r. nincn 1 -I mnwn i echnicai Paper 2392 i February 1985 Effects of Empennage Surface Location on Aerodynamic d Characteristics of a Twin- Enyint:L Afterbody Model With Nonaxisyrnrnet ric Wozzles I I NASA Technical Paper 2392 1985 Effects of Empennage Surface Location on Aerodynamic Characteristics of a Twin- Engine Afterbody Model With Nonaxisymmetric Nozzles Francis J. Capone and George T. Carson, Jr. Langley Research Center Hampton, Virginia National Aeronautics and SDace Administration Scientific and Technical Information Branch Summary craft because of adverse interactions originating from empennage surfaces, base areas, actuator fairings, and The effects of empennage surface location and ver- tail booms (refs. 1 to 5). tical tail cant angle on the aft-end aerodynamic charac- A comprehensive program to study the interference teristics of a twin-engine fighter-type configuration have effects of empennage surfaces on single- and twin-engine been determined in an investigation conducted in the fighter afterbody/nozzle drag has been conducted at the Langley 16-Foot Transonic Tunnel. The configuration Langley Research Center (refs. 6 to 11) because these in- featured two-dimensional convergent-divergent nozzles terference effects can account for a major portion of to- and twin vertical tails. The investigation was conducted tal aft-end drag. These studies, which are summarized at different empennage locations that included two hor- in references 12 and 13, were conducted with configu- izontal and three vertical tail positions. Vertical tail rations with conventional axisymmetric nozzles. Little cant angle was varied from -10' to 20" for one selected information is currently available on empennage effects configuration. Tests were conducted at Mach numbers on configurations with advanced nozzle concepts. from 0.60 to 1.20 and at angles of attack from -3' to This paper presents results from an investiga- Nozzle pressure ratio was varied from jet off (1) to 9". tion of the effects of horizontal and vertical tail approximately 9, depending upon Mach number. An position on twin-engine fighter aft-end drag with a analysis of the results of this investigation was made at model which had nonaxisymmetric (two-dimensional a tail deflection of 0'. 0'. convergent-divergent) nozzles. This exhaust system has Tail interference effects were present throughout the the potential to satisfy many different mission require- test range of Mach numbers and were found to be either ments with less installation penalties than axisymmet- favorable or adverse, depending upon test condition and ric nozzles (refs. 14 to 16). The present study was model configuration. At a Mach number of 0.90, ad- part of an overall research program that also deter- verse interference effects accounted for a significant per- mined nonaxisymmetric nozzle thrust reverser perfor- centage of total aft-end drag. Interference effects on the mance (ref. 17) and effects of thrust reversing on hori- nozzle were generally favorable but became adverse as zontal tail effectiveness (ref. 18). This investigation was the horizontal tails were moved from a mid to an aft conducted in the Langley 16-Foot Transonic Tunnel at position. The effects of vertical tail position on aft-end Mach numbers from 0.60 to 1.20, at angles of attack drag were usually dependent on Mach number and con- from -3" to go, and at nozzle pressure ratios up to 9. figuration. Generally a forward position of the vertical Horizontal tail incidence angle was varied from 0" to tails produced the lowest total aft-end drag. The config- - 10". uration with nonaxisymmetric nozzles had lower total aft-end drag with tails of€ than a similar configuration Symbols with axisymmetric nozzles at Mach numbers of 0.60 and 0.90. At a Mach number of 0.60, the nonaxisymmetric Model forces and moments are referred to the stabil- nozzle configuration had lower drag with tails-on than ity axis system with the model moment reference center the axisymmetric nozzle configuration but unfavorable located 4.45 cm above the model centerline at fuselage interference caused higher drag at a Mach number of station 91.6 cm, which corresponds to 0.25E. All co- 0.90. A decrease in total aft-end drag occurred as ver- efficients are nondimensionalized with respect to qooS tical tail cant angle was varied from -10' to 20°. or q,SE. A discussion of the data reduction procedure and definitions of the aerodynamic force and moment Introduction terms and the propulsion relationships used herein are presented in the appendix. The symbols used in the The mission requirements for the next generation computer-generated tables are given in parentheses in fighter aircraft may dictate a highly versatile vehicle ca- the second column. pable of operating over a wide range of flight conditions. These aircraft will most likely be designed for high ma- neuverability and agility, will operate in a highly hostile &b,l model cross-sectional environment, and will possess short take-off and land- area at FS 113.67 and ing characteristics to operate from bomb-damaged air- FS 122.56, cm2 fields. These aircraft require variable geometry nozzles model cross-sectional area to provide high iniernai nozzie performance; thus, im- at FS 168.28, cm2 portant aft-end parameters such as closure and local boattail angles continuously change throughout the op Aseal.1 cross-sectional area erating range of Mach number, angle of attack, and enclosed by seal strip engine pressure ratios. Large drag penalties can result at FS 113.67 and from integration of the propulsion system into the air- FS 122.56, cm2 cross-sectional area en- cm total aft-end aerody- closed by seal strip at namic pitching-moment FS 168.28, cm2 coefficient total aft-end drag Cm,aft afterbody (plus tails) coefficient pitching-moment coefficient afterbody (plus tails) drag coefficient Cm,n nozzle pitching-moment coefficient nozzle drag coefficient total aft-end pitching- tail drag coefficient Cm,t moment coefficient (in- drag-minus-thrust coeffi- cluding thrust compo- cient, C(D--F) G CD at nent), C%,? = C, at NPR = '1 (jet off) NPR = 1 CD at CL = 0 E wing mean geometric chord, 44.42 cm increment in empen- nage interference drag D f friction drag, N coefficient on afterbody total aft-end axial force, N (es. (A1311 FA axial force measured by increment in empennage FA,M b a1 interference drag coeffi- main balance, N cient on nozzle (eq. (A12)) FA,mom momentum tare axial force increment in empennage due to bellows, N interference drag coef- FA,Sbal axial force measured by ficient on total aft end afterbody shell balance, N (eq. (All)) Faft afterbody (plus tails) axial increment in empennage force, N interference drag coef- ficient on afterbody at Fi ideal isentropic gross CL = 0 thrust increment in empennage F' thrust along body axis, N interference drag coeffi- M free-stream Mach number cient on nozzle at CL = 0 NPR nozzle pressure ratio, increment in empennage interference drag coeffi- pt,jlpm cient on total aft end at m measured mass-flow rate, CL = 0 kg/sec ideal isentropic gross mi ideal mass-flow rate, thrust coefficient kg/sec total aft-end aerodynamic Pea,l average static pres- lift coefficient sure at external seal at FS 113.67, Pa afterbody (plus tails) lift coefficient Pes,2 average static pres- sure at external seal at nozzle lift coefficient FS 122.56, Pa total aft-end lift coefficient (including thrust com- Pea,3 average static pres- ponent), CL,t CL at sure at external seal at NPR = 1 FS 168.28, Pa average internal static Test-section plenum suction is used for speeds above a pressure, Pa Mach number of 1.05. A complete description of this facility and operating characteristics can be found in average jet total pressure, reference 19. Pa free-stream static pressure, Model and Support System Pa Details of the general research, twin-engine fighter free-stream dynamic afterbody model and wing-tip-mounted support system pressure, Pa used in this investigation are presented in figure 1. wing reference area, Photographs of the model and support system installed 4290.00 cm2 in the Langley 16-Foot Transonic Tunnel are shown in figure 2. A sketch of the wing planform geometry is gas constant, 287.3 J/kg-K presented in figure 3. thickness-chord ratio The wing-tip model support system shown in fig- ure 1 consisted of three major portions: the twin jet total temperature, K support booms, the forebody (nose), and the wing- (ALPHA angle of attack, deg centerbody combination. These pieces made up the nonmetric portion (that portion of the model not ratio of specific heats, mounted on the force balance) of the twin-engine fighter 1.3997 for air at 300 K model. The fuselage centerbody was essentially rect- horizontal tail deflection, angular in cross section having a constant width and positive leading edge up, height of 25.40 cm and 12.70 cm, respectively. The four deg corners were rounded by a radius of 2.54 cm. Maxi- mum cross-sectional area of the centerbody (fuselage) leading-edge sweep angle, was 317.04 cm2. The support system forebody (or nose) deg was typical of a powered model in that the inlets were vertical tai! cant angle, faired over. For these tests, the wings were mounted positive tip out, deg above the model centerline (model has capability for both high or low wing mount). The wing had a 45" Abbreviations: leading-edge sweep, a taper ratio of 0.5, an aspect ra- tio of 2.4, and a cranked trailing edge (fig. 3). The ASME American Society of NACA 64-series airfoil had a thickness ratio of 0.067 Mechanical Engineers near the wing root to provide a realistic wake on the BL buttock line, cm afterbody.