Document
AIAA 99-2670
Experimental Investigation of Convoluted Contouring
for Aircraft Afterbody Drag Reduction
K. A. Deere and C. A. Hunter
NASA Langley Research Center
Hampton, VA
35th AIAA/ASME/SAE/ASEE
Joint Propulsion Conference & Exhibit
June 20-24, 1999 / Los Angeles, CA
For permission to copy or republish, contact the American Institute of Aeronautics and Astronautics 1801 Alexander Bell Drive, Suite 500, Reston, VA 20191-4344 AIAA 99-2670
Experimental Investigation of Convoluted Contouring for Aircraft
Afterbody Drag Reduction
Karen A. Deere † and Craig A. Hunter † NASA Langley Research Center Hampton, VA ABSTRACT An experimental investigation was performed in the NASA Langley 16-Foot Transonic Tunnel to determine the aerodynamic effects of external convolutions, placed on the boattail of a nonaxisymmetric nozzle for drag reduction.
Boattail angles of 15° and 22° were tested with convolutions placed at a forward location upstream of the boattail curvature, at a mid location along the curvature and at a full location that spanned the entire boattail flap. Each of the baseline nozzle afterbodies (no convolutions) had a parabolic, converging contour with a parabolically decreasing corner radius. Data were obtained at several Mach numbers from static conditions to 1.2 for a range of nozzle pressure ratios and angles of attack. An oil paint flow visualization technique was used to qualitatively assess the effect of the convolutions. Results indicate that afterbody drag reduction by convoluted contouring is convolution location, Mach number, boattail angle, and NPR dependent. The forward convolution location was the most effective contouring geometry for drag reduction on the 22° afterbody, but was only effective for M < 0.95. At M = 0.8, drag was reduced 20 and 36 percent at NPRs of 5.4 and 7, respectively, but drag was increased 10 percent for M = 0.95 at NPR = 7.
Convoluted contouring along the 15° boattail angle afterbody was not effective at reducing drag because the flow was minimally separated from the baseline afterbody, unlike the massive separation along the 22° boattail angle baseline afterbody.
INTRODUCTION engine pressure ratio. However, mechanical variable area nozzles, intended to improve internal nozzle performance, also alter the external aft-end shape, closure, and boattail For an afterbody of a typical fighter aircraft that angle, often at the expense of external aerodynamic accounts for about 35 percent of the total aircraft length, performance. In addition, attempts to shorten the length it is often surprising that the afterbody is responsible for of the propulsion system to improve efficiency, decrease up to 50 percent of the total aircraft drag at transonic weight, and reduce skin friction drag can result in a short, conditions. Approximately half of the afterbody drag steep boattail geometry at a dry power setting. The trade results from adverse interference effects when integrating off is a geometry that can encourage flow separation and the propulsion system with the airframe, and from result in excessive drag. In such cases, external flow pressure drag on the afterbody (refs. 1-2). While separation can occur at subsonic conditions when the propulsion airframe integration is of equal importance, boundary layer cannot overcome the adverse pressure this study focuses on the reduction of afterbody pressure gradient on the boattail. At transonic speeds, flow drag.
acceleration along the boattail terminates with a shock, and the resulting shock-boundary layer interaction leads to Fighter aircraft encounter a wide range of flight separation.
conditions to perform a required mission. To maintain high performance, the nozzle geometry must be Previous work (ref. 3) has shown that convoluted continuously optimized for changes in Mach number and contouring can alleviate, and in some cases eliminate, shock-induced boundary layer separation at transonic †Aerospace Engineer, Configuration Aerodynamics Branch, conditions. Convoluted contouring has also shown Aerodynamics Competency. Member AIAA promise for alleviating separation in subsonic Copyright © 1999 by the American Institute of Aeronautics and applications (refs. 4-5). Studies have indicated that Astronautics, Inc. No copyright is asserted in the United States under convolutions can reduce bluff body drag by nearly 75 Title 17, U.S. Code. The U.S. Government has a royalty-free license to exercise all rights under the copyright claimed herein for percent. In the current investigation, convoluted government purposes. All other rights are reserved by the copyright contouring was placed on a nozzle afterbody, shown in owner American Institute of Aeronautics and Astronautics AIAA 99-2670 figure 1, to evaluate the potential for reducing drag by M = 1.2. A detailed description of the facility and alleviating separation. operating procedures are found in reference 6.
Model Installation and Support System The nozzle afterbody was attached to the single- engine propulsion simulation system and mounted in the tunnel on a sting strut, as shown in figure 2. The model was composed of a forebody that covered the high- pressure plenum, a centerbody that covered the low- pressure plenum and instrumentation section, and the nozzle afterbody. Hardware downstream of MS 26.9 was metric, or on a six-component force balance. The nozzle afterbody started at MS 54.486, and is defined as x/L = 0 in the figures of pressure coefficient.
Figure 1. Convoluted contouring on an aircraft afterbody.
A standard grit application method was used on the model forebody to transition the boundary layer to NOMENCLATURE turbulent flow. Number 100 silicon carbide grit particles were sparsely applied to a 0.1-inch wide strip located 1.5 A exit area, 6.944 in inches aft of the model nose tip.
e A reference area, 42.396 in ref 15 ∞ MS 69.70 22 ∞ MS 66.57 A throat area, 4.972 in MS 0.00 MS 54.486 MS 45.00 MS 26.9 (Metric break) t Centerbody A /A expansion ratio, 1.397 e t Low pressure plenum Instrumentation Forebody Nozzle Transition section Afterbody section CD total drag Flexible seal High pressure plenum (metal bellows) Choke plate C skin friction drag D,f Tunnel C pressure drag, C - CD,f centerline D,p D C pressure coefficient, p - p / q p • • 8 equally spaced Total temperature probe nozzles exiting radially Airflow L nozzle length, inches Total pressure rake M free stream Mach number MS model station, inches NPR nozzle pressure ratio, pt,j /p • 45 ∞ p surface static pressure, psi 45 ∞ p average jet total pressure, psi t,j p free-stream static pressure, psi Figure 2. Single-engine propulsion simulation system • q free-stream dynamic pressure, psi with the nozzle afterbody on a sting-strut mount.
• x/L normalized axial location along afterbody a angle of attack, deg Propulsion Simulation System An external high-pressure air system provided a APPARATUS AND EXPERIMENTAL METHODS continuous flow of clean, dry air at a stagnation temperature of approximately 540°R to the nozzle. High- Wind Tunnel pressure air was routed through six air lines in the support system to the nonmetric high pressure plenum.
To minimize the axial momentum generated by This investigation was conducted in the 16-Foot transferring the air from the nonmetric, high-pressure Transonic Tunnel at NASA Langley Research Center.
plenum to the metric, low-pressure plenum, the air was This single-return, continuous-flow, atmospheric wind discharged radially through eight equally spaced sonic tunnel has a slotted octagonal test section and continuous nozzles. Flexible metal bellows were used to seal the air air exchange. Variable airspeeds allow testing from Mach system between the plenums and minimize pressurization numbers of 0.3 to 1.2. Test section plenum suction is forces. Data were corrected for pressure force tares during used for Mach numbers above 1.05. The Reynold's 6 6 the data reduction process. The air flowed from the low- number per foot varies from 3x10 at M = 0.6 to 4x10 at pressure plenum, through a choke plate, into the American Institute of Aeronautics and Astronautics AIAA 99-2670 3 2 instrumentation section, and exhausted through the test z = -0.0027705 x + 0.0627048 x - 0.301426 x + 2.9007 (3) nozzle.
Model Description 15° Nozzles 22° Nozzles This research effort was conducted to investigate the Baseline 354.17 288.64 effect of convoluted contouring on the external afterbody Forward 429.01 351.25 of an exhaust nozzle, for alleviating separation and Mid 494.38 390.21 reducing total drag. The addition of convolutions to the Long 526.88 421.82 external surface of the baseline afterbody shape increased surface area, and hence, skin friction drag. Wetted area estimates for each afterbody are shown in Table 1, and Table 1: Nozzle wetted area, square inches.
equivalent "flat plate" skin friction estimates are given in Table 2. Nozzles with boattail angles of 15° and 22° were M=0.6 M=0.8 M=0.9 M=1.2 tested with convolutions placed at a 'forward' location 15° Baseline 0.019 0.018 0.018 0.017 upstream of the boattail curvature, at a 'mid' location 15° Forward 0.024 0.023 0.022 0.021 along the curvature, and at a 'long' location that spanned 15° Mid 0.028 0.027 0.027 0.025 the entire boattail flap. The convolution locations are 15° Long 0.031 0.029 0.029 0.027 illustrated in figure 3. Each baseline nozzle (no 22° Baseline 0.015 0.014 0.014 0.013 convolutions) had a parabolic, converging contour with a 22° Forward 0.019 0.018 0.018 0.017 parabolically decreasing corner radius.
22° Mid 0.021 0.020 0.020 0.019 22° Long 0.024 0.023 0.022 0.021 Convoluted contours were designed based on previous work (refs 3-5), and were characterized by vertical walls with constant radius hills and valleys. In Table 2: Nozzle skin friction drag coefficient, C D,f the spanwise direction, each convolution cycle had an amplitude to period ratio of 0.6:1, or equivalently, a hill/valley height to width ratio of 2.4:1 (fig. 4(a)). In BOATTAIL the streamwise direction, the convolution run was defined ANGLE
15 ∞ 22 ∞
by a flattened bell curve and had a 7 to 10 percent height CONVOLUTION LOCATION to length aspect ratio (fig. 4(b)). Contours were generated on an imaginary flat plate and then computationally faired onto the baseline afterbody shape.
NONE
All nozzle afterbodies had a convergent-divergent i nt e rn a l g eo m et r y w it h a n e xp a ns i o n r at i o o f
FORWARD
A /A = 1.397 and a design nozzle pressure ratio (NPR) of e t 5.4. Drawings of the internal moldlines are shown in figure 5. The convergent section occurred for x ≤ 3 inches in the xy plane (fig. 5(a)) and was defined with
MID
curve 2 (eqn. 1).
3 2 y = 0.093037 x - 0.418667 x + 1.756 (1)
LONG
The internal geometry was designed with the divergent section in the xz plane to accommodate the depth of the Figure 3. Matrix of boattail angle and convolution convolution valley in the xy plane. The divergent section location.
was dependent on the nozzle length L and was defined with curve 3 for x ≥ 3 inches (fig. 5(b)) . Equations 2 and 3 define curve 3 for the 15° and 22° boattail angle nozzle afterbodies, respectively.
3 2 z = -0.0011509 x + 0.0313863 x - 0.157244 x + 2.70633 (2) American Institute of Aeronautics and Astronautics AIAA 99-2670 Instrumentation An internal six-component strain-gauge balance was used to measure external and internal forces and moments acting on the model downstream of MS 26.9 (fig. 2).
Weight flow was measured with a multiple-critical venturi located in the high-pressure air system. Jet total pressure and total temperature were measured with a ten- probe total pressure rake and an iron-constantan (a) Spanwise profile thermocouple located in the instrumentation section. Jet total pressure and venturi static pressures were measured with individually sized transducers. Nozzle internal and external static pressures were measured with electronically-scanning pressure (ESP) modules located in the model forebody.
The convoluted afterbodies had three longitudinal, (b) Streamwise profile rows of external pressure orifices, one located near the sidewall, one near the centerline on the top of a Figure 4. Spanwise and streamwise convolution profiles. convolution, and one near the centerline in the valley between two convolutions. The last pressure tap in each row, x/L = 1, was located in the aft-facing, base end region. Baseline nozzles had two rows of longitudinal y pressure orifices, one located near the sidewall and one at L the centerline. All nozzles had seven internal pressure orifices.
3.000 Angle of attack was measured with an accelerometer 1.756 0.500 x in the strut head and corrected for sting deflections and 0.040 TE # E V tunnel flow angularity. The average flow angularity R C U measured in the tunnel was 0.1° of upflow.
3.000 Data Acquisition and Reduction (a) Side View All data for the wind tunnel parameters and the model were recorded simultaneously. Steady-state data L were obtained by averaging 50 frames of instantaneous 3.000 data sampled at a rate of 10 Hz. Final force and moment data were obtained by correcting each measured balance component for model weight tares, for balance component interactions, and for jet-off balance interactions. A 7.252 detailed description of the data reduction procedure is x given in reference 7.
4.972 Data were taken with increasing nozzle pressure CU R V E # ratio (NPR), which is the ratio of jet total pressure p to t,j 0.100 free-stream static pressure p . Balance-measured, thrust- ∞ removed, total drag coefficient C was comprised of D z pressure drag coefficient C and skin friction drag D,p (b) Top View coefficient C . Skin friction drag was computed by the D,f method of Frankl and Voishel for compressible, turbulent Figure 5. Nozzle internal moldlines.
flow on a flat plate (ref. 7). Balance-measured pressure drag was obtained by subtracting skin friction drag from American Institute of Aeronautics and Astronautics AIAA 99-2670 balance-measured, total drag. Coefficients were calculated shallow 15° afterbody at all NPR for both subsonic (fig.
by normalizing drag force with free-stream dynamic 6(b)) and supersonic (fig. 7(b)) conditions.
pressure q and a reference area of 42.396 inches.
∞ Pressure coefficient ( C ) distributions are shown as a Pressure coefficient distributions along the p function of normalized axial location x/L , where pressure centerline of the 15° and 22° baseline afterbodies for coefficient is defined as the difference between surface various NPRs at M = 0.8 are shown in figure 8. The static pressure p and free-stream static pressure, nozzle afterbody starts at MS 54.486, which is defined as normalized by the free-stream dynamic pressure x/L = 0. The pressure tap located at x/L = 1 was not an afterbody static pressure tap, but was located in the aft- Flow Visualization facing, base end of the nozzle. The data indicates that the flow remained attached along the 15° boattail afterbody at Oil paint, thinned with linseed oil, was used to this Mach number. A strong shock, followed by shock- visualize flow patterns along the nozzle surfaces. It induced separation near x/L = 0.7, occurred along the provided an excellent aid for interpreting pressure and steep 22° boattail afterbody. The 22° baseline afterbody force data. A row of paint dots, alternating in color, were produced more drag because the steeper boattail resulted in applied normal to the flow direction at the nozzle connect more expansion near x/L = 0.5 and pressures were lower station. Alternating the paint colors provided easier on the aft-facing afterbody, compared to the 15° baseline detection of streamline patterns. The tunnel set condition afterbody.
was held for approximately 2-3 minutes to allow the flow to move and dry the paint along the streamline path. Pressure coefficient distributions along the Photographs were taken after the tunnel was shut down. centerline of the 15° and 22° baseline nozzle afterbodies for various NPR at M = 1.2 are shown in figure 9.
Test Schedule Compared to the M = 0.8 case, the pressure coefficient was lower for both boattail angles over the last 40 percent Data were acquired at Mach numbers of 0.6, 0.8, of the afterbody at supersonic conditions, which resulted 0.9, 0.95, and 1.2, and model angles of attack ( a ) of -5°, in higher drag. Based on the pressure data in figure 9, it 0°, 5° and 10°. Data were taken at a constant Mach appears that the flow had enough momentum to remain number and a , while varying nozzle pressure ratio from 1 attached to the shallow, 15° afterbody until x/L = 0.88, (jet-off) to 12. resulting in flow separation over the last 12 percent of the afterbody. The flow along the steeper, 22° afterbody appeared to separate near x/L = 0.7. Drag was higher on RESULTS the 22° afterbody than on the 15° afterbody because the pressures were much lower over a larger region of the boattail (30 percent compared to 12 percent). In addition, Data are presented with effect of boattail angle along the steep 22° boattail had more aft-facing area for the the baseline nozzle afterbodies first, followed by the pressures to act in the axial direction. Therefore, the effects of adding convolutions to the 22° and the 15° convolutions had more of a chance to effect the adverse boattail angle afterbodies. All data presented were flow conditions along the 22° baseline afterbody than the acquired at zero degree angle of attack.
15° baseline afterbody.
Baseline Comparisons The effects of nozzle pressure ratio on total drag are greater on the 15° afterbody than on the 22° afterbody at The 15° baseline nozzle afterbody had less total drag M = 1.2 (fig. 7). The pressure data indicates larger at both M = 0.8 and 1.2 than the 22° baseline nozzle differences in pressure recovery along the aft end of the afterbody, as shown in figures 6 and 7, respectively. The 15° afterbody, especially near x/L = 0.8, compared with 15° baseline nozzle afterbody was 15.18 in. long and had the 22° afterbody separation. The shock located more wetted surface area than the 22° baseline nozzle downstream of x/L = 0.7 appeared to get stronger (higher afterbody, which was 12.09 in. long. Therefore, the 15° pressures) as NPR increased, which resulted in a decrease baseline afterbody produced more skin friction drag than in drag (fig. 7).
the 22° baseline afterbody at all Mach numbers (see Table 2). After removing the contribution of skin friction drag Streamline patterns along the 22° baseline afterbody from total drag, it is obvious that the shorter, steeper 22° at M = 0.8 and M = 0.95 for a NPR = 7 are shown in afterbody produced more pressure drag than the longer, figures 10 and 11, respectively. The paint covered only 70 percent of the boattail, indicating massive separation American Institute of Aeronautics and Astronautics AIAA 99-2670 over 30 percent of the afterbody at both Mach numbers. aft 20 percent of the boattail at M = 0.8 (fig. 19(a)), Even though the streamline patterns look remarkably which resulted in thrust on the aft-facing boattail. The similar, total drag was 38 percent higher at M = 0.95. convoluted contouring delayed separation at M = 0.95, The difference in nozzle drag can be explained with the but the expansion near x/L = 0.6 was greater and the pressure distributions along the afterbody. The pressure shock was stronger relative to the baseline (fig. 20), coefficient distributions along the 22° baseline afterbody resulting in increased drag. The flow did not recover to centerline and sidewall for M = 0.8 and M = 0.95 at an positive values of pressure coefficient along the 22° NPR = 7 are shown in figure 12. The shock was stronger forward convoluted afterbody as it did under M = 0.8 at M = 0.95, as seen by the quick, steep increase in conditions. The last pressure tap was in the trailing edge pressure near x/L = 0.7. After the flow expanded at x/L = surface, and was pressurized by the internal flow at 0.55, the flow attempted to adjust to ambient conditions, underexpanded conditions.
but separated from the afterbody downstream of the shock near x/L = 0.8 for M = 0.8 and near x/L = 0.7 for M = Convoluted contouring was not effective at 0.95 conditions. Drag was higher at M = 0.95 because decreasing total drag at supersonic conditions because of the pressures were lower than the M = 0.8 case for x/L > wave drag (fig. 16). At supersonic conditions the 0.7. In addition, pressures acting on the last 50 percent convolutions created shock waves that reduced total of the afterbody have more impact on drag than x/L < 0.5 pressure, increased entropy, and inevitably, increased drag.
because of the larger aft-facing area. As shown in figures 14 and 16 at NPR = 5.4, the forward convolutions decreased drag 5 percent at M = 0.9, but 22° Boattail Angle, Convoluted Afterbody increased drag nearly 20 percent at M = 1.2. The pressure coefficient distributions along the 22° baseline and forward convoluted nozzle afterbodies at M = 1.2 and The effect of convoluted contouring on the total NPR = 7 are shown in figure 21. The convolutions drag of the 22° boattail angle afterbody, at freestream caused pressurization on the forward-facing surfaces Mach numbers of M = 0.8, 0.9, 0.95, and 1.2, is shown between x/L = 0.2 and 0.4, increasing drag relative to the in figures 13-16, respectively. Results indicate that the baseline. Similar to the M = 0.95 case (fig. 20), the effectiveness of convoluted contouring on the afterbody convolutions delayed separation, but caused the expansion boattail is dependent on freestream Mach number, NPR, of the flow along the centerline at x/L = 0.7 to reach and convolution location. None of the convolution lower C , which resulted in more drag than the baseline locations reduced drag relative to the baseline at M = 1.2.
p afterbody produced. The flow on both the baseline and The forward convolutions were the most effective at forward convoluted afterbody separated with negative reducing drag for M < 0.95. At M = 0.8, the forward values of pressure coefficient (drag on the aft-facing convolutions reduced total drag nearly 20 percent at the surface).
design nozzle pressure ratio, NPR = 5.4. In general, the D long convolutions increased skin friction and pressure 15° Boattail Angle, Convoluted Afterbody drag relative to the baseline afterbody for M > 0.8.
The effect of convoluted contouring on the total Convoluted contouring in the forward location drag of the 15° boattail angle afterbody, at freestream energized the boundary layer and the flow remained Mach numbers of M = 0.6, 0.9, and 1.2, is shown in attached further downstream on the afterbody. The delay figures 22-24, respectively. The convolutions were not of separated flow was "visualized" with oil paint, and the effective on the 15° afterbody. As expected, the results are shown in figures 17 and 18, for an NPR = 7, convolutions increased skin friction drag (see Table 2), at M = 0.8 and 0.95, respectively. The streamlines on but also increased nozzle pressure drag at all Mach the baseline afterbody (figs. 10 and 11) indicated more numbers. As with the 22° convoluted nozzle afterbodies separation than on the forward convoluted afterbody at at M = 1.2, the 15° convoluted nozzle afterbodies had both Mach numbers (figs. 17 and 18). However, greater total drag than the baseline due to wave drag.
compared to the baseline afterbody, the forward After removing the contribution of skin friction drag from convolutions reduced drag 36 percent at M = 0.8, but the measured total drag, it is obvious, as shown in figure increased drag 10 percent at M = 0.95. The pressure 25, that pressure drag was higher for the convoluted cases coefficient distributions in figures 19 and 20 for M = 0.8 than for the baseline afterbody. As discussed previously, and M = 0.95, help to illustrate the Mach number the 15° baseline afterbody had much less, and in some dependence of the forward convolutions. The forward cases, no separated flow to influence relative to the 22° convolutions energized the flow to allow pressure baseline afterbody.
recovery to positive values of pressure coefficient over the American Institute of Aeronautics and Astronautics AIAA 99-2670 0.2 0.2 15° 0.1 C 0.15 0 p -0.1 C 0.1 D -0.2 15° NPR=3.11 15° NPR=5.452 -0.3 15° NPR=7.975 22° NPR=3.018 0.05 -0.4 22° NPR=5.375 22° NPR=7.991 22° 15° Baseline -0.5 22° Baseline 0 -0.6 0 2 4 6 8 10 0 0.2 0.4 0.6 0.8 1 1.2 NPR x/L (a) Total drag coefficient Figure 8. Centerline pressure coefficient distributions, 15° and 22° baseline afterbodies at selected NPR, M = 0.8.
0.2 0.2 0.1 0.15 15° 15° NPR=2.025 C p 15° NPR=4.008 15° NPR=5.446 C -0.1 0.1 D,p 15° NPR=7.009 15° NPR=9.009 -0.2 22° NPR=2.01 22° NPR=4.047 -0.3 0.05 22° NPR=5.396 22° NPR=6.993 15° Baseline -0.4 22° Baseline 22° NPR=8.963 22° -0.5 0 2 4 6 8 10 -0.6 NPR 0 0.2 0.4 0.6 0.8 1 1.2 x/L (b) Pressure drag coefficient Figure 9. Pressure coefficient distributions along the Figure 6. Baseline nozzle afterbodies, M = 0.8. centerline, 15° and 22° baseline afterbodies, M = 1.2.
0.4 15° Baseline 22° Baseline 0.35 C 0.3 D 0.25 0.2 0 2 4 6 8 10 12 14 NPR (a) Total drag coefficient Figure 10. 22° Baseline afterbody at M =0.8, NPR = 7.
0.4 15° Baseline 22° Baseline 0.35 C 0.3 D,p 0.25 0.2 0 2 4 6 8 10 12 14 NPR (b) Pressure drag coefficient Figure 11. 22° Baseline afterbody at M =0.95, NPR = 7.
Figure 7. Baseline nozzle afterbodies, M = 1.2.
American Institute of Aeronautics and Astronautics AIAA 99-2670 0.2 0.3 0.1 C 0 0.25 p -0.1 C -0.2 0.2 D -0.3 -0.4 0.15 Baseline Forward M=0.8 Mid -0.5 M=0.95 Long -0.6 0.1 0 0.2 0.4 0.6 0.8 1 1.2 0 2 4 6 8 10 12 14 x/L NPR (a) Centerline Figure 15. Effect of convolutions on total drag, 22° boattail angle, M = 0.95.
0.2 0.1 C 0.4 p -0.1 0.35 -0.2 C D -0.3 0.3 -0.4 M=0.8 Baseline -0.5 M=0.95 0.25 Forward Mid Long -0.6 0 0.2 0.4 0.6 0.8 1 1.2 0.2 x/L 0 2 4 6 8 10 12 14 NPR (b) Sidewall Figure 16. Effect of convolutions on total drag, 22° Figure 12. 22° Baseline comparison, M = 0.8 and M = boattail angle, M =1.2.
0.95, NPR=7.
0.2 0.15 C D 0.1 Baseline 0.05 Forward Mid Long 0 2 4 6 8 10 12 14 NPR Figure 13. Effect of convolutions on total drag, 22° Figure 17. Forward convolutions delay separation on the boattail angle, M = 0.8.
22° afterbody at M = 0.8, NPR = 7.
0.2 0.15 C D 0.1 Baseline 0.05 Forward Mid Long 0 2 4 6 8 10 12 14 NPR Figure 14. Effect of convolutions on total drag, 22° Figure 18. Forward convolutions delay separation on the boattail angle, M = 0.9.
22° afterbody at M = 0.95, NPR = 7.
American Institute of Aeronautics and Astronautics AIAA 99-2670 0.4 0.4 22° Baseline 0.2 0.2 22° Forward, Hill 22° Forward, Valley 0 0 C p C -0.2 -0.2 p -0.4 -0.4 22° Baseline -0.6 -0.6 22° Forward, Hill 22° Forward, Valley -0.8 -0.8 0 0.2 0.4 0.6 0.8 1 1.2 0 0.2 0.4 0.6 0.8 1 1.2 x/L x/L (a) Centerline (a) Centerline 0.4 0.4 22° Baseline 0.2 0.2 22° Forward 0 C 0 p C p -0.2 -0.2 -0.4 -0.4 -0.6 -0.6 22° Baseline 22° Forward -0.8 -0.8 0 0.2 0.4 0.6 0.8 1 1.2 0 0.2 0.4 0.6 0.8 1 1.2 x/L x/L (b) Sidewall (b) Sidewall Figure 19. Forward convolutions delay separation on the Figure 21. Pressure coefficient distributions along the 22° afterbody at M = 0.8, NPR=7. 22° baseline and forward convolution afterbodies, M = 1.2, NPR=7.
0.2 0.2 Baseline Forward 0 Mid 0.15 Long -0.2 C p C D 0.1 -0.4 0.05 22° Baseline -0.6 22° Forward, Valley 22° Forward, Hill -0.8 0 0 0.2 0.4 0.6 0.8 1 1.2 0 2 4 6 8 10 NPR x/L (a) Centerline Figure 22. Effect of convolutions on total drag, 15° boattail angle, M = 0.6.
0.2 0.2 Baseline Forward Mid 0.15 C Long -0.2 p C D 0.1 -0.4 -0.6 22° Baseline 0.05 22° Forward -0.8 0 0.2 0.4 0.6 0.8 1 1.2 x/L 0 2 4 6 8 10 12 14 NPR (b) Sidewall Figure 23. Effect of convolutions on total drag, 15° boattail angle, M = 0.9.
Figure 20. Forward convolutions delay separation on the 22° afterbody at M = 0.95, NPR = 7.
American Institute of Aeronautics and Astronautics AIAA 99-2670 22° afterbody, but the convolutions were only able to 0.4 reduce drag for M < 0.95. At M = 0.8, drag was reduced Baseline Forward 20 and 36 percent at NPRs of 5.4 and 7, respectively.
Mid 0.35 Long Drag was increased 10 percent for M = 0.95 at NPR = 7.
C In either case, the convolutions delayed separation, but D 0.3 only when the pressure on the boattail recovered to higher v al ue s t ha n t he b as el in e, w as t he d ra g r ed uc ed . At 0.25 M = 0.8, pressure recovered to positive values of pressure coefficient on the aft-facing boattail, which resulted in a 0.2 0 2 4 6 8 10 12 14 thrust component over part of the afterbody, or drag NPR reduction.
Figure 24. Effect of convolutions on total drag, 15° boattail angle, M = 1.2.
None of the convolution locations were effective at decreasing drag on the 22° afterbody for M > 0.95, due to 0.4 supersonic wave drag. Shocks formed on the forward- Baseline facing contours at M = 1.2, increasing pressure drag Forward 0.35 Mid relative to the baseline. In addition, the contouring Long delayed separation, but the flow expanded to even lower C 0.3 D,p pressures than on the baseline, increasing drag further.
0.25 REFERENCES 0.2 0 2 4 6 8 10 12 14 NPR 1. Corson, B. W., Jr., and Runckel, J. F.: Figure 25. Effect of convolutions on nozzle pressure Exploratory Studies of Aircraft Afterbody and drag, 15° boattail angle, M = 1.2.
Exhaust-Nozzle Interaction . NASA TM X-1925, 1969.
2. Runckel, J. F.: Interference Between Exhaust System and Afterbody of Twin-Engine Fuselage CONCLUSIONS Configurations . NASA TN D-7525, 1974.
3. Hunter, C. A.: An Experimental Analysis of Adding convolutions to the baseline nozzle Passive Shock-Boundary Layer Interaction Control afterbodies increased wetted area and therefore, increased for Improving the Off-Design Performance of Jet skin friction drag, one component of total drag. The Exhaust Nozzles . Master of Science Thesis, George effectiveness of the convolutions at decreasing pressure Washington University/NASA Langley Research drag, the other component of total drag, was convolution Center, September 1993.
location, Mach number, boattail angle, and NPR 4. Presz, W. M., Jr.; Werle, M.; and Paterson, R.
dependent.
W.: Trailing Edge Separation/Stall Alleviation .
Journal of Propulsion and Power, Volume 25, 1987.
The 22° baseline afterbody had more separation 5. Presz, W. M., Jr. et al.: Rippled Afterbody along the boattail (due to the short, steep boattail Performance Study . United Technologies Research geometry) than the longer, more gradual 15° baseline Center Report UTRC87-27, September 1987.
afterbody. At M = 1.2, the flow was separated over the 6. Capone, Francis J.; Bangert, Linda S.; Asbury, last 30 percent of the 22° baseline afterbody, and over the Scott C.; Mills, Charles T.; and Bare, E. Ann: The last 12 percent of the 15° baseline afterbody. Therefore, NASA Langley 16-Foot Transonic Tunnel - the convolutions had an opportunity to influence a larger Historical Overview, Facility Description, region of separated flow on the 22° boattail angle Calibration, Flow Characteristics, Test afterbody. In fact, none of the convolution locations were Capabilities. NASA TP-3521, 1995.
effective at decreasing drag on the 15° boattail angle 7. Mercer, C. E., Berrier, B. L., Capone, F. J., and afterbody at any Mach number.
Grayston, A. M.: Data Reduction Formulas for the 16-Foot Transonic Tunnel at NASA Langley The forward convolution location was the most Research Center, Revision 2. NASA TM-107646, effective contouring geometry for drag reduction on the 1992.
American Institute of Aeronautics and Astronautics