable-wing-sweep fighter model. The model was tested in the Langley 16-foot transonic tunnel at Mach numbers of 0.6 to 1.3 and angles of attack from about -20 to 60 and in the Langley 4-foot supersonic pres- sure tunnel at a Mach number of 2.2 and an angle of attack of 00. Compressed air was used to simulate nozzle exhaust flow at jet total-pressure ratios from 1 (jet off) to about 21. The results of this investigation show that the aircraft drag can be significantly reduced by replacing the basic interfairing with a modified interfairing.
17. Key Words (Suggested by Author(s)) 18. Distribution Statement Afterbody drag Unclassified - Unlimited Boattail drag Interfairing modifications Fighter model STAR Category 02 22. Price* 19. Security Classif. (of this report) 20. Security Classif. (of this page) 21. No. of Pages Unclassified Unclassified 126 $5.75
For sale by the National Technical Information Service, Springfield, Virginia 22151 ON EFFECTS OF NOZZLE INTERFAIRING MODIFICATIONS LONGITUDINAL AERODYNAMIC CHARACTERISTICS OF A TWIN-JET, VARIABLE-WING-SWEEP
FIGHTER MODE L
By David E. Reubush and Charles E. Mercer Langley Research Center
SUMMARY
the effects of nozzle inter- A wind-tunnel investigation has been made to determine a twin-jet, fairing modifications on the longitudinal aerodynamic characteristics of tested in the Langley 16-foot tran- variable-wing-sweep fighter model. The model was of attack from about -20 to 60 and sonic tunnel at Mach numbers of 0.6 to 1.3 and angles number of 2.2 and an angle of in the Langley 4-foot supersonic pressure tunnel at a Mach nozzle exhaust flow at jet total- attack of 00. Compressed air was used to simulate pressure ratios from 1 (jet off) to about 21. The results of this investigation show that the aircraft drag can be significantly reduced at both subsonic and transonic speeds by modifications to the basic interfairing nozzle exit plane and addition of a (termination of interfairing upstream of the cruise pitching contoured body to the center) without significantly affecting airplane lift or engine nacelles upstream of moment. The addition of vortex generators mounted on the the nozzles reduced nozzle drag but resulted in an overall increase in drag.
INTRODUCTION
engines within their fuse- Military fighter aircraft are often configured to have twin rear of the fuselage. In contrast to the lages with the exhaust nozzle exits located at the arrangement offers compactness podded engine arrangement, this type of engine-fuselage it generally presents and a reduction of the one-engine-out stability problem, although and the nozzles. The flow other problems, such as difficulty in integrating the airframe and includes disturbances over the aft portion of a typical fighter configuration is complex and for carrier-based from such sources as horizontal and vertical tails, ventral fins, expansion and succeeding recompression airplanes, the tail hook which interacts with the change on the nozzle boattails. With such a complex flow field, often a relatively minor in some aircraft component can have a major effect on the aircraft drag. One aircraft component which can be relatively easily changed and offers promise of a large payoff in drag reduction is the interfairing between the nozzles. The purpose of this investigation was to determine the effects on airplane aerodynamic characteristics (drag, lift, and pitching moment) which result from modifications to the interfairings between the nozzles of a variable-wing-sweep fighter airplane configuration. Also, the addition of vortex generators on the engine nacelles just upstream of the nozzles was investigated for possible drag reduction.
The model used in this investigation had two fixed wing-sweep positions: 220 for subsonic speeds and 680 for supersonic speeds. Exhaust nozzles representative of power settings for cruise, partial afterburning, and maximum afterburning for two different engine packages were utilized. Nozzle exhaust flow was simulated by use of high- pressure air at about room temperature. In addition to the basic interfairing, six alter- nate interfairings were investigated. Data for the model with the various nozzles and basic interfairing have been previously reported in reference 1, and data obtained through the use of an aerodynamic model of this configuration with some of these alternate inter- fairings and others have been reported in reference 2. This investigation was conducted in the Langley 16-foot transonic tunnel at Mach numbers from 0.6 to 1.3 and in the Langley 4-foot supersonic pressure tunnel at a Mach number of 2.2. Angle of attack was varied from -2o to 60 in the Langley 16-foot tunnel and was held constant at 00 in the Langley 4-foot tunnel. The jet total-pressure ratio was varied from 1 (jet off) to about 21, depending on Mach number.
SYMBOLS
All force and moment coefficients are referenced to the stability-axis system and are based on the geometry of the model having a wing leading-edge sweep of 200. The origin of this axis system is at fuselage station 0.9127 m and water line 0.3175 m. All reference dimensions are given in meters; model dimensions are shown in centimeters.
Ae nozzle exit area, m 2
At nozzle throat area, m 2 b wing span, 1.6289 m
C D afterbody-nozzle drag coefficient, Aft-end drag + nozzle drag q2S
2 CD,n nozzle drag coefficient obtained from integration of nozzle pressures, Drag of two nozzles qS
Aft-end lift + nozzle lift CL afterbody-nozzle lift coefficient, Aft-end lift + nozzle lift q S
Cm afterbody- nozzle pitching-moment coefficient, Aft-end pitching moment + nozzle pitching moment q SE
ACx incremental coefficient due to a change in model configuration from a base- line to a modified configuration (x is a dummy variable) c mean aerodynamic chord of wing, 0.2490 m
1 reference length from airplane nose to tip of tail, 1.5685 m
M free-stream Mach number
Pt,j jet total pressure, N/m2
2 p0 free-stream static pressure, N/m q, free-stream dynamic pressure, N/m 2
2 S wing reference area, 0.3645 m x longitudinal distance from model nose (station 0), positive rearward, m a angle of attack, deg (see fig. 2(b))
0 angle of radius from nozzle center line to nozzle surface pressure orifice (clockwise positive for left nozzle, counterclockwise positive for right nozzle; facing upstream, 00 is at top of nacelle), deg
A wing sweep angle, deg
Abbreviations:
A/B afterburning 3 BL buttock line
FS fuselage station max maximum
S. L. sea level
WL water line
APPARATUS AND PROCEDURE
Wind Tunnels
This investigation was conducted in the Langley 16-foot transonic tunnel and in the Langley 4-foot supersonic pressure tunnel. The Langley 16-foot transonic tunnel is a single-return, continuous, atmospheric wind tunnel with a slotted octagonal test section. The tunnel speed is continuously variable between Mach numbers of 0.2 and 1.3. The Langley 4-foot supersonic pressure tunnel is a single-return, continuous wind tunnel with a stagnation pressure range of 27.58 kN/m 2 to 206.84 kN/m 2 and a stagnation tempera- ture range of 309 K to 322 K. By use of interchangeable nozzle blocks, the Mach number can be varied from 1.25 to 2.2.
Model
Photographs of the model mounted in the Langley 16-foot transonic tunnel and in the Langley 4-foot supersonic pressure tunnel are shown in figure 1. A sketch showing the principal dimensions of the model is shown in figure 2(a). The model was supported in the Langley 16-foot tunnel by a thin sweptback strut attached to the bottom of the fuselage just aft of the nose, as shown in figure 2(b). The strut blended into a sting which had a constant cross section beginning at the intersection with the strut trailing edge and extend- ing downstream to a station well aft of the model. Model details and dimensions are pre- sented in figure 3.
The model was tested with two wing-sweep positions: 220 for subsonic speeds (M < 1.0) with extendible glove vanes retracted and horizontal tails normally set at 00, and 680 for transonic and supersonic speeds (M ? 1.0) with glove vanes extended and hori- zontal tails set at -20. (Tails were set at -20 at transonic and supersonic speeds to insure that the balance did not foul out.) The inlets, located on each side of the fuselage, maintained true geometric lines but were closed to flow passage a short distance inside the inlet lip. The model consisted of three parts: the forebody and wings, the aft fuse-
4 lage and empennage (hereafter referred to as the afterbody), and the engine exhaust noz- zles. The forebody and wings were rigidly attached to the support system and were not metric. The afterbody was the metric portion of the model and started at the model metric break (station 1.1261 m); it included the horizontal and vertical tails, ventral fins, tail hook and fairing, aft fuselage, and interfairing between the engines. The metric break is indicated in the sketches of figure 2 and can be seen in the photographs shown in fig- ure 1. A flexible teflon strip inserted into slots machined into the metric and nonmetric portions of the model was used as a seal at the metric-break station to prevent flow through the gap between the afterbody and the forebody.
Two different sets of exhaust nozzles representing various power settings of two different nozzle types were tested. One set represented various power settings of a convergent-divergent iris type of nozzle (type A) and the second set, various power set- tings of a convergent-divergent balance-beam type of nozzle (type B). Photographs and geometric details of these nozzles are shown in figures 1 and 4, respectively. The noz- zle exhaust flow was simulated by use of a high-pressure compressed air system similar to that described in reference 3. The nozzles have been given configuration numbers, which conform to the configuration numbers used in references 1 and 4 and are as follows:
Configuration Nozzle Power setting Ae A t number type
03 A Cruise 1.05 07 A Maximum afterburning 1.21 09 B Cruise 1.02 10 B Sea-level maximum afterburning 1.19 11 B Transonic maximum afterburning 1.37 18 B Supersonic maximum afterburning 1.41
For this investigation, the model was supplied with seven interchangeable interfair- ings: the basic one and six alternates. (Note that the interfairing is defined as that part of the fuselage afterbody located between the engine nacelles and nozzles.) Details of the interfairings are presented in figure 3 and photographs of the various interfairings installed on the model are shown in figure 1. Interfairings 1, 2, 4, and 5 were of similar design, an uncambered wedge with varying droop angles (inclination of upper and lower surfaces to the horizontal). Interfairing 1 was of the same length as the basic interfairing, whereas the trailing edge of interfairings 2, 4, and 5 terminated slightly upstream of the nozzle exit plane (type A cruise). Interfairing 3 was a truncated basic interfairing but was not as short as interfairing 2. Interfairing 6 was interfairing 4 with the addition of a con- toured centerbody extending aft of the wedge trailing edge.
5 In addition to the various interfairings, the model was supplied with vortex genera- tors which could be mounted on the engine nacelles just upstream of the nozzles. The vortex generator blades were 0.762 cm long, 0.635 cm high, and 0.038 cm thick. They were arranged in opposing pairs with each blade inclined 150 to the free stream and their midpoints separated by 0.826 cm. The rings supporting the vortex generators could be installed above the interfairing and horizontal tail (fig. 1(s)) or below, or both sets could be installed at the same time.
Instrumentation
External static-pressure orifices were located on the exhaust nozzles as indicated in table I. In addition, internal static-pressure orifices were located in the afterbody cavity and at the seal station in the gap between the forebody and afterbody. The total pressures and temperatures of the jet simulation air were measured in each tail pipe by use of a total-pressure probe and a thermocouple. Forces and moments on the metric portion (afterbody) of the model were obtained by use of a six-component strain-gage bal- ance. An electronic flowmeter was used to measure the air mass flow rate to the nozzles.
Tests
Data were obtained for Mach numbers from 0.6 to 1.3 at angles of attack from -20 to 60 and for a Mach number of 2.2 at 00. The average Reynolds number per meter varied from 1.00 x 107 at M = 0.6 to 1.41 x 107 at M = 1.3 in the 16-foot tunnel and was about 1.19 x 107 at M = 2.2 in the 4-foot tunnel. The jet total-pressure ratio was varied from 1 (jet off) to about 21, depending on Mach number.
Transition was fixed on the model by means of 3.2-mm-wide strips of No. 120 car- borundum grains. The transition strips were located on the ventral fins and on the horizontal- and vertical-tail surfaces at a distance of 5.08 mm measured normal to the leading edge. The transition strips on the wing were located as shown in figure 5. A 3.2-mm-wide ring of transition grit was also located 13.5 mm aft of the nose of the fuselage.
Tests were conducted with the model equipped with the basic interfairing and six alternate interfairings and with the various nozzles listed previously. In addition to the tests with the various interfairings, other variables were investigated at subsonic speeds. One nozzle-interfairing combination was tested with horizontal-tail settings of +20 and -20. Another nozzle-interfairing combination was tested with two arrangements of vortex gen- erators mounted on the engine nacelles just upstream of the nozzles (fig. 1(s)).
6 Data Reduction force and Model data recorded on magnetic tape were used to compute standard are referenced to the ,ressure coefficients. All force and moment data in this paper of attack was corrected itability axes through the airplane center of gravity. Model angle of attack was cor- or support deflection due to loads and for tunnel upflow. The angle loads and by using the lift •ected by calibrating the support system for deflection due to for the various :urves from reference 5 to calculate the loads on the entire configuration data from references 6 :onditions. No correction was made for strut interference since system. mnd 7 indicate that the effect is small for a similar type of support axial force The afterbody axial force was obtained from the reading for balance :orrected for pressure-area terms which consisted of internal-cavity and seal-cavity from pressure measure- forces. The axial force on the exhaust nozzles was obtained pressure orifice (loca- ments by assigning an incremental projected area to each nozzle tions shown in table I) and summing the incremental forces. of the nozzle The afterbody force and moment increments and the drag increment the force or moment boattail due to interfairing variation were obtained by subtracting interfairing coefficient value obtained with a particular nozzle configuration and the basic and one of the alternate inter- from the value obtained with the same nozzle configuration the fairings. The increments due to the other test variables were obtained by subtracting combination with- force and moment coefficients obtained with a given nozzle-interfairing obtained with the same out the variation (for example, no vortex generators) from those nozzle-interfairing combination with the variation (for example, with vortex generators).
RESULTS AND DISCUSSION
Basic Force and Moment Data
Basic afterbody-nozzle force and moment coefficient data for the various configura- force and moment tions investigated are presented in figures 6 to 28. Afterbody-nozzle previously reported in data for the configurations with the basic interfairing have been reference 1 and therefore are not presented here. These figures present the afterbody- ratio for the various nozzle force and moment data as a function of the jet total-pressure these aerodynamic forces and Mach numbers and angles of attack. It should be noted that of the model (afterbody-nozzle moments represent only those measured on the aft portion and do not include forces and combination, approximately one-third of the model length) of this report moments on the wings or forward portion of the fuselage. Since the purpose to the alternates, these is to investigate the effect of changing from the basic interfairing data will be discussed. basic data will not be discussed and only the incremental
7 Incremental Force and Moment Data Figure 29 presents typical engine operating pressure-ratio schedules with Mach number for the two nozzle types used in this investigation. Data have been cross-plotted at these jet total-pressure ratios, and from these cross plots, incremental obtained data have been and are presented in figures 30 to 39. Effects of interfairing modifications.- The primary purpose of this investigation was to determine the effects of modifications of the interfairing between the nozzles on aircraft drag. The incremental changes in aerodynamic characteristics from those of the basic configuration are shown in figures 30 to 35 for all configurations tested. The subsonic drag increments (ACD) in figure 30 for the truncated basic interfairing (interfairing 3) show that this relatively minor change in the interfairing was not success- ful since it generally resulted in increased drag for both the type A and type B cruise noz- zles. Conclusions derived from examination of just the drag increment interfairings for the other (which all result in reduced drag for both nozzle types) could be erroneous because changes in aircraft lift and pitching moment Therefore, could result in a trim drag penalty. to obtain a valid answer as to which interfairing gives the best performance, both lift (fig. 31) and pitching-moment (fig. 32) increments must be examined in conjunc- tion with the drag increments. Interfairings 1, 2, and 5 do not give the best performance since they exhibit rela- tively large changes in the afterbody-nozzle lift and pitching moment at all subsonic speeds for nozzle type B and these changes would result in a trim drag fairing 4, penalty. Inter- a symmetrical uncambered wedge ending slightly upstream of the type A cruise nozzle exit, reduced the drag coefficient of the airplane aft end by about 0.0020 type B cruise (with nozzles installed) at subsonic speeds without significantly affecting the lift and pitching moment. However, interfairing 4 may not be a practical configuration because of the necessity of shielding the tail hook for carrier-based requirement aircraft and of the for additional volume to house the chaff and flare dispensers and the fuel dump.
Interfairing 6 (interfairing 4 with the addition of a contoured body at the center provide a more to gradual change in airplane area distribution in the vicinity of the nozzle exits, protection for the tail hook, and additional required volume) has a very slight drag advantage compared with interfairing 4 at subsonic speeds, whereas the lift and pitching- moment characteristics are about the same. At transonic speeds (M = 1.2, fairing fig. 34), inter- 4 produces a less beneficial drag increment for both type A and type B nozzles than interfairing 6. At supersonic speeds (M = 2.2, fig. 35), interfairing 2, which was the only interfairing modification tested at those speeds, shows little effect geometry on any of of interfairing the aircraft characteristics and this small effect is probably typical of any of the short wedge interfairings. Therefore, it must be concluded that interfairing 6 8 performance is the best of the interfairings investigated and does provide significant advantages at both subsonic and transonic speeds. (ACD, fig. 33) with the total drag incre- By comparing the nozzle drag increments n , percentage of drag improvement due ments (ACD, fig. 30) it can be seen that the largest the nozzles. This is probably due to the to the interfairing modifications is obtained on the recompression which occurs fact that at subsonic speeds with the short interfairings, help pressurize the nozzle boattail, whereas at the end of the interfairings can feed over to occurs downstream of the nozzles. At with the long basic interfairing, the recompression with help of a wave-drag transonic speeds, the addition of the contoured body (designed to smooth the trailing edge of interfairing 6 helps program) which extends downstream of rearward. Therefore, inter- out the area distribution without moving the recompression subsonic and transonic speeds. This fairing 6 is successful in reducing drag at both result agrees with that found in reference 2. nozzles, these results show that for For configurations with relatively wide-spaced nozzle interfairing should end upstream of the best performance at subsonic speeds, the at transonic speeds, the interfairing for sub- the nozzle exits. For the best performance of a body at the center to allow for a smooth sonic speeds may be modified by the addition subsonic performance. area distribution without adversely affecting the Because of the flow separation on the Effects of the addition of vortex generators.- aircraft drag without modifying the basic cruise nozzles, an attempt was made to improve generator configurations mounted on the interfairing by testing the model with two vortex of the nozzle boattail. One configuration engine nacelles just upstream of the beginning nacelles and the other only had them on had vortex generators completely encircling the and the horizontal tails. The results the top half of the nacelles between the interfairing and 37. Neither of these two vortex generator of this investigation are shown in figures 36 reduced nozzle drag, configurations reduced aircraft drag; however, both configurations was generally slightly more than as was expected. The full set of vortex generators drag increased because the drag twice as effective as the half set. However, the overall twice the value of the reduction in nozzle on the vortex generators was generally about generators was generally slightly drag. (See figs. 36 and 37.) The full set of vortex more than twice as bad as the half set. the possible drag pen- Effects of horizontal-tail incidence.- In an effort to evaluate at subsonic speeds to allow the use of inter- alties associated with retrimming the aircraft at M = 0.7 and 0.8 with the horizontal tails fairing 2, for example, the model was tested results (figs. 38 and 39) show that a tail set at 12o as well as at the normal 00. The for any of the interfairings deflection of 20 or less is sufficient to retrim the aircraft angle can have significant effects on aircraft investigated. However, this change in tail tail setting. These results tend to support drag, depending on aircraft angle of attack and 9 the effectiveness of interfairing 6, which changed aircraft lift and pitching moment very little, as the best type to improve airplane range capability.
CONCLUSIONS
An investigation of the effects of interfairing modifications and of other minor con- figuration modifications on the longitudinal aerodynamic characteristics of a twin-jet, variable-wing-sweep fighter model has indicated the following conclusions: 1. The replacement of the basic interfairing with a shortened symmetric uncambered wedge with a contoured body at the center resulted in significant drag reductions at both subsonic and transonic speeds without significantly affecting the aircraft lift or pitching- moment characteristics.
2. For the best performance at subsonic speeds of a similar configuration with relatively wide-spaced nozzles, the nozzle interfairing should end upstream of the nozzle exits. For the best performance at transonic speeds, the interfairing for subsonic speeds may be modified by the addition of a body at the center to allow for a smooth area distri- bution without adversely affecting the subsonic performance. 3. The addition of vortex generators on the engine nacelles upstream of the nozzles reduced nozzle drag but resulted in an overall drag-increase.
Langley Research Center, National Aeronautics and Space Administration, Hampton, Va., December 11, 1974.
10 REFERENCES
on 1. Mercer, Charles E.; and Reubush, David E.: Sting and Jet Interference Effects Longitudinal Aerodynamic Characteristics of a Twin-Jet, Variable-Wing-Sweep Fighter Model at Mach Numbers to 2.2. NASA TM X-2825, 1973.
2. Putnam, Lawrence E.: Effect of Nozzle Interfairings on Aerodynamic Characteristics of a Twin-Engine Variable-Sweep-Wing Fighter Airplane at Mach 0.60 to 2.01. NASA TM X-2769, 1973.
3. Mercer, Charles E.; and Berrier, Bobby L.: Effect of Afterbody Shape, Nozzle Type, and Engine Lateral Spacing on the Installed Performance of a Twin-Jet Afterbody Model. NASA TM X-1855, 1969. a 4. Reubush, David E.; and Mercer, Charles E.: Exhaust-Nozzle Characteristics for to 2.2. Twin-Jet, Variable-Wing-Sweep Fighter Airplane Model at Mach Numbers NASA TM X-2947, 1974.
5. Putnam, Lawrence E.: Effects of Modifications and External Stores on Aerodynamic Characteristics of a Fighter Airplane With Variable-Sweep Wing and Twin Vertical Tails at Mach 0.65 to 2.01. NASA TM X-2366, 1974. of 6. Re, Richard J.; Wilmoth, Richard G.; and Runckel, Jack F.: Investigation of Effects Afterbody Closure and Jet Interference on the Drag of a Twin-Engine Tactical Fighter. NASA TM X-1382, 1967.
7. Runckel, Jack F.; Lee, Edwin E.; and Simonson, Albert J.: Sting and Jet Interference Effects on the Afterbody Drag of a Twin-Engine Variable-Sweep Fighter Model at Transonic Speeds. NASA TM X-755, 1963.
11 TABLE I.- EXTERNAL NOZZLE ORIFICE LOCATIONS
[Configuration number in parentheses]
x/l x/1 0, x/1 0, x/1 x/1 X/ deg (03) (07) deg (09) deg (10) (11) (18) Engine type A nozzles Engine type B nozzles Left nozzle 0 0.935 0.932 12 0.909 12 0.919 0.919 0.919 20 .905 .909 12 .936 12 .937 .937 .937 35 .935 .932 12 .970 12 .955 .957 .959 35 .949 .940 54 .909 54 .919 .919 .919 35 .963 .947 71 .936 71 .937 .937 .937 50 .905 .909 60 .948 60 .943 .943 .943 75 .949 .940 60 .957 60 .955 .957 .959 75 .963 .947 60 .970 60 .974 .973 .974 95 .935 .932 108 .909 108 .919 .919 .919 135 .905 .909 108 .948 108 .937 .937 .937 135 .935 .932 108 .957 108 .943 .943 .943 135 .949 .940 108 .970 108 .974 .973 .974 180 .905 .909 180 .909 180 .919 .919 .919 180 .963 .947 180 .936 180 .937 .937 .937 225 .905 .909 180 .948 180 .943 .943 .943 225 .935 .932 180 .957 180 .955 .957 .959 225 .949 .940 180 .970 180 .974 .973 .974 264 .905 .909 228 .909 228 .919 .919 .919 264 .921 .921 228 .948 228 .937 .937 .937 264 .945 .937 228 .970 228 .955 .957 .959 275 .963 .947 276 .957 276 .955 .957 .959 287 .905 .909 290 .927 290 .926 .929 .929 287 .921 .921 324 .909 324 .919 .919 .919 287 .945 .937 324 .936 324 .937 .937 .937 315 .935 .932 324 .948 324 .943 .943 .943 315 .949 .940 324 .957 324 .955 .957 .959 315 .963 .947 324 .970 324 .974 .973 .974 350 .905 .909 348 .919 348 .919 .919 .919 Right nozzle 0 0.949 0.940 36 0.936 84 0.926 0.929 0.919 0 .963 .947 36 .948 84 .943 .943 .943 35 .905 .909 36 .957 88 .974 .973 .974 75 .935 .932 228 .936 252 .943 .943 .943 135 .963 .947 264 .927 252 .974 .973 .974 180 .921 .921 264 .943 264 .909 .909 .909 180 .935 .932 264 .957 264 .926 .929 .929 180 .949 .940 276 .970 290 .909 .909 .909 225 .963 .947 290 .943 300 .943 .943 .943 315 .905 .909 290 .957 300 .974 .973 .974
12 (a) Top rear view of nozzle type A, maximum afterburning power setting, with basic interfairing.
L-71-6677
(b) Bottom rear view of nozzle type A, maximum afterburning power setting, with basic interfairing. transonic tunnel Figure i.- Photographs of model installed in Langley 16-foot and 4-foot supersonic pressure tunnel. 13 L-74-3939 (c) Top rear view of nozzle type A, cruise power setting, with typical short wedge interfairing.
L-74-3941 (d) Bottom rear view of nozzle type A, cruise power setting, with typical short wedge interfairing. Figure 1.- Continued. 14 L-74-3947 with interfairing 2. (e) Rear view of nozzle type A, cruise power setting,
L-74-3943 setting, with interfairing 3. (f) Top rear view of nozzle type A, cruise power Figure 1.- Continued. 15 15 ORIGINAL PAi OF POOR QUALI'Tf L-74-3940 (g) Rear view of nozzle type A, cruise power setting, with interfairing 4.
L-74-3922 (h) Rear view of nozzle type A, cruise power setting, with interfairing 6. Figure 1.- Continued.
16 L-74-3925 (i) Top rear view of nozzle type B, cruise power setting, with typical short wedge interfairing.
L-71-6672 (j) Bottom rear view of nozzle type B, cruise power setting, with typical short wedge interfairing. Figure 1.- Continued. 17 L-74-3920 (k) Top rear view of nozzle type B, cruise power setting, with interfairing 3.
L-74-3918 (1) Bottom rear view of nozzle type B, cruise power setting, with interfairing 3.
Figure 1.- Continued.
18 L-74-3921
(m) Rear view of nozzle type B, cruise power setting, with interfairing 3.
L-74-3926
(n) Rear view of nozzle type B, cruise power setting, with interfairing 4. Figure 1.- Continued. 19 L-74-3928 (o) Rear view of nozzle type B, cruise power setting, with interfairing 5.
L-74-3931 (p) Top rear view of nozzle type B, cruise power setting, with interfairing 6.
Figure 1.- Continued.
20 L-74-3932 (q) Bottom rear view of nozzle type B, cruise power setting, with interfairing 6.
L-74-3933 (r) Rear view of nozzle type B, cruise power setting, with interfairing 6. Figure 1.- Continued.
21 L-74-3944 (s) Rear view of nozzle type A, cruise power setting, with basic interfairing and vortex generators on top of nacelles.
L-71-6659 (t) Rear view of nozzle type B, supersonic maximum afterburning power setting, with interfairing 2 in 4-foot tunnel; model mounted on side wall. Figure 1.- Concluded.
22 ,--1
FS 91.271 / 200*/
Wing pivot
Glove vane
FS 49.741 FS 0 BL 22.648 83.079 Bottom view Interfairing
13.547 162.885
Top view 480 e
inlet bleed Wing fences 0 E_ 24.896 BL 32.407 i i 68' FS 88.897 \
156.808
Wing fence
WL 27.767
8.110
149.528
power setting nozzles installed. (a) Model with nozzle type A and maximum afterburning support. Figure 2.- Sketch of model and geometric details of model All dimensions are in centimeters unless otherwise specified.
23 FS 112.607 FS 141.499
48.514 FS O
5.080
Angle-of-attack reference
< - 156.808
WL 27.767 .° 35.941 51.029 2.54 5 - f .08 15.24
(b) Geometric details of model support. Figure 2.- Concluded.
24 6522,648 1 FS 76.949 6.225 radius
14. 322
200 V p 35.373 Wing pivot
13.250
9.364
--- 21.635
b/2 = 81.443 BL 0 (a) Wing.
FS 55.668 '4
BL 14.499 220
.079 radius (typical)
Section A-A
(b) Glove vane.
Figure 3.- Details of model. All dimensions are in centimeters unless otherwise specified.
25 ,< 31.456
3506.103 3.5 <-- 16. 487
9. 490 BL 19.051 (NACA 65A004.643)
.791
22.488 23. 852 24.171
(NACA 65A003.16) 6.709
WL 29. 845 36.257 > FS 122.02 (c) Horizontal tail.
FS 137. 269 8 --- - 5.419 28.448 .948
8 4 7-1 28 . 44 8 -W L 24 . 10 0 = 6.129
4. 60 3.861
(d) Ventral fin.
1.1.727727 793 diameter 1.3721.372
" .736 .763 1.518.541.
WL 29 . 20 9 71 7 1.575 2.591 370 8.87 .839
9.271 28.701 FS 126. 893 (e) Tail hook and fairing.
Figure 3.- Continued. 26 FS 156.808 32.387 50
(NACA 65AO004.5) WL 56.304
21.671 21.591
/
26.033 8L 11.644
(f) Vertical tails (left shown). Vertical tails are toed in 10.
Exit of type A cruise nozzles Exit of type A maximum afterburning nozzles
I. LA --- BL 0 Section A-A B,
I I_ I I Section B-B
152.291 FS 156.845 FS 141.499 FS 149. 09 FS FS 138.748 FS 145.415 FS 150.707 FS 154.940
Section C-C
WL 30.996 Section D-D
Section E-E (g) Basic interfairing.
Figure 3.- Continued.
27 -- BL 0
FS 137.054 FS 156.846
59 .406
WL 31.220
5019, Interfairing I (h) Interfairing 1.
BL O
FS 137.054 FS 150.034
2045 2.588 .427 - - - 1 W L 31.220
Interfoiring 2
2.588 .211 - - W L 31.220
Interfairing 4
2.588 306 7045' .-- WL 31.220
1058! -A k .635 239' Interfairing 5 (i) Interfairings 2, 4, and 5. Figure 3.- Continued.
28 -- BL O
FS 148.590 FS 137.054 FS 151.342
4.798 radius
-- - WL 31.115
BL 0 and 1.693
4.077 radius
- -- WL 31.115
BL 3.387
3.960 radius S--WL 31.115
BL 5.080
Interfairing 3 (j) Interfairing 3.
Figure 3.- Continued.
29 A- B- C D E F\
- BL 0 Section A -A
_ E F
Section B - B
FS 143. 299 1 B C -FS 155.998 4 FS 145. 415 FS 153. 882 FS FS 147.531 FS 151.766 Section C - C FS 149. 648 FS 156. 845 FS 150.971
+-
Section D - D
WL 30.996 - +-f
Section E - E
Section F - F (k) Interfairing 6.
Figure 3.- Concluded. Typical
8.466 5.385 5.664 10.686
Configuration 03, cruise
10.792 > FS 141.499
8.466 8. 8.88 10.686 7.
Configuration 07, max A/B FS 141.499
(a) Nozzle type A, cruise and maximum afterburning nozzles.
Figure 4.- Sketches of nozzle configurations. All dimensions are in centimeters.
31 FS 141.499 .00 11.811
200 30'
10.688 28'
8,466 5.194 5.425
Configuration 09, cruise
FS 141.499 12.169 2108 8.892 12.108 S8062 75
5*45'S4 3.5863 5:T5,34 3.533 4.021S3.439 4.233 10024 3439 4.153
Configuration 10, sea-level max A/B Configuration II, transonic max A/B
<- 12. 154
-8.90030
122 52' 4.6729 3.899 4.762
Configuration 18, supersonic max A/B (b) Nozzle type B.
Figure 4.- Concluded.
32 64602
Js:119 I3.
,508
1.35f ~ ~~ 1351_-- 333
5.- Sketch showing transition location on upper wing surface for configurations Figure 5.- Sketch showing transition location on upper wing surface for configurations with A = 22o . All dimensions are in centimeters. with A = 220. All dimensions are in centimeters.
33 a, deg .08 ...... - : 60.0 .08 4.3
.08.0 1-jq ITl T
oJ044,4 o J7v
i- -i4~ 06- IiV~~
.012 00
.010 .002
.008 0 ,i/Pm P,i/
(a) M= 0.60. Figure 6.- Effect of jet total-pressure ratio on afterbody-nozzle aerodynamic forces and moments for type A cruise nozzles and interfairing 1. .08 T,' +;, .[ a,o.o deg .o ,~,r-rttcTr,_ -- A 4.34.3 trtrf4' LuV4~T L4 44-1 - .04 .04 L Lk ; rTi - - Cm H-' "+I t . 'L + I f i .012 F 14 -- 4 L .. . ,,
'[ ;-+--+ .04 -E-i~~ V TL.I .006 -.-10 ,i Di -. - i
-. 04 L -.
.o~o:+__ :+-1 - -1 .014 00 '--'L 0060 f -H-H-jt
.0 12't"'t '-it .004 lt
D Figure 6.- Continued.D, n
.010 .00 0 2 3 4 0 I 2 3 4 5 .008 3 5 0 2 3 4 5
Pt, i/poo Pt, i/Po (b) M =0.70. (b) M 0.70.
Figure 6. - Continued. a, deg .04 O 0.0 .12
S4.3
-.08 08
.0 6 --. 04 -. 08
-4-
------... .2 0 1 t It - -M- . 0 5 6.- Continued. -Figure 0 4 -. --- 0 0 - - - -.-..- -.- - - -- 0
0 ~~I"1 2..... 2 3 Pt a j/ :: P,:/ c ll~~~~~l~ l l ll l 0 .80.ltt l i i
11111Fiur 6.-~ Continued.I ... a, deg O 0.0 .12 O7 A 4.3
.08 CL -. 04
-LCm
.04
.016
.01 4 .006
CD .012 .004
.0120 .002
.010O .008 0
.008 00 2 3 4 5 0 I 2 3 4 5 Pt, /Po Pt, j/pOc
(d) M = 0.90.
Figure 6.- Concluded. a, deg .04 El 0.0 .08 A 4.3
CL O .04 Cm
-. 04 0 -
.016 -- -. 04
.014 .006"' ::' l
CD .0 2-t 1 .004
CD, n .010 .002 .
.008 O 0 2 3 4 5 0 2 3 4 5
Figure 7.- Effect of jet total-pressure ratio on afterbody-nozzle aerodynamic forces and moments for type A cruise nozzles and interfairing 2. a, deg 0 -2.0 0.0 .I2 .04 O 0 2.0 A 4.3 Lb 6.0 .08
.04 -. 04 Cm
-.08
.016 -. 04
-. 08 .014
.004 CD .012
CD, n .002 .010
4 5 .008 0 0 1 2 3 4 0 2 3 4 5 pt, j/pcO Pt, j/Poo
(b) M = 0.70.
Figure 7.- Continued. a, deg .04 - HE 0.0 .12 A 4.3
0 .08
-.04 -- Cm .04
-. 08 0 i ti
-- - -- I {I I I |I I I .01 6 -H I------k_t- p -. 04 m+ ------+-- - -
.0.14 ...... 006
------,, ,,,, , -+ H k H + - t o .012 -...- - -d .004
CD, n .010 ---- .002 ------7------gu- r --- n nu d
.008 0 0 2 3 4 5 0 1 2 3 4 5
(c) M = 0.80. Figure 7.- Continued. a,deg O 0.0
4.34 # t .08
.0.044
-..040800 .040
EI IL
-. 08 .014.
CO .012
.010 Co,n .002
.008 2 O4 3 4 5 0 I 2 3 4 5 0 2
Pt,/Poo Pt, j/P, (d) M = 0.90.
Figure 7.- Concluded. a, deg 0C 0.0 .08 1 ' -- 2.0 .08
CL 0 Cm 0
t ill I I
-.04 -. 04 ------
.016 ...... Go~n .006 i , ,_ -
CD-- I - - - .o~- -C ------o - - I+ . ItII .012 ---_ .0022- .012/ p0----- .0 I0 ------0 0 0 Figue8,I Effct2 o etto------u--ioo-ftrod-oz-3 4 5 0 1eodnmi 2 oce3 n 4 5 m oment - -o P .t,St, ------zl s an i te f ir ng 3 Pj (a) M = 0.60. Figure 8.- Effect of jet total-pressure ratio on afterbody-nozzle aerodynamic forces and moment for type A cruise nozzles and interfairing 3. a, deg O 0.0 4 2.0 .08 .08 A 4. 3
-0.04 .0 .6
- 04 -. 04
.0.00
.014 .004
CD KCD, n .012 .002
.010 0 2 3 4 5 0 I 2 3 4 5
C~D "p,/p% Pt,i/
(b) M = 0.70.
Figure 8.- Continued. a, deg O 0.0 .08 0 2.0 .08 A 4.3
L4 --0--. - - --,-,- - i 4 - .04-- --. , ,+
C Cm
-. 04 -. .04 - -
.O 6 ,l II .I006 i
[ T I I,' I I I I 11 IIr
.0 14 -- - -- .004 - - .01 2 .002 : ,. ------II ------I I ------.P0, -AAj1/p~ 00 p,,/pPt /
0 I 2 3 4 iue8- 5 ontin44- 0 I 2 3 4 5
(c) M = 0.80. Figure 8.- Continued. k------
OEO . 0 .08 0 .08
L-. 04 -Cm .04 k
- .08
.008 .018 08
.016 .006
CD .01 4 CD, n.004
.002 .012
0 I 2 3 4 0 1 2 3 4 5 Pt.j/Po Pt, /POD (d) M = 0.90. Figure 8.- Concluded. a,deg E 0.0 .08 2.0 .12
.04 0O..
.0 1-6 --- 04------
.014 .. .006
CO .01 2 .04
C0,
.0 10 .002
.008 0 01 2 3 4 5 0 2 3 4 5 Po , L Pt,j/ (a) M = 0.60.
Figure 9.- Effect of jet total-pressure ratio on afterbody-nozzle aerodynamic forces and moments for type A cruise nozzles and interfairing 6. Dashed line indicates possible fouling. a, deg
.08 12 A 4.3
o Cm .04
- .016 - 0 - -4 -4r
.o6 .-. .008 .0
0 1 2 3 4 5 0 I 2 3 4 5 Pt, j/% t,i/ 0 (b) M = 0.70. Figure 9.- Continued. a, deg CL 0.0 02.04 4a3
LH!L C.0 - .04 Cm
.012 ,IHF H ie -H------(I! - ' .006 ------i,,ii
.C010 ------.004
+HH-I F ------CD, n,n .-u , ., .J, ,,
------7 1 1 1 li .008 .002
.006 O 0 I 2 3 4 5 0 I 2 3 4 5
Figure 9.- Continued. 0.0 D .12 0 2.0 +H ......
. 48- 08 ------. L' - 0 - i I 1 -' T
- .08 .04
.016 - - ' 0
.014 .006
CD .012 - .004 CD, n
.010 -1- .002
.008 O 0 I 2 3 4 5 O I 2 Pt, j/Po Pt,/P (d) M = 0.90.
Figure 9.- Concluded. CJi
a, deg .08 l 0. 0 12
0 . .04 - I T
- .04 - - -0-,
CD .0 - n---.04
.04 .0
D . 12Dn .0 0T i I , co
O I 2 3 4 5 o I 2 3 4 5
(a) M = 0.60. Figure 10.- Effect of jet total-pressure ratio on afterbody-nozzle aerodynamic forces and
moments for type A cruise nozzles, basic interfairing, and vortex generators on top of nacelles. a, deg
S120 .1 ! ii CD, .08OO
.08 0 .02
CL .04 t i
i1ii i!r itt - gru -.--- "n i }
-. 0 4 4i144 .006
0 2 3 4 5 0 2 3 4 5 .012 o
ulT
Figure 10.- Continued. a, deg ,I2 O 0.0 .08 4.3
.08 .04
CL .04 - 0-O
0 ' _ -. 04
-. 04 .006
.016 .004
0 .014 002 ..
.012 0 0 I 2 3 4 5 0 2 3 4 5
Figure 10.- Concluded. a, deg 11 0.0 12 2. .08 1 - i4A:K 4.3 I ...
.04.04
08 04 .04 0
C -
012 0 I 2 3 4 5
Pt, j/Po Pt, j/Poo (a) M = 0.60. A Figure 11.- Effect of jet total-pressure ratio on afterbody-nozzle aerodynamic forces and moments for type cruise nozzles, basic interfairing, and vortex generators encircling nacelles. a,deg .12 0.0 .08 4..3 i l l6.0 i i04ll l il, l
CO 4i l i.02 I I S I 2 3 4 5 I 2 3 4 5
.08 .04
04 ure 11.- Continued.08 .018 il ...... 006 ......
.014 --- -- ! -00o . - ;. .. ------
Figure 11.- Continued. a, deg O 0.0 .08 .08 .082.0 A 4.3 L 6.0 . 7 .0.0 4
Cm O
-004 -.04 .04 .0 .020 - 8
.0 18 ...... 0 0 6 + ......
c...... l .o C, n
.01 2 .006 0 2 3 4 5 0 I 2 3 4 Pt, /P Pt, i/Pa
(c) M = 0.80. Figure 11.- Continued. a,deg O 0.0 .04 2 2.0 .08 ------A 4.3 L 6.0 - 0i ))l 0 4 - ...
-. 04 Tm F . .
-. O8~~~~~~~~ -. 08 ^,,t 41lllllllllllll- -,-.- .04 -! ----r -11-;------
.020 ' -. 08
. 18 .006
D.06"" .004 ----- 5 C. .01 6 .004 ._.
I I------Li u II .0 14 II I II I I C, n .0 02lF' ----- .014 .0802
.01 2 O O 2 13 0 ' 2 4 5
(d) M = 0.90. Figure 11.- Concluded. a,deg 0 0.0 .08 .08 0 2.0 # 4.3 L 6.0 .04 .04
CL 0 0
Cm
-. 04 - .04
.018 - .08
.016 -. 2
.004 CD .014
.012 CO, n .002
.010 O ,' 0 I 2 3TICL 4 EI~ f 5 o 2 3 4 5
(a) M = 0.70. forces and moments for type A Figure 12.- Effect of jet total-pressure ratio on afterbody-nozzle aerodynamic cruise nozzles, interfairing 2, and horizontal tails set at +20. cO a,deg I- - -- El 0. .08 A 4.3
.04 ------.04------.--I I - I------I l CL -
0 Cm 0
!i 1l t l l l l l l l /= -. 04 -. 04 3 L I IITl FF ------
.016 -.08 ---- {----
.O 4 Rt.lll I I II rL Rl.006
C0 ..012 .--- .004
C n
.0I 0 I.002
.008 O 0 I 2 3 4 5 0 I 2 3 4 5
/J0o Pt, I/Po (b) M = 0.80. Figure 12.- Concluded. a,deg O 0.0 .04 0 2.0 .16 A 4.3 6.0
0 .12
-. 04 Cm .08
-. 08 .04
.016 ll 10O
.014 .006 ill
CD .012 .004 CD, n
.008 O 0 I 2 3 4 5 0 2 3 4 5
(a) M = 0.70.
Figure 13.- Effect of jet total-pressure ratio on afterbody-nozzle aerodynamic forces and moments for type A cruise nozzles, interfairing 2,and horizontal tails set at -20. a, deg .04 O 0.0 .6 ------4.3
0 -- +-- .12
.01 6 n 0Or
.014 .006
CD, n
.0101111 .002 I II #a]!!!~ ~ (b M; 0.80.~lj!; ' i
.008 O
(b) M 0.80.
Figure 13.- Concluded. a, deg .08 lO 0.0 .16
CL
0 .08
-. 04 .04
.0 14 O
.002 T F4
0
0)! " " !I
O 1 2 4 5 O0 2 3 4 5 Pt, j/% , Pt, j /P 0 (a) M = 0.80. forces and moments for type A Figure 14.- Effect of jet total-pressure ratio on afterbody-nozzle aerodynamic maximum afterburning nozzles and interfairing 2. .04 a,deg .16 O 0.0 S2.0 A 4.3 0 EL 6.0 .1 2
CL -. 04 .08
-. 08 Cm .04
.01 6 I I I
.014 -. 04
.012 -. 08
CD .010 .004
.008 CD,n .002
0 2 3 4 5 0 2 3 4 5
(b) M = 0.90.
Figure 14.- Continued. a, deg 2.0 08 .08 4.3 .0
.04 .042
CLm
.008 .0 6
.024
D - -PCD, n .022 .004
.002 .0200
.01 80 3 4 5 1 2 3 4 5 6
(c) M = 1.20.
Figure 14.- Continued. .08 a, deg .04 0 2.0 .08
------0- - - fRlRfRT- - - -- A 4.J -- 44
.04 -. 04
------4
-.0 :-.024, .0...... 008 ...... (d) I I1 1 IIft1 1 1 f
-Figure 14.- Continued.
------Continue------
.018 O 0 .08
Cm CL .04 - 04
-. 08 0'
.006 .020
.004 .0 ) 8 CD CD, n .002 .016
.014 0 4 8 12 16 20 24 0 0 4 8 12 16 20
Pt, i/Po Pt, j/ (e) M = 2.20.
Figure 14.- Concluded.
U' a, deg S2.0
----- "+H' f ---- - +
-.----04 -..-... .. , - - - , .04
6 #-# 008 ,,,2------H.0 CD nil 4- --
- -. 024 ------t~ - i .006
------
.022 c n.04n 0, ,+++ T~l ------
Pto -/P----- P .020 Fiur 15.- Efec.... of to-pressure"rati on afeboy.002 nozl aeoyai forceso ,,,d -mome-n-s for...tyeAmxmmatrunn oze n neig 4 atl M' = 1.20. .018 . ... 0 2 3 4 5 6 0 I 2 3 4 5 6
Figure 15.- Effect of jet total-pressure ratio on afterbody-nozzle aerodynamic forces and moments for type A maximum afterburning nozzles and interfairing 4 at M = 1.20. 0 2.0 .08 .08 o4.3 - j- Fill 0.04
.04 04 Cm
C- 04
-. 04
S.010 .02 60
.024 .0
.020 .004
.008 .020O
.018 0.002 4 5 6 0 1 2 3 4 5 6 o 2 3 Ptj/P Pt, aerodynamic forces and moments for type A Figure 16.- Effect of jet total-pressure ratio on afterbody-nozzle maximum afterburning nozzles and interfairing 6 at M = 1.20. a, deg .08 O 0.0 .08 .-- - - A 4.3 .
04 .04
.04 - .04
0 .00
.012 0 0....- - { -
.008 04
0 2 3 4 5 o 2 3 45 Pt, j/P Pt, j/Po (a) M = 0.60. Figure 17.- Effect of jet total-pressure ratio on afterbody-nozzle aerodynamic forces and moments for type B cruise nozzles and interfairing 1. , deg 08 .08 Cm .04 o 4.3
.04 .006
.004
.012 .004 CD C, n
.010 .002
5 3 .008 0 4 2 0 2 3 4 5 Pt, i /p Pt, j/P0 (b) M = 0.70. Figure 17.- Continued. a, deg .04 0.0 08 .00MA 4.3
0 I ------T- ---,-O
-.o0 .04
!,41.00 6 , d. 1 0 -- - T
F!_ t fl F I- F .0 0 002I
.008 S I 2 3 4 5 0 0 2 3 4 5 P, /Po 'o Pt, j
(c) M =o.80. Figure 17.- Continued. Figure 17.- Continued. a, deg . 4 4.3 0 ..
CD -. 04 .08--. - 0 .08 .04
.008 0
.016 0
.04 2 3 4 5 0 2 3 4 5
CD .012 .oo4 ,-+H CD, n
.010 .002 - I,H
.008 0 4 5 0 2 4 5 0 2 3 4 St, //Pco (d) M = 0.90.
Figure 17.- Concluded. a, deg 04 3 0.o0 .08 A 4.3
0 .04
L-.04 0 --
-. 08 r _-.04
.010 .002
.08 .0
0 2 3 4 5 0 I 2 3 4 5 Pt, j/P Pt /P (a) M = 0.60. Figure 18.- Effect of jet total-pressure ratio on afterbody-nozzle aerodynamic forces and moments for type B cruise nozzles and interfairing 2. .04 0.0 .08
.04
Cm CL -.04 0 -.04
- .04 - .08
.0 I 4.06
.01.004 CDC, n
.0 .002
(b) M = 0.70.
Figure 18.- Continued. a, deg .08 1O 0.0 I 2 S4.3
.04 0
F- 1
L.01 .0064
-1- 7.08.08 -.04 14,1 -. 04 : :0 4
iii- .012 , 2 -. 004
P, j t
Fiure 18.- Continued. a, deg
0 JA 4.3
CL - .04 Cm .04
.046.006 - Cm+- , -:-- .0,
.010 .00
.008 M 2 3 4, 0 3 4 5 0 2 3 4 5 0 2 Pt, j/PCO Pt, /o (d) M = 0.90.
Figure 18.- Concluded. a,deg O 0.0 .08 0 2.0 .04 A 4. 3
CL m
O ---- -. 04
-. 04 - -.-.0808--
.016I .006I I I
.0 14 .00
.01 2 .002
.010 1 0O 0 1 2 3 4 5 0 2 3 4 5 Pt, j/P Pt, j/p (a) M = 0.60. Figure 19.- Effect of jet total-pressure ratio on afterbody-nozzle aerodynamic forces and moments for type B cruise nozzles and interfairing 3. Dashed line indicates possible fouling. a, deg O 0.0 .08 2.0 .08 A 4.3
.04 .04
CL .04 Cm 0 Mf O c
C-- -- .0 0 4,, l l ,, i , {[ [
-.04 -~~.004 ,, Lti ------
.014 .004
.042 .002
. 2 3 4 -17 5 . 0 2 3 4 5
(b) M = 0.70. Figure 19.- Continued. Figure 19.- Continued.
-ZR a, deg O 0.0
.04 .0- .04
-, C m .04 2.0 -. 04
. .008
.04 .0406
CD .012 C, n .004
.0 10 .002
.008 0 0 1 2 3 4 5 0 I 2 3 4 5 P,j/o Pt, j/P (c) M = 0.80.
Figure 19.- Continued. a, deg O 0.0 .04 , , , ,O 2.0 .12
CL Cm IF-.04 ..fT.04
-. 08 - - O
.0 16 .008
.01 4 .006
CD .012 -0 I, n . 4 Co'' . 12!!!!!!!! Co .O0ll .... ,,,,+
.010 : .002
.008 0 I 2 3 4 5 0 I 2 3 4 5 0
(d) M = 0.90. Figure 19.- Concluded. o
a, deg O 0.0 .04 2.0 .08 n 4.3
Ll II i i r
- 08 ll -04
.014D: ++1:: ------.08 1 1t l I ll l 0j ii0 I I I II I ilI IIl lI l III------ll l l ' i :. 0 IIIDI I III, InI I II i l
: : S ::3:: l l i l l l IIIII02 .00 0i ii2tll3ll4l5 0 l1ll lll l 4 5l .012 . .002
CD
.010 C0 0
.008 -.002 S I 2 3 4 5 0 I 2 3 4 5 Pt, j/ co Pt, J "c (a) M = 0.60. Figure 20.- Effect of jet total-pressure ratio on afterbody-nozzle aerodynamic forces and moments for type B cruise nozzles and interfairing 4. a, deg El 0.0 .08 .04 2.0 SA 4.3
.04 0
-. 04
- .04 - .08
-. 08 .014
.002 .012
CD
.010 CO, n O
.008O I 2 3 4 5 ,002 0 3 4 Pt, o Pt,i/Po (b) M = 0.70.
Figure 20.- Continued. a,deg O 0.0 .08 2.0 .08 A 4.3 i
o Cm 0
-. 04 ll l ill.04 -:-
,0 14 I-- - - -.08
.008 -. 002 .01I .002
CD .0 I CD,n -104O .008 I! II.I00I2 0 2 4 50 2 4! ll ll l l llPtll ll l/ lPoo: ll l l l Pt:::: O LJIIIIIIII~~~~illcl l ll l l l l MI~ 0.i8i0 .i i .0 2 3 4 5 Fiur 20,.- Coniud 0 2 3 4 5
(c) M = 0.80. Figure 20.- Continued. a, deg 0 0.0 0O0 2.0 .08
#m .04 CL -. 04
.0A1 4 .006
.012 .004 CD CD, n .002 .010 .00
0 ttrIrt1 .01 rl 1_ _ )T.0080 2 4 5 02
(d) M = 0.90.
Figure 20.- Concluded. a, deg o 0.0 .04 . 2.0 .04
CL 0O 0
Cm
-. 04 -.04
.01 4 -. 08
.01 2 l l l fll l l l l l.002
.O .0-. .H..,.' ,10,, , , II , ,, , , - CDnDIf I0q , ," ," , ' ,' , '" , ,," , , , , ,,f' H ++H , ,f , ,H, , ,,,, ,,, , ,,+( l l l, , -
.008 -. 002 0 2 3 4 5 0 1 2 34 5
(a) M = 0 .60.
Figure 21.- Effect of jet total-pressure ratio on afterbody-nozzle aerodynamic forces and moments for cruise nozzles and interfairing 5. Dashed type B cruise nozzles and interfairing 5. Dashed line indicates possible fouling. a, deg o 0.0 .04
A 4.3 ------: ,ILI -:-
CL O O Cm - - I- -I-l I -- - I-- - 4-
-. 04 < 2.0 -. 04
0-.04 T ... .014 :
- - -,-- ,, ,, , ,, ,,----L-, L LL L ~ ,,-,, ,, , ,, , .01.0 1 2 .002
II I II I I .l0 2 .00 iiiiI~l2,, , lin~l-,, , . , r- _ _ -4 5 0 , I ,,2, , I I, , ,, - H 41 1 151 t i ," J - '1 N l I 1 i~tiiiiiii It I/ I I 1 IL I ~Poo .010, . t ... .,,, .... ' b).M.0. 70.n ~ ~ l l I., - ,,, .008 2 3 4 5 0 I 2 3 4 5 I IIIIII III III IIII I igr 21. C ontIinIIued.II15 - III Il']d II Pt, ,/P,, I/IIo Pt, CID, ,,,.-. +.. , , , , ,, ,,- (b) M 0.70. , , , i , , ,, C."l lJ / IIIIIIlIII1 I|IIIul] - = I I ~ i l l l l I I I Figure 21.- Continued.
CJ1 a, deg . 0.0 .02.0 .04 A 4.3
.0 4l I II ......
O -. 04
-. 04 -. 08
.014 ,,,,,,,, .004 III II II ------0 . . .
Cl 0 ll C s l .012' ' i- -l 0 Il 1 t
,,,,,,,,,l(c), i , ,0 M,, .8..
II~lFigurel 2l1. -lConlnlll
.008 I I -. 002 1 1 1 0 I 2 3 4 5 0 I 2 3 4 5
(c) M = 0.80. Figure 21.- Continued. a, deg 0 0.0 .08 2.0 .08
.04 .04
-. 04 .04
.014 .006
.012 .004 C, n l CD 0 I2 3 5H
.010 .002
.008 5 0 2345 - CO 'l l l , , ,, ,, , , , , l 1 1' 1 I~ l l l --1~l
(d) M = 0.90.
Figure 21.- Concluded. 00
a, deg 0 0.0 .04 lil4.3 .0 8
-+------H l llk .I II,
C 00 .0
S.0l08 C D, O I
------.006 - - .002 I I 0 2 3 4 5 0 1 2 3 4 5
cruise nozzles and interfairing 6. a, deg 0 0.0 .08 .08 A 4.3 E
.04 .04 Cm
- 04
-. 04-.04
.01 2 .004
.002 .010
CD CO, n
.008 0
.006 .002 0 2 3 40 5 Pt, j/PO Pt, j/po (b) M = 0.70. Figure 22.- Continued.
co a, deg .08 0 0.0 .08 A 4.3
.04 .04 - - - - - I - -l- - - CL Cm ------
-.04 -.04 - -
.012 .004
0 2 , j/--3 4 5 0 --2 ,/%-- 3 4 5 Dt 0, n/Pm (c) M = 0.80. Figure 22.- Continued. a, deg o 0.0 0 .2
CL -.04 -- .08 Cm
-.08 .04
.014 0O
.012 .004
CD .010 .002 CD, n
.008 0
-. 002 .006 0 2 3 4 5 0 2 3 4 5 Ptjj/Poo Pt, /Poo (d) M = 0.90. Figure 22.- Concluded. .04 a, deg .I2 0 0.0 0 2.0
S04 6.0 .08
cLCm
-. 08 0O -
.014 -. 04 ...... --
.012 4-! -41- -.002
.008 - c, o0 J -
.006 -. 002 0 2 3 4 5 0 2 3 4 5
(a) M = 0.60.
Figure 23.- Effect of jet total-pressure ratio on afterbody-nozzle aerodynamic forces and moments for type B sea-level maximum afterburning nozzles and interfairing 2. a, deg l 0.0 .04 2.. 0 . I 2
- 0 .0.08LO-K 8r'
-.0 .0-04 Cm .0
20 4.01 1 -. 044.3 .014
.0 2.0 -. 08
n .008: - CD 0
-.0 .00612 3 4 0 , 2 3 4 5 0 , 2
Pt, j/Po Pt, j pO
(b) M = 0.70.
Figure 23.- Continued. a, deg o 0.0 .04 < 2.0 12
0 .08 .04 .04
-704 . . .04 -
------
.01 4 - .04
.01 2 E-.08
Co .010 .002
.008 f 0 0 s
.006 -. 002 0 I 2 3 4 5 0 I 2 3 4 5 t,J / . Pt Pao
(c) M = 0.80. Figure 23.- Continued. a, deg o 0.0
.04 2. - .08 -- L 6.0
.08 .0
S.o0 0
CD n .008 t, n-. 002
.014 -. 004 .012 I 2 4 0 I 2 3 4 5
(d) M = 0.90.
Figure 23.- Concluded. a, deg .08 4 2.0 .08
0 - K T, + 4.3
F:i~~ii17 ." -- ,.:, , '"- ...... --
M t .. "MI ,1 A lU-22 - .04 ..04 . . . .
j JJ t T - T~I
L L~
.04 -~~~.04 qf___ _
.026 . .. . .008 ......
0 3 2 5 6 1 4 U
Pt,jta -P4,'- P- TII1.0 .0gr24o 4Efcto TR etttl-rsue tranonicmaxiumai aterbrnigo 6 nozlesandzzearoyaicfreo nte*_4-1 2i admmet3+ 4 ortp 6 Pt, i/, p%,i
(a) M = 1.20. Figure 24.- Effect of jet total-pressure ratio on afterbody-nozzle aerodynamic forces and moments for type B transonic maximum afterburning nozzles and interfairing 1. a, deg .0 2.0 .0
.08:Fi~i~t- -t~-t-,~ - - tF'IT l T
.04 .04 0
2 35 - 6I
0- I 2 O I i p,.024 0 P,/ ,
(b) M = 1.30. Figure 24.- Concluded. a, deg 0 2.0 A 4.3
.08 .08
CcI- ---- R -W I --F~ ------i------.04 1 .04 ,"-
L -m 0 0
-. 04 -. 04
.0246 .008
-A-
.02'4 .006
C00 CD, n
.020 .004
02 3 4 5 6 7 0 2 3 4 5 6 7
(a) M = 1.2. Figure 25.- Effect of jet total-pressure ratio on afterbody-nozzle aerodynamic forces and moments for type B transonic maximum afterburning nozzles and interfairing 2. a, deg 0 1.9 A 4.0
.08 .08
.0.04 CCm
-. 04 l-.04
.006
CD
.018. 0 1 2 30 4 6 7 p,,j/p PtIj/ p
(b) M = 1.3.
Figure 25.- Concluded.
co a, deg .12 0 2.0 .12
.08 .08 Cm
CL .04 .04
0 0
-. 04 .008
.0214 .006
CD, n
C0 .022 .004
.0200 .002 0 2 3 4 5 6 7 0 2 3 4 5 6
Figure 26.- Effect of jet total-pressure ratio on afterbody-nozzle aerodynamic forces and moments for type B transonic maximum afterburning nozzles and interfairing 4 at M = 1.20. a, deg S2.0 .08 A 4.3 .08
.024 Tfll I.008 .02 ...... ------2L ------. 04 ...... -. 04
------II
1[ ll ln -4'iCDlll Pt, ./.. Pt, j .. i- ii i - -i- ---r ------,0 4 ------+ CD , '.020 .004 :
02 3 4 5 6 0 I 2 3 4 5 6
Figure 27.- Effect of jet total-pressure ratio on afterbody-nozzle aerodynamic forces and moments for type B transonic maximum afterburning nozzles and interfairing 6 at M = 1.20.
0 o
.080
.020 .006
.018i l .004
CD , n
.014 III I --- - .04 o 4 8 12 16 20 24 o 4 8 12 6 20 24 tJj/Po Pt, j/Po
Figure 28.- Effect of jet total-pressure ratio on afterbody-nozzle aerodynamic forces and moments for type B supersonic maximum afterburning nozzles and interfairing 2 at M = 2.20. Aerodynamic model flow-through Pt, jPoo ------Nozzle type A pt,j/Po
Nozzle type B Pt, j/Poo 14
12
0103
6
103 Nozzle type A Nozzle type B Nozzle type B Cruise Cruise S. L. max A/B
.002
.002
a 20 .004
.002
ACC D
-.002 110
.004
a 4. 30 .004
.002
-.002--
-. 004 1 2 3 6 1 2 3 4 5 6 2 Interfairing (a) M = 0.60. Figure 30.- Afterbody-nozzle drag increment due to changing from basic interfairing to various modified interfairings.
104 Nozzle type B. Nozzle type A Nozzle type B S. L. max A/B NozzCruisetype A Cruise Cruise
.002
0
-. 002
-. 004
= 20 .004
.002
AC0
-.002
-. 004
a = 4.3 .004
.002
-. 002
-. 004 1 2 3 6 1 2 3 4 5 6 2 Interfairing (b) M = 0.70.
Figure 30.- Continued.
105 Nozzle type A Nozzle type B Nozzle type B Cruise Cruise S. L. max A/B o 0o .002
-. 002
-. 004
a = 20 .004
.002
-. 002 M
-. 004
a 4. 30 .004
.002 - -
.002
-. 004 L 1 2 3 6 1 2 3 4 5 6 2 Interfairing (c) M = 0.80. Figure 30.- Continued.
106 Nozzle type A Nozzle type A Nozzle type B Nozzle type B Cruise Max A/B Cruise 000 .002
-. 002 MEE\
-. 004
S20 .004
.002
AC 0 0 - I I
-. 002
-. 004
a = 4. 3 .004
.002
.002 1 2 3 6 2 1 2 3 4 5 6 2 l nterfai ring
(d) M = 0.90.
Figure 30.- Concluded.
107 Nozzle type A Nozzle type 8 Nozzle type B Cruise Cruise S.L. max A/B
.02
0=00
.02
.02
-,02
a - 4. *
-. 02 1 2 3 6 1 2 3 4 5 6 2 Interfairing (a) M = 0.60. Figure 31.- Afterbody-nozzle lift increment due to changing from basic interfairing to various modified interfairings.
108 Nozzle type A Nozzle type B Nozzle type B max A/B Cruise Cruise S. L.
.02
-. 02
a - 2 .02
.02
a 4.3o .02
1 2 3 6 1 2 3 4 5 6 2 I nterfairing (b) M = 0.70. Figure 31.- Continued.
109 Nozzle type A Nozzle type B Nozzle type B Cruise Cruise S. L. max A/B
.02
a 230 -.02
(c) E = 0.80.
Figure 31.- Continued.
110 A Nozzle type A Nozzle type B Nozzle type B Nozzle type S. L. max A/B Cruise Max A/B Cruise o OP .02
-. 02
.02
a -4.30
-. 02 1 2 3 6 2 1 2 3 4 5 6 2 Interfairing (d) M = 0.90.
Figure 31.- Concluded.
111 Nozzle type A Nozzle type B Nozzle type B Cruise Cruise S. L. max A/B a=0 .02
.022
-. 04
a = 2 .02
m 0 M
.02------
a =4. 3' .02 -
0 - - -
.02
-. 04 1 2 3 6 1 2 3 4 5 6 2 I nterfairing (a) M = 0.60.
Figure 32.- Afterbody-nozzle pitching-moment increment due to changing from basic interfairing to various modified interfairings.
112 Nozzle type A Cruisee B Nozzle type B Cruise S. L. max A/B 00
.02
AC- - - -
-.02
a 4. 30
-.02
.04 2 1 2 3 6 1 2 3 4 5 6 Interfairing (b) M = 0.70. Figure 32.- Continued.
113 Nozzle type A Nozzle type B Nozzle type 8 Cruise Cruise S. L.max A/B o=0o .02
0 -a" --
-.02
.04
a 20 .02
Figure 32.- Continued.
.0214 a= 4.30
.04 -
Interfairing (c) M = 0.80.
Figure 32.- Continued.
114 Nozzle type A Nozzle type B Nozzle type B Nozzle Max A/B Cruise S. L. max A/B a 20 .04
0
.02
1 2 3 6 2 1 2 3 4 5 6 2
Interfairing (d) M = 0.90.
Figure 32.- Concluded.
115 Nozzle type A Nozzle type B Nozzle type B Cruise Cruise S. L. max A/B
.002
-. 002 -
-. 004
a =20 .004
.002
AC D, n 0 --
S.002
-.004
a = 4.30 .004
.002
------
. 0I
-.004-
1 2 3 6 1 2 3 4 5 6 2 Interfairing (a) M = 0.60.
Figure 33.- Nozzle drag increment due to changing from basic interfairing to various modified interfairings.
116 A Nozzle type B Nozzle type B Nozzle\type A Cruise S. L. max A/B Cruise a 0o .002
-. 002
-. 004
a = 20 .004
S002
-. 002
-. 004
a = 4.3 .004
.002
.002
.004 2 3 4 5 6 1.02 3 6 1 2 Interfai ring (b) M = 0.70.
Figure 33.- Continued.
117 Nozzle type A Nozzle type B Nozzle type B Cruise Cruise S.L. max A/B = 00 .002
-.002
.004
a = 20 .004
.002
ACD n 0
-.002
a 4.30
.002
-.004 1 2 3 6 1 2 3 4 5 6 2 Interfairing (c) M = 0.80. Figure 33.- Continued.
118 Nozzle type B Nozzle type B Nozzletype A Nozzle type A = Cruise Max A/B a0 Cruise .002
-. 002
-. 004
a = 20 .004
002
ACD, n 0
-. 002
-. 004
= 4.30 .004
.002
0
-.002
-. 004 2 3 6 2 1 2 3 4 5 6 2
Interfairing
(d) M = 0.90. Figure 33.- Concluded.
119 Nozzle type A Nozzle type B Max A/B Transonic max A/B .002
-. 002
.002
ACD, 0 - -
-. 002
.06
.04
ACL .02
-. 02
.02
m 0
-.02 2 4 6 1 2 4 6
Interfairing Figure 34.- Increments in afterbody-nozzle aerodynamic characteristics due to changing from basic interfairing to various modified interfairings at M= 1.20; a =20.
120 .002
AOD o
.002
ACD, n O
ACL 0 -
- .02
.02
ACm 0 - -
-. 02 Nozzle type A Nozzle type B Max A/B Supersonic max A/B characteristics Figure 35.- Increments in afterbody-nozzle aerodynamic due to changing from basic interfairing to interfairing 2 at M = 2.2; a =00.
121
\\ ODn AOL
ACD ACm .002 .02 ao a 00
0 -- 77 -7 0-
-.002 -.02
ACD, n ond Lan ACD and .002- a = 4.30 ACm .02 - a= 4.30
0 0
-. 002 -. 02 M = 0.60 M = 0.70 M =0.80 M = 0.60 M 0.70 M 0.80 (a) Nozzle and afterbody-nozzle drag. (b) Afterbody-nozzle lift and pitching moment.
Figure 36.- Increments in afterbody-nozzle aerodynamic characteristics due to addition of vortex generators to top of nacelles with type A cruise nozzles and basic interfairing. .004 ACCo
.002
0
-.002
S=20 .004
.002
0
ACD -. 002
or a 4.30 AC D,n .004
.002
-.002
a - 60 .004
.002
002
-.002 M =0.6 M =0.7 M =0.8 M 0.9
(a) Afterbody-nozzle and nozzle drag. aerodynamic characteristics due to Figure 37.- Increments in afterbody-nozzle nozzles addition of vortex generators encircling nacelles with type A cruise and basic interfairing.
123 ON x ALL
.02
= ° 0 2
.02
ACL -. 02 or AC a = 4. 3 m .02
C------
-. 02
a 60 .02
.02 M=0.6 M=0. 7 M=0.8 M 0.9 (b) Afterbody-nozzle lift and pitching moment.
Figure 37.- Concluded.
124 .004
.002
ACD 0
-. 002
.002
ACD,n 0
.04
ACL .02
0
0
-.02
AC m -. 04
0 a0OO a2 0 a-4.3o a-6 a-00 a-4.30 0.8 M- 0.7 M aerodynamic characteristics due to Figure 38.- Increments in afterbody-nozzle cruise nozzles and deflecting horizontal tails from 00 to +20 with type A interfairing 2. 125 .002
AC0D 0
-. 002
.002
AC, n
.0
ACL -
-. 04
.06
.04
AC
.02
a 0' a 20 a 4.3 oa 66 a0 aa 4. 30 M 0.7 M 0. 8 Figure 39.- Increments in afterbody-nozzle aerodynamic characteristics due to deflecting horizontal tails from 00 to -2o with type A cruise nozzles and interfairing 2.
126 NASA-Langley, 1975 L9802