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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. From BL 27.94 outboard to the support booms, however, wing thickness ratio increased from FS fuselage station (axial 0.077 to 0.10 to provide adequate structural support location described by for the model and to permit transfer of compressed air distance in centimeters from the booms to the model propulsion system. from model nose) The metric portion of the model aft of FS 113.67, Fwd forward supported by the main force balance, consisted of the internal propulsion system, afterbody, tails (not shown HT horizontal tails in fig. l),and nozz!es. The afterbody lines (boattail) VT vertical tails were chosen to provide a length of constant cross sec- tion aft of the nonmetric centerbody and to enclose the WL water line, cm force balance and jet simulation system while fairing smoothly downstream into the closely spaced nozzles. Apparatus and Procedure The afterbody shell from FS 122.56 to FS 168.28 and tai! surfaccs (whcn insta!!cd) '::crc attrtchcd t3 a= after- Wind Tunnel body force balance which was attached to the main force This investigation was conducted in the Langley balance (fig. 1). The main force balance in turn was 16-Foot Transonic Tunnel, a single-return atmospheric grounded to the nonmetric wing-centerbody section. wind tunnel with a slotted octagonal test section and The nozzles were attached directly to the main force continuous air exchange. The wind tunnel has contin- balance through the propulsion system piping. Three uously variable airspeed up to a Mach number of 1.30. clearance gaps (metric breaks) were provided between
3 the nonmetric and the individual metric portions (after- sonic nozzles equally spaced around the supply pipe. body and nozzies) of the model at FS 113.67, FS 122.56, This method is designed to eliminate any transfer of ax- and FS 168.28, to prevent fouling of the components ial momentum as the air is passed from the nonmetric upon each other. A flexible plastic strip inserted into to the metric portion of the model. Two flexible metal circumferentially machined grooves in each component bellows are used as seals and serve to compensate the impeded flow into or out of the internal model cavity axial forces caused by pressurization. The cavity be- (fig. 1). tween the supply pipe and bellows is vented to model 1 In this report, that section of the model aft of internal pressure. The airflow is then passed through FS 122.56 is referred to as the total aft end (includes the tailpipes into the transition sections and then to afterbody, tails when installed, and nozzles). That the exhaust nozzles. (See fig. 1.) section of the model from FS 122.56 to FS 168.28 is The nonaxisymmetric (two-dimensional convergent- referred to as the afterbody, and that section aft of divergent) nozzle used in this investigation is shown in FS 168.28 is considered the nozzles. A skin-friction drag figure 8. The nozzle simulated a dry-power or cruise adjustment to the axial force results of the main balance operating mode with a design NPR of about 3.5. The was made for the section of the model from FS 113.67 nozzle throat area (17.48 cm2) and expansion ratio to FS 122.56. (See appendix.) (1.15) were sized to be consistent with advanced mixed The afterbody had provisions for mounting both the flow turbofan cycles. The ratio of total throat area to twin vertical tails and horizontal tails in three axial maximum body cross-sectional area was 0.11, and the positions. The vertical tails at a cant angle of 0", nozzle throat aspect ratio was 3.45. This nozzle was were tested in three positions-forward, mid, and aft- one of a series of nozzles tested in the study reported as shown in figure 4. With the vertical tails in the in reference 20. Nozzle static performance, ideal thrust mid position, cant angles of -lo", lo", and 20" were coefficients, and scheduled pressure ratios are presented also tested. The horizontal tails were only tested in in figure 9. the mid and aft positions which are about at the same positions as those of references 9 to 12. Note that both Instrumentation the vertical and horizontal tails have smaller tail spans Forces and moments on the metric portions of the when installed in the aft position than when they are model were measured by two six-component strain- installed in the other positions. gauge balances. The main balance measured forces and Sketches of the horizontal and vertical tails are pre- moments resulting from nozzle gross thrust and the ex- sented in figures 5 and 6, respectively. These tail sur- ternal flow field over that portion of the model aft of faces were sized to be representative of current twin- FS 113.67. The tandem afterbody shell balance mea- engine fighter aircraft. Individual root fairings (fillers) sured forces and moments resulting from the external contoured the tails to the afterbody at each tail posi- flow field over the afterbody and empennage surfaces tion. Clearance gaps were provided between the nozzles from FS 122.56 to FS 168.28. The tandem balance ar- and horizontal and vertical tails (aft position) in order rangement permits the separation of model component to prevent fouling between the main and afterbody bal- forces for data analysis. ances (fig. 4). These tail surfaces (without fairings) were Eight external seal static pressures were measured in also used in the investigations of references 9 to 11. the seal gap at the first metric break (FS 113.67). All orifices were located on the nonmetric centerbody and Twin-Jet Propulsion Simulation System spaced symmetrically about the model perimeter. An The twin-jet propulsion simulation system is shown additional five orifices, positioned symmetrically about in figure 1. An external high-pressure air system pro- the right side of the model measured seal gap pressures vides a continuous flow of clean, dry air at a controlled at the second metric break (FS 122.56). The final seal temperature of about 306 K at the nozzles. This high- pressures were measured by two sets of surface taps, pressure air is brought into the wind-tunnel main sup both consisting of two orifices, each an equal distance port strut where it is divided into two separate flows and fore and aft of the third metric break (FS 168.28). In passed through remotely operated flow-control valves. addition to these external pressures, two internal pres- These valves are used to balance the total pressure in sures were measured at each metric seal. These pressure each nozzle. measurements were then used to correct measured ax- The divided compressed airflows are piped through ial force and pitching moment for pressure-area tares as the wing-tip support booms, through the wings, and discussed in the appendix. into the flow-transfer bellows assemblies (fig. 1). A Chamber pressure and temperature measurements sketch of a single flow-transfer bellows assembly is taken in the supply pipe, upstream of the eight sonic shown in figure 7. The air in each supply pipe is dis- nozzles (fig. 7), were used to compute mass-flow rates charged perpendicularly to the model axis through eight for each nozzle. Instrumentation in each charging sec-
4 tion consisted of a stagnation-temperature probe and Horizontal tails aft, variable vertical tail position, a total-pressure rake. Each rake contained four total- and & = 0" ...... 11 pressure probes. (See fig. 8.) Horizontal tails aft, vertical tails mid, and variable All pressures were measured with individual pres- q5t ...... 12 sure transducers. Data obtained during each tunnel Variation of total aft-end aerodynamics with a for- run were recorded on magnetic tape and reduced with Horizontal tails mid, variable vertical tail position, standard data reduction procedures. Typically, for each and 4t = 0" ...... 13 data point, 50 samples of data were recorded over a pe- Horizontal tails aft, variable vertical tail position, riod of 5 sec and the average was used for computational and 4t = 0" ...... 14 purposes. Horizontal tails aft, vertical tails mid, and variable 4t ...... 15 Tests Summary data: This investigation was conducted in the Langley Total aft-end drag, CL = 0, and l6-Foot Transonic Tunnel at Mach numbers from 0.60 empennage location ...... 16 to 1.20 and at angles of attack from -3" to 9". Nozzle Interference drag terms, Ch = 0, and pressure ratio varied from 1 (jet off) to 9, depending empennage location ...... 17 upon Mach number. Basic data were obtained by vary- Interference drag terms, a = O", ing nozzle pressure ratio at zero angle of attack and and horizontal tails mid ...... 18 by varying angle of attack at fixed nozzle pressure ra- Interferfence drag terms, a = 0", tios. The investigation was conducted with different and horizontal tails aft ...... 19 empennage locations that included two horizontal and Comparison with axisymmetric nozzle three vertical tail positions. Vertical tail cant angle was configurations with a = 0" ...... 20 varied from -10" to 20" for one selected configuration. Total aft-end drag, CL = 0, and vertical Horizontal tail incidence was varied for selected config- tail cant angle ...... 21 urations from 0" to -10". Reynolds number based on Interference drag terms, a = O", the wing mean geometric chord varied from 4.4 x lo6 and vertical tail cant angle ...... 22 to 5.28 x lo6. All tests were conducted with 0.26-cm-wide boundary-layer transition strips consisting of No. 120 Discussion silicon carbide grit sparsely distributed in a thin film of lacquer. These strips were located 2.54 cm from the Basic Data tip of the forebody nose and on both upper and lower The basic data obtained during this investigation surfaces of the wings and empennage at 5 percent of the are presented in figures 10 to 15 for the various con- root chord to 10 percent of the tip chord. figurations tested at 6h = 0" only. Two types of data presentation are made to illustrate the effects of nozzle Presentation of Results pressure ratio and angle of attack. First, the variation The results of this investigation are presented in of total aft end, afterbody, and nozzle aerodynamic drag both tabular and plotted form. Table 1 is an index to and lift coefficients with nozzle pressure ratio at a = 0" the tabular results contained in tables 2 to 17. The com- is presented in figures 10 to 12. Second, the variation of puter symbols appearing in these tables are defined in total aft-end aerodynamic lift and drag coefficients with the section "Symbols" with their corresponding math- angle of attack is presented in figures 13 to 15 at jet-off ematical symbols which are described in the appendix. conditions and at a scheduled pressure ratio for each Plotted data are presented only at 6h = 0" because sub- Mach number (fig. 9(c)). Test parameters not shown sequent analysis cannot be made at a constant lift coef- in plotted form, such as total aft-end pitching-moment ficient. Because this investigation was conducted with coefficient, or obtained on configurations investigated a partially metric model, data were obtained at essen- at tail deflections other than 0" have been tabulated tially three different ranges of lift coefficient as tail de- (tables 2 to 17). flection was varied fro= 0" tg -10". Basic 2nd s?111?=lary In genera!, the v.riation Cf teta! &end drag coef- data for selected conditions at 6h = 0" are presented in ficient Co with nozzle pressure ratio for any particular figures 10 to 22 as follows: configuration (figs. 10 to 12), follows expected trends Figure (e.g., see refs. 6 and 9). Total aft-end drag decreases Variation of aft-end aerodynamics at a = 0" with initial jet operation up to nozzle pressure ratios with NPR for- of 2 to 3. This decrease in total aft-end drag is primar- Horizontal tails mid, variable vertical tail position, ily a result of a decrease in drag on the nozzles partic- and +t = 0' ...... 10 ularly at M = 0.60 and 1.20. This decrease in nozzle 5 drag is caused by a reduction in the external flow expan- The trends of zero-lift total aft-end empennage in- sion required at the nozzle exit as the exhaust flow fills terference drag coefficient (fig. 17) are the same as the nozzle base region. As nozzle pressure ratio is fur- for CD,, as vertical tail position is varied for the test ther increased, there is an increase in nozzle drag and, Mach numbers. Note that (ACD,~~),is negative at hence, total aft-end drag. Except at M = 0.60, nozzle M = 0.60 and 1.20; this indicates favorable interfer- drag subsequently decreases again with additional in- ence. At M = 0.60, this favorable interference is due creases in NPR. The change in drag trend with increas- to the tails-on nozzle drag being lower than the tails-off ing NPR (generally a drag increase) at NPR above 2 nozzle drag. At this Mach number, favorable interfer- to 3 is probably caused by exhaust flow entrainment ef- ence is a result of favorable interference on the nozzles. fects on the external nozzle flow, whereas the drag de- However, the favorable interference effects on the total creases at the higher nozzle pressure ratios (NPR > 3) aft end at M = 1.20 are caused from favorable inter- are caused by a compression at the nozzle exit created ference on the afterbody and not the nozzles (fig. 17). by the increased thickness of the exhaust flow plume. Similar results at M = 1.20 were found in reference 9. These trends with increasing nozzie pressure ratio are At M = 6.96, favorable interference on iiie nuzzks typical for jet-powered models (ref. 6). also occurred. However, the favorable interference ef- The effects of angle of attack on total aft-end aero- fects on the nozzles are negated by adverse interference dynamic characteristics shown in figures 13 to 15 are now present on the afterbody (fig. 17), which is greatly also typical for partially metric afterbody propulsion aggravated as the vertical tails are moved from the for- models. A single break occurs in the lift curves at ward to the mid position. With the vertical tails in a FZ 3' at M = 0.60, whereas at M = 1.20, there are the mid position, afterbody interference drag is 57 per- breaks at a = 0" and 3". At M = 0.90, the lift curves cent and total aft-end interference drag is 47 percent are nonlinear. Total aft-end drag polars also exhibit of the total aft-end zero-lift drag Co,,. As the verti- ' characteristics that are typical for afterbody propul- cal tails are moved from the forward to mid location sion models. A typical shape of the drag at M = 0.90 at M = 0.60 and 0.90, there is a decrease in nozzle probably results from changes in the wing downwash on drag (fig. 10) and a favorable increase (more negative) the afterbody and empennage surfaces in the transonic in nozzle interference drag increment. This trend is range. similar to that obtained previously on a single-engine configuration (ref. 6) but opposite to that obtained on Effect of Empennage Location a twin-engine configuration (ref. 9). Both of these stud- ies utilized models with axisymmetric nozzles. The effects of twin vertical tail longitudinal position on total aft-end zero-lift drag coefficients and individ- Horizontal tails aft. As shown in figure 16, ual zero-lift interference drag increments are presented a large increase in total aft-end zero-lift drag occurs in figures 16 and 17, respectively, for the two horizon- subsonically as the vertical tails are moved from the tal tail positions investigated. These values were deter- forward to mid position. Further movement of the mined by interpolation at CL = 0 from data obtained vertical tails to the aft position then results in a decrease as angle of attack was varied at constant nozzle pressure in CD,,. A similar trend was observed at M = 1.20, ratio (fig. 13(b),typical). These nozzle pressure ratios but the changes in CD,, are small. With the horizontal correspond to the schedule shown in figure 9(c). In ad- tails aft, the configuration with the vertical tails forward dition, the effect of nozzle pressure ratio on individual produced the lowest total aft-end drag at all Mach interference drag coefficients at a = 0" is shown in fig- numbers. In general, configurations with staggered tail ures 18 and 19 for the various configurations tested. arrangements (vertical tails forward, horizontal tails Note that these interference drag data are at a = 0' aft) have been found to have lower total aft-end drag (because of the method used to obtain data) rather than for both single- and twin-engine configurations (refs. 6 CL = 0; thus, absolute levels may differ from those pre- and 9). sented in figure 17. Examination of zero-lift individual interference drag increments shows a large effect of moving the horizontal Horizontal tails mid. There are no definite trends tail from the mid to aft position (fig. 17). With the to total aft-end zero lift drag CD,, as the vertical tails horizontal tails in the aft position, the increment in are moved from the forward to the mid position (fig. 16) empennage interference drag coefficient on the nozzle is for the test Mach numbers. The lowest value of CD,, always unfavorable, even though nozzle drag coefficient was measured for this investigation at M = 0.60 and is still negative at M = 0.60 and 0.90 (fig. 11). This NPR = 3.5 with all tail surfaces in the mid position probably indicates a reduction in pressure recovery on (fig. 16). The lowest jet-off (NPR = 1) value of CD,, the nozzles or flow separation on the nozzle sidewalls also occurred for this configuration (fig. 13(a)). when the horizontal tails are located adjacent to the nozzles. As the vertical tails are moved from the forward to the aft position, there is an increase in Drag from present Drag from empennage interference drag on the nozzle (except at study for- reference 9 for- M = 0.6 with vertical tails aft) and an increase in Schedule Schedule nozzle drag (fig. 11). This trend is opposite to the one M Jet off NPR Jet off NPR noted with the horizontal tails mid and is the same as 0.60 0.0039 0.0035 0.0050 0.0041 that reported in reference 9. .90 .0032 .0028 ,0040 .0030 At M = 0.9, there is a sharp increase in both the 1.20 .0166 .0158 .0150 .0122 total and afterbody empennage interference drag terms (fig. 17) as the vertical tails are moved from the for- ward to mid position. Similar results were also found with the horizontal tails in the mid position. Although M = 0.60. However, at M = 0.90 and M = 1.20, the reasons for this behavior are not known, one pos- the configurations of the present study nearly always sible explanation is that the adverse interference is a had higher drag than that of reference 9 because of result of the vertical tails being misaligned with the lo- unfavorable tail interference effects. In general, the cal flow field of the wing/forebody at the mid position. trends in Co with vertical tail movement are similar Reference 10 indicates that total aft-end drag was ex- to that of reference 9. tremely sensitive to vertical tail toe angle. The vertical Effect of Vertical Tail Cant Angle tail toe angle was 0" for the present investigation. At lifting conditions (figs. 13 and 14), the configu- The effects of vertical tail cant angle on total aft-end ration with the vertical tails forward generally had the zero-lift drag coefficient and increments in empennage lowest jet-on drag coefficient throughout the Mach num- interference drag coefficient at cr = 0" are presented in ber and angle-of-attack ranges except at M = 0.60. At figures 21 and 22. As can be seen, increasing tail cant this Mach number, the configuration with the vertical angle from -10" to 20" reduced total aft-end zero-lift and horizontal tails mid had the lowest drag over the drag coefficient and, in general, reduced each of the drag angle-of-at tack range. interference terms over the test Mach number range. Figure 15 also shows that reductions in drag coefficient Comparison with other data. A comparison of the from increasing tail cant angle also occur over the test total aft-end drag coefficient at CY = 0" of the present angle-of-attack range. Similar resulh were obtained in study with that of the configuration of reference 9 reference 10. is made in figure 20. Both of these configurations were identical up to FS 122.56 and used the same tail Conclusions surfaces. The afterbody (FS 122.56 to FS 168.28) of reference 9 was designed to have axisymmetric nozzles An investigation has been conducted in the Langley installed at FS 168.28. Since the nozzles of both these 16-Foot Transonic Tunnel to determine the effects of investigations had the same nozzle throat areas and empennage surface location and vertical tail cant angle expansion ratios, the afterbody closure ratios (ratio on the aft-end aerodynamic characteristics of a twin- of twice throat area to maximum body cross-sectional engine fighter-type configuration. The configuration area) were the same. featured two-dimensional convergent-divergent nozzles In order to make the comparisons between the two and twin vertical tails. The investigation was conducted configurations, it is first necessary to discuss the to- at different empennage locations that included two tal aft-end drag characteristics of both configurations horizontal and three vertical tail positions. Vertical tail without tails. Additional drag differences between the cant angle was varied from -10" to 20" for one selected two configurations are caused by tai! interference ef- configuration. Tests were conducted at Mach umbers fects since the drag of the tails is essentially the same. from 0.60 to 1.20 over an angle-of-attack range from Tails-off total aft-end drag coefficients at Q = 0" are -3" to 9". Nozzle pressure ratio was varied from jet off presented in the table on this page. As can be seen, (1) to approximately 9, depending upon Mach number. the nonaxisymmetric nozzle configuration of the present An analysis of the results of this investigation at a tail study has lower drag at M = 0.60 and 0.90 because of deflection of 0" indicates the following conclusions: the nozzie instaiiation. Tie higher drag ai Il.i = 1.20 I.4 rn.idii iuiar fa1 eace effects weie present thiwtghoiit is attributed to poor cross-sectional area distribution the test range of Mach numbers and were found to characteristics. be either favorable or adverse, depending upon test As seen in figure 20, the present configuration always condition and model configuration. At Mach number had lower total aft-end drag than that of reference 9 for of 0.90, adverse interference effects accounted for a a!! the codinations of empennage surfaces tested zt significant percentzge of total aft-end drag.
7 2. Interference effects on the nozzle were generally the axisymmetric nozzle configuration but unfavorable favorable but became adverse as the horizontal tails interference caused higher drag at a Mach number were moved from a mid to an aft position. of 0.90. 3. The effects of vertical tail position on aft-end 6. A decrease in total aft-end drag occurred as drag were usually dependent on Mach number and vertical tail cant angle was varied from -10" to 20". configuration. Generally, a forward position of the vertical tails produced the lowest total aft-end drag. 4. The configuration with nonaxisymmetric nozzles Langley Research Center had lower total aft-end drag with tails off than a sim- National Aeronautics and Space Administration ilar configuration with axisymmetric nozzles at Mach Hampton, VA 23665 numbers of 0.60 and 0.90. October 3, 1984 5. At a Mach number of 0.60, the nonaxisymmetric nozzle configuration had lower drag with tails on than
8 Appendix forward seal rim and interior pressure forces, respec- tively. In terms of an axial-force coefficient, the second Data Reduction and Calibration Procedure term ranges from -0.0001 to -0.0007 and the third term varies f0.0075, depending upon Mach number Calibration Procedure and pressure ratio. The internal pressure at any given set of test conditions was uniform throughout the in- The main balance measured the combined forces and no cavity flow was indicated. moments due to nozzle gross thrust and the external side of the model; thus, flow field of that portion of the model aft of FS 113.67. The momentum tare force FA,^^^ is a momentum tare The tandem shell balance measured forces and moments correction with jets operating and is a function of the due to the external flow field exerted over the afterbody average bellows internal pressure that is a function of and tails between FS 122.56 and FS 168.28. the internal chamber pressure in the supply pipes just Force and moment interactions exist between the ahead of the sonic nozzles (fig. 7). Although the bellows flow-transfer bellows system (fig. 7) and the main force were designed to minimize momentum and pressuriza- balance because the centerline of this balance was below tion tares, small bellows tares still exist with the jet the jet centerline (fig. 1). Consequently, single and com- on. These tares result from small pressure differences bined loadings of normal and axial force and pitching between the ends of the bellows when internal velocities moment were made with and without the jets operat- are high and also from small differences in the forward ing with ASME calibration nozzles. These calibrations and aft bellows spring constants when the bellows are were performed with the jets operating because this pressurized. The last term Df (eq. (Al)) is the fric- condition gives a more realistic effect of pressurizing tion drag of the section from FS 113.67 to FS 122.68. the bellows than capping the nozzles and pressurizing A friction drag coefficient of 0.0004 was applied at all the flow system. Thus, in addition to the usual balance- Mach numbers. interaction corrections applied for a single force balance Afterbody axial force is computed from a similar under combined loads, another set of interactions were relationship as follows: made to the data from this investigation to account for the combined loading effect of the main balance with the bellows system. These calibrations were performed over a range of expected normal forces and pitching mo- ments. Note that this procedure is not necessary for the afterbody forces because the flow system is not bridged Since both balances are offset from the model center- by the tandem shell balance. line, similar adjustments are made to the pitching mo- ments measured by both balances. These adjustments Data Adjustments are necessary because both the pressure area and bel- lows momentum tare forces are assumed to act along In order to achieve desired axial-force terms, the ax- the model centerline. The pitching-moment tare is de- ial forces measured by both force balances must also termined by multiplying the tare force by the appro- be corrected for pressure-area tare forces acting on the priate moment arm and subtracting the value from the model and the main balance corrected for momentum measured pitching moments. tare forces caused by flow in the bellows. The exter- nal seal and internal pressure forces on the model were obtained by multiplying the difference between the av- Model Attitude erage pressure (external seal or internal pressures) and The adjusted forces and moments measured by both free-stream static pressure by the affected projected balances were transferred from the body axis (which area normal to mnde! axis. The momentum tare force lies in the horizontal tail chord plane) of the metric was determined from calibrations with the ASME noz- portion of the model to the stability axis. Attitude of zle prior to the wind-tunnel investigation. the nonmetric forebody relative to gravity was deter- Axial force minus thrust was computed from the mined from a calibrated attitude indicator located in main balance axial force with the following relationship: the model nose. Angle of attack a,which is the angle between the afterbody centerline and the relative wind, FA - = FA,M~~!+ (pea,: - px)(Azb,i - -Asez!,:) Fj was determined by applying terms for afterbody deflec- -k (Pi - poo)Aiseal,l - FA,mom -k Df tion, caused when the model and balance bent under (All aerodynamic load, and by a flow angularity term to the where FA,J,fbal includes all pressure and viscous forces, angle measured by the attitude indicator. The flow an- internal and external, on both the afterbody and thrust gularity adjustment was 0.1", which is the average angle system. The second and third terms account for the measured in the Langley 16-Foot Transonic Tunnel.
9 Ideal Thrust cD,n = CD - CD,ait (-48)
The ideal isentropic gross thrust of each nozzle can cm,n = cm - Cm,ait (A91 also be determined if the mass-flow rate for each nozzle Tail Interference Terms is known. The effective discharge coefficients of the eight sonic nozzles (fig. 7) forward of each of the nozzle Vertical and horizontal tail drag was defined as the tail pipes were determined and used for measuring mass sum of form drag plus skin-friction drag for M 5 0.90 flow. and wave drag plus skin-friction drag for M > 1.00. The total ideal isentropic gross thrust or exhaust jet The subsonic form factors for the tails were calculated momentum for both nozzles is with the equation:
Form factor = 1+ 1.44(t/c) + 2(t/~)~ (A10) The individual fairings required for each tail location were also included in the skin-friction and wave-drag where Ti? the rate the is mass-,9ew =?easured i:: flew- caicuiations. Vaiues of cD,tails are given in tabie 18. transfer assemblies and pt,j is the average jet stagnation The tail interference terms used in this report are pressure for both nozzles. consistent with those used in references 6 and 9. The total tail interference increment on the aft end was Thrust-Removed Characteristics determined from The resulting force and moment coefficients (in- cluding thrust components) from the main balance in- clude total lift coefficient CL,~,drag-minus-thrust co- where (C~)t~il~on is the measured total aft-end drag for efficient C(D--F), and total pitching-moment coefficient a given configuration, (CD)tajls off is the measured aft- Cm,t. Force and moment coefficients from the tandem end drag for the same afterbody/nozzle configuration shell balance are afterbody (plus tails) lift coefficient with the tails removed, and CD,tails is the computed CL,ait, afterbody drag coefficient CD,aft, and afterbody value of tail drag as discussed previously. Hence this pitching-moment coefficient Cm,aft. total tail interference increment includes the interfer- Thrust-removed aerodynamic force and moment co- ence effects of one empennage surface on another, of efficients for the entire model were obtained by deter- the afterbody/nozzles on empennage surfaces, and of mining the components of thrust in axial force, normal empennage surface on the afterbody/nozzles. It also force, and pitching moment and subtracting these val- includes drag increments associated with misalignment ues from the measured total (aerodynamic plus thrust) of the empennage surfaces with the afterbody flow field. forces and moments. These thrust components at for- The empennage interference effects on the nozzles alone ward speeds were determined from measured static data were found from the following equation: and were a function of the free-stream static and dy- namic pressure. Thrust-removed aerodynamic coeffi- ACD,an = (CD,n)tails on - (CD,n)tails off (A12) cients are where the nozzle drags are obtained from equation (A8). CL = CL,~- Jet lift coefficient (-44) This empennage interference increment, then, is the re- sult of changes in nozzle external pressure distributions CD = C(D-F) + Thrust coefficient (A5) resulting from adding empennage surfaces to an after- body/nozzle configuration. The empennage interfer- C, = C,,t - Jet pitching moment coefficient (A6) ence increment on the afterbody alone was then defined Nozzle coefficients are obtained by simply combin- to be the difference between the empennage interference ing the measured results from both force balances as increments on the total aft end and the nozzles alone or follows: CL,n = CL - CL,aft (A71
10 References Nozzle Aerodynamic Characteristics. NASA TP-2158, 1. Nichols, Mark R.: Aerodynamics of Airframe-Engine 1983. Integration of Supersonic Aircraft. NASA TN D-3390, 11. Bare, E. Ann; and Berrier, Bobby L.: Investigation of In- 1966. stallation Effects on Twin-Engine Convergent-Divergent 2. Glasgow, E. R.: Integrated Airframe-Nozzle Perfor- Nozzles. NASA TP- 2205, 1983. mance for Designing Twin-Engine Fighters. AIAA Pa- 12. Leavitt, Laurence D.: Effects of Various Empennage Pa- per No. 73-1303, Nov. 1973. rameters on the Aerodynamic Characteristics of a Twin- 3. Runckel, Jack F.: Interference Between Exhaust System Engine Afterbody Model. AIAA-83-0085, Jan. 1983. and Afterbody of Twin-Engine Fuselage Configurations. 13. Berrier, Bobby L.: Empennage/Afterbody Integration NASA TN D-7525, 1974. for Single and Twin-Engine Fighter Aircraft. AIAA- 4. Richey, G. K.; Surber, L. E.; and Laughrey, J. A.: 83-1126, June 1983. Airframe/Propulsion System Flow Field Interference 14. Capone, Francis J.: The Nonaxisymmetric Nozzle-It Is and the Effect on Air Intake and Exhaust Nozzle Per- for Real. AIAA Paper 79-1810, Aug. 1979. formance. Airframe/Propulsion Interference, AGARD- 15. Nelson, B. D.; and Nicolai, L. M.: Application of Multi- CP-150, Mar. 1975, pp. 23-1-23-31. Function Nozzles to Advanced Fighters. AIAA-81-2618, 5. Berrier, Bobby L.; and Staff, Propulsion Integration Sec- Dec. 1981. tion: A Review of Several Propulsion Integration Features 16. Capone, Francis J.; and Reubush, David E.: Effects of Applicable to Supersonic- Cruise Fighter Aircraft. NASA Varying Podded Nacelle-Nozzle Installations on Ransonic TM X-73991, 1976. Aeropropulsive Characteristics of a Supersonic Fighter 6. Berrier, Bobby L.: Effect of Nonlifting Empennage Sur- Aircraft. NASA TP-2120, 1983. faces on Single-Engine Afterbody/Nozzle Drag at Mach 17. Carson, George T., Jr.; Capone, Francis J.; and Numbers From 0.5 to 2.2. NASA TN D-8326, 1977. Mason, Mary L.: Aeropropulsive Characteristics of Non- 7. Burley, James R., 11; and Berrier, Bobby L.: Investiga- axisymmetric-Nozzle Thrust Reversers at Mach Numbers tion of Installation Effects on Single-Engine Convergent- From 0 to 1.20. NASA TP- 2306, 1984. Divergent Nozzles. NASA TP-2078, 1982. 18. Capone, Francis J.; and Mason, Mary L.: Interference 8. Burley, James R., 11; and Berrier, Bobby L.: Effects Effects of Thrust Reversing on Horizontal Tail Effective- of Tail Span and Empennage Arrangement on Drag of a ness of a Twin-Engine Fighter Aircraft at Mach Numbers npical Single-Engine Fighter Aft End. NASA TP-2352, From 0.15 to 0.90. NASA TP-2350, 1984. 1984. 19. Peddrew, Kathryn H., compiler: A User’s Guide to the 9. Leavitt, Laurence D.: Effect of Empennage Location on Langley 16-Foot Ransonic lhnnel. NASA TM-83186, Twin-Engine Afterbody/Nozzle Aerodynamic Characteris- 1981. tics at Mach Numbers From 0.6 to 1.2. NASA TP-2116, 20. Yetter, Jeffery A.; and Leavitt, Laurence D.: Effects 1983. of Sidewall Geometry on the Installed Performance of 10. Leavitt, Laurence D.; and Bare, E. Ann: Effects of Twin- Nonaxisymmetric Convergent-Divergent Exhaust Nozzles. Vertical- Tail Parameters on Twin-Engine Afterbody/ NASA TP-1771, 1980.
11 TABLE 1. INDEX TO DATA TABLES
Position of- Table Horizontal tails Vertical tails 4t, deg 6hy deg 2 Off Off 3 Mid Forward 0 0 4 Mid Mid 0 0 5 Aft Forward 0 0 6 Aft Forward 0 -5 7 Aft Forward 0 - 10 8 Aft Mid 0 0 9 Aft Mid 0 -5 10 Aft Mid 0 - 10 11 Aft Aft 0 0 12 Aft Aft 0 -5 13 Aft Mid - 10 0 14 Aft Mid 10 0 15 Aft Mid 20 0 16 Aft Mid 20 -5 17 Aft Mid 20 - 10 ORIGMAL PACE ’is aE POOR QUALITY
13 7 .z U z D U
z c 2 N U
c k 4 r: V
I- LL 4 a V
I- LL < J u
x U
a V
J L’
b- I U
c LL I c Y u
#- J U
4 X a J -l
a P 2
I u 4 I: ORIGINAL PAGE ki QE POOR QUALITY,
15 e-- DU cc
p;' 0 Ft
"c 2; 35
16 uE
42 rm 3
4 m d
Frl c 0 V I%
J V
c LL I C b C’
-I n J <
a a Z
17 2 I: u
2 n u
z J V
I- IL a S u
G- #- cr LL 4 a Q m u t! 8
S e4 u x0
a u
b TL U
t- J p: L‘
4 x a J U
a a 2
18 ...... Ill II11l1I I Ill I1 1111 cn d 8 cll
4 cn d 8
e: 0
z4 V 0
a ...... 2 -~F~~-----~~~~m--RR~-----~~l.-l~R
20 4-0.maw*cm+LnJflJb f-almr-r-lnmmn*ooom~ Ot00tOCOCC3CC00 30003CCOC030VfC...... D1818DB88181bD1 om-mwrnv-w~rn~~=-m --mmJm-tNCr-mm-e OCCOOCCGC3=CCCC occcccocccccccc...... 8D888BDD Db88
o--~oC~-~cc--co 3...... s 5 3 c c c c c GC c c G 5 8 8fl P7W(Fr)8fl.cIm 8 8
;c e.0. c - rs? 7 s f =?-e6 c c c Co.~mC3CCCtlrJmmm --Nflm-c.---m*mrm......
21 I c Lj
2 -1 k- i) e
I- & 4 I Li
t U Q c Lj
s: 2 u
d O x c d V 2 tns b tn U wp: c x V
c J L’ .- do 40 a I w II n Cll 2 m-E; U
s n n I
1 L 4 > 22 z L u
I1I
L 3 0
2 4c -I cn L‘ d 2 t LL a r V
e LL a C E U 4 cn + LL d c 2 2 u z3 L u 2 ...... 0 I X C u
-I U
t > c
c u. a c- U
F 0 3 a h’ 9 b N - C C 0 G P b - rC )r) J e J wt-v c v u rr! crwtn c c n!s o c 3 c- -I mwmwQOmJNCutwwinhrc-fNoC V CCCCCCCCCCOCCCCCOCOO...... DamDI#IaI IaIaaIII
4 I n J a
IY P 2
I L: Q I
23 z S L'
2 0 u
z m J I4 V H 2 Gl d
+ iL a C u
C U
-1 e: V wE 8 d 4 c x I V L'
LA B c b U
c u
4 I a -1 <
U a z
24 2 I u
z c U
c LL 4 L U
c L 4 -n u 4c cn d ir 4 -1 2 U
c u
J u
c I V
c U m c Y u
c 2 V
4 I n J 4
~C--NmCmu.mc-"O II mccccQ.coI0.Q.c cc c QI V -R~NNQ~Q~~CSDQL~ 4 ...... 4 ...... I 4-d-W
25 I11 IllIIIltlllDIIDIIllIlI DICI IIIB
26 -L U
iA 4 I LJ
t- b- Lr Fr 4 c 4 b
I- LL -c U
Lr u
c LL I 0 v
c d- o
3 L 4 ....4...... 0...... a...... S
27 d. 0 V
2 ry A 0 U c
c LL U 2 U
?- iA c r L U
c iA Q 2 u
z U
c u
c E Li
c U
-.t u
P n L
28 TABLE 18. TAIL DRAG COEFFICIENTS
HT VT dt M CD,t ails Mid Forward 0 0.6 0.0032 .9 .0031 1.2 .0092 Mid 0 0.6 0.0033 .9 .0032 1.2 .0095 Aft Forward 0 0.6 0.0033 .9 .0032 1.2 .0093 Mid -10, 0, 10, and 20 0.6 0.0034 .9 .0033 1.2 .0095 Aft 0 0.6 0.0035 .9 .0034 1.2 .0097
29 30 ORiGflAL PAGE E5 DE POOR QUALITY
31 32 33 .'. .\
\ 1
0W \ N' N \ I \
\
34 WL 32.15- +- I I I I I I
WL 6.35- I
W? -6.35-
Figure 4. Location of horizontal and vertical tails.
35 , b M N
m t: L
36 I
iI
I I I I I I I I
......
37 c 0 v) C .-0 Y 8 a .-W YL E 0 Y .-W .wI E C 0 C E c2 .-L m
0 L W v)W c -c 3
C W $ M 5
eL m r-
c 0
m IC IC 1; L
8 r5
N00 . 0:
39 1.04
1.00
3 y W
.96
.92
.88
1.00
-I-? fi. .96 I
.92 * 2 3 4 5 6 7 NPR
(a) Static performance.
Figure 9. Nozzle characteristics
40 . 16
. 14
. 12
C F, i - 10
.08
.06
.04
-02 2 3 4 5 6 7 8 9 NPR
(b) Ideal thrust coefFicients.
Figure 9. Continued.
41 NPR
M
(c) Scheduled nozzle pressure ratios.
Figure 9. Concluded.
42 Ve rt i ca I t ai Is 0 Forward 0 Mid .020
,016
'D .012 cL
.008
.004
,020
.016
.02 c 'D,aft ,012
.008 -. 02 cL'aft -. 04OL- .004 1 .02
0 'D, n 'L, n
-. 02
-.04 0 2 4 6 8 10 NPR NPR
(3) .!! = n:c;n,
Figure 10. Effect of vertical tail position on afterbody aerodynamic characteristics for horizontal tails mid. dt = 0". and a = 0".
43 Vertical tails 0 Forward 0 Mid
r "L
.020 .04
016 .02
'D,aft .012 $,aft o
,008 -. 02
.004 -. 04
,004 .02
'D.n 0 cL, " '0
-.004 -. 02
-. 008 -. 04 0 2 4 6 8 1 NPR NPR
(b) M = 0.90.
Figure IO. Continued.
44 Verti cal t ai I s o Forward o Mid .032 .04
,028 .02
cD .024 cL 0
,020 -. 02 1
.016 -.04 T 1 ,024 .04
.020 .02
%,aft .016 ‘L,aft o
.012 -_02
,008 -. 04
,012 .02
,008 0 ‘D, n ‘L, n
.004 -. 02
0 -. 04 0 2 4 6 8 10 0 2 4 6 8 10 NPR NPR
(c) M = 1.20.
Figure 10. Concluded.
45 Vertical tails o Forward 0 Mid 0 Aft
,020 .04
,016
'D,aft ,012 'L, aft 0
,008 -. 02
,004 -. 04 .02
Ill I 1 I -.004 } I I -. 02
.04 0 2 4 6 a 10 NPR NPR
(a) M = 0.60.
Figure 11. Effect of vertical tail position on afterbody aerodynamic characteristics for horizontal tails aft, 4t = 0". and Q = 0".
46 Vertical tails 0 Forward 0 Mid 0 Aft .020 .04
.016 .02
‘D ,012 cL 0
,008 -. 02
,004 -. 04
,020 .04
.016 .02
cD. aft ,012 CL,aft 0
,008 -. 02
,004 -.04
.004 .02
C cD, n 0 L, n 0
-. coy14 -. 02
-. 008 -. 04 0 2 4 6 8 10 NPR NPR
(b) M = 0.90.
Figure 11. Continued.
47 Vertical tails 0 Forward 0 Mid 0 Aft .04
.02
r 'L 0
,020 I+ -. 02
,016 L -. 04 .024 .04
,020 .02
'D,aft ,016 5,aft 0
.012 -. 02
.008 -. 04
,012
.008 %," cL, " .004
n 0 2 4 6 8 10 NPR NPR
(c) A4 = 1.20.
Figure 11. Concluded.
48 .020 .04
I .016 .02
'D ,012 cL 0
,008 -. 02
.004 -. 04
f020 .04
,016 .02
'D, aft ,012 CL,aft 0
,008 -. 02
-. 04
.004 .02
'D,n 0 cL," 0
-. 004 -. 02
-. 008 -_04 0 2 4 6 10 0 12 4 6 10 t a a NPR NPR
(a) M = 0.60.
Figure 12. Effect of vertical tail cant angle on afterbody aerodynamic characteristics for horizontal tails aft, vertical tails mid. and a = 0".
49 Ot, deg
0 -10 00 3 10 !l 20
‘020 rrr .016
cD ,012
.@I8
,004
,020 .04
,016 .02
‘D,aft .012 0
,008 -. 02
,004 -.04 ,004 .02
‘D, n 0 0
-. 004 -. 02
-. 008 -. 04 0 2 4 6 8 10 0 2 4 6 8 10 NPR NPR
(b) A4 = 0.90.
Figure 12. Continued.
50 U+, deg
0 -10 00 0 10 6 20 .032 .04
.028 .02
‘D ,024 ‘L 0
.020 -. 02
.016 -. 04
,024 .04
.020 .02
‘D,aft ,016 ‘L,aft 0
.012 -. 02
,008 -.04
,012 .02
.008 0 ‘D. n ,n
.004
0 0 2 4 6 8 10 0 2 4 6 8 1 NPR NPR
(c) M = 1.20.
Figure 12. Concluded.
51 00 ru0
w 8
ru0 9
L P
20
0
S 2 0 E .-0
00
al07 uu d
0
0 B I
52 U 000 8 0
V
53 54 n 0
U 0 m U 0 A $1 H 0 8 0 N0 0U 0a' 8 1
2 0
55 56 0s:
e03 T 0 T
sN
I- 00 cu0
.z Yc *0
0 w’ 000 N0 8 0 I I I
57 N 3-n
0 0N
024
-I i 8
0 U-4 c 0
c 0
0
U 0 m U N0 N0 U0 a' 0 23 0 0 , I
2 V
58 59 60 0
61 \o 0m
0cum
1 R U 0 00 N 0 a' c-.( 2 3 0 8 B 0 0N B 0 I I I
62 T
63 N 0rn
m 0N
0Nu
<.I0 0
$1 0
I'€3- 000 0
2
00
arn u= d
0
U 0 m f 2 - 0 8
64 0x
0 cu0
0
0
m 0
65 0w
0
m UW d
\o 0W (v0 B 0 I I I n 0
f 000 8
67 \o 0Ln
(v 0In
W U0
I
8
we (v s II
._ m 0 cu0
0 -J0 0(v (v0
I
68 \o 0In
(u 0In
00 0-3
-3 0v
N 0m
00 N0
-N 9
-t-I I 0 0a3 8 0-3 s 0(u
-1 0
69 M
-- 0.60 _____ .90 - 1.20 Horizontal tail position Mid
*032.028 rT4
.024
.020
.016
.012
.008
.0°40 Fwd Mid El Fwd Mid Aft Vertical tail position
70 M 0.60 --- --_ .90 - -I- -- - 1.20 Horizontal tail position Mid Aft .008
.004
(“D,it’o 0
-. 004
.008
004
(“D,ia’o 0
-. 004
FWd Mid Fwd Mid Aft Vertical tail position
, Figure 17. Summary of effect of empennage location on individual interference drag increments at C, = 0 for scheduled NPR. & = 0”.
71 Vertical tails Forward ----- Mid
M - 0.60 M = 0.90 .008 rn rn
.004
"D, it
0
-.004
,008
,004
"D, ia 0
-. 004
,004
bC D,in 0
-. 004 1 3 5 1 3 5 7 1 3 5 7 9 NPR NPR NPR
Figure 18. Effect of vertical tail position on individual interference drag increments for horizontal tails mid, q5t = 0". and a = 0".
72 Vertical tails Forward _----Mid - -- Aft
M = 0.60 M 9 0.90 M = 1.20 .008
.004
“D, it 0
-. 004
008
.004
“D, ia 0
-. 004
,004
“D,in 0
-. 004 1 3 5 1 3 5 7 3 5 7 9 NPR NPR NPR
Figure 19. Effect of vertical tail position on individual interference drag increments for horizontal tails aft, dt = 0”. and a = 0”.
73 M Present study Ref. 9 0.60 0 0 .90 0 1.20 0 + Horizontal tail position Mid Aft At scheduled At scheduled Jet off NPR Jet nff N?!? .028
.024
.020
,016
.012
.008
0F wd Mid *Oo41 F wd Mid Aft F wd Mid Aft Vertical tail position
Figure 20. Comparison of total aft-end drag of present study with configuration with axisyrnrnetric nozzles. Q = 0". and
$I = 0".
I 74 M 0.60 -_ ___ .90 - 1.20 .032
.028
.024
.020
C D, 0 .016
.012
.008
.004
0 -10 0 10 20 @+, deg
Figure 21. Summary of effect of vertical tail cant angle on total aft-end zero-lift drag coefficient for scheduled NPR. horizontal tails aft, and vertical tails mid. -10 ----- 0 -___lo 20 M = 0.60 M-0.90 ,008
.004
“D, it
0
-. 004
,008
.004
“D, ia
0
-. 004
.004
“D,in 0
-. 004 1 3 5 1 3 5 7 1 3 5 7 9 NPR NPR NPR
Figure 22. Effect of vertical tail cant angle on individual interference drag increments for horizontal tails aft. vertical tails mid, +t = 0”. and a = 0”.
76 1. Report NO. 2. Government Accession No. 3. Recipient's Catalog No. NASA TP-2392 4. Title and Subtitle 5. Report Date EFFECTS OF EMPENNAGE SURFACE LOCATION February 1985 ON AERODYNAMIC CHARACTERISTICS OF A TWIN- ENGINE AFTERBODY MODEL WITH NONAXISYMMETRIC 6. Performing Organization Code NOZZLES 505-43-90-07 7. Author@) 8. Performing Organization Report No. Francis J. Capone and George T. Carson, Jr. L-15825
~ 9. Performing Organization Name and Address 10. Work Unit No. NASA Langley Research Center Hampton, VA 23665 l-----d11. Contract or Grant No. 13. Type of Report and Period Covered 12. Sponsoring Agency Name and Address National Aeronautics and Space Administration Technical Paper Washington, DC 20546 14. Sponsoring Agency Code 15. Supplementary Notes I
16. Abstract An investigation has been conducted in the Langley 16-Foot Transonic Tunnel to determine the effects of empennage surface location and vertical tail cant angle on the aft-end aerodynamic characteristics of a twin- engine fighter-type configuration. The configuration featured two-dimensional convergent-divergent nozzles and twin-vertical tails. The investigation was conducted with different empennage locations that included two horizontal and three vertical tail positions. Vertical tail cant angle was varied from -10' to 20' for one selected configuration. Tests were conducted at Mach numbers from 0.60 to 1.20 and at angles of attack from -3O to 9'. Nozzle pressure ratio was varied from jet off to approximately 9, depending upon Mach number. Tail interference effects were present throughout the range of Mach numbers tested and were found to be either favorable or adverse, depending upon test condition and model configuration. At a Mach number of 0.90, adverse interference effects accounted for a significant percentage of total aft-end drag. Interference effects on the nozzle were generally favorable but became adverse as the horizontal tails were moved from a mid to an aft position. The configuration with nonaxisymmetric nozzles had lower total aft-end drag with tails-off than a similar configuration with axisymmetric nozzles at Mach numbers of 0.60 and 0.90.
17. Key Words (Suggested by Authors(s)) Tail interference Empennage location Until February 1987 Twin noise Afterbody drag Nonaxisymmetric nozzles Subject Category 02 19. Security Classif.(of this report) 20. Security Classif.(of this page) 21. No. of Pages 22. Price Unclassified Unclassified 77
NASA-Langley, 1985