[American Institute of Aeronautics and Astronautics 29th AIAA Applied Aerodynamics Conference -...

American Institute of Aeronautics and Astronautics

1

Near Field Acoustic Test on a Low Boom Configuration in Langley’s 4x4 Wind Tunnel

Thomas R. Wayman1, Kenrick A. Waithe2, Donald C. Howe3

Gulfstream Aerospace Corporation, Savannah, GA, 31402, U.S.A.

Linda Bangert4, and Floyd Wilcox5

NASA Langley Research Center, Hampton, VA 23681, U.S.A.

Experimental aerodynamic data are presented for a wind-tunnel model of a low-boom aircraft concept conducted in the NASA Langley 4- by 4-foot Unitary Plan Wind Tunnel. This aircraft concept designed for experimental investigation and computational tool development was tested to generate a near field acoustic signature. Data have been obtained at freestream Reynolds numbers of 3.5x106

Nomenclature

per foot and Mach numbers of 1.50, 1.60 and 1.80 for the configuration. Force and moment, surface pressure and surface visualization data were obtained but the focus of this paper will be to describe the on-track and off-track near field acoustic signatures only.

AoA = Model angle of attack (degrees) BL = Boundary Layer Dewpt = test section dew point (°F) or humidity h = spacing between model and acoustic signature probes (inches) L = Model reference length (inches) M = Mach number p = test section static pressure (psf) p0

P01 (P = test section total (stagnation) pressure (psf)

01

P02 (P) = Survey probe pressure, off-track - y/L = 0.223 (psi)

02

P03 (P) = Survey probe pressure, on-track – y/L = 0.000 (psi)

03

P) = Survey probe pressure, off-track – y/L = 0.481 (psi)

ref

q = Reference Pressure (psi)

T= test section dynamic pressure (psf)

0

Re = Reynolds number based on unit length = test section total (stagnation) temperature (°F)

y = lateral displacement (inches) ∆p = Differential pressure (measured – reference) φ (phi) = Model roll angle (degrees)

1 Aero/Performance Specialist, Supersonic Technology Development, P.O. Box 2206 M/S R-07, AIAA Associate Fellow. 2 Applied Aerodynamics, Supersonic Technology Development, P.O. Box 2206 M/S R-07, AIAA Senior Member 3 Applied Aerodynamics Technical Fellow, Supersonic Technology Development, P.O. Box 2206 M/S R-07, AIAA Associate Fellow. 4 Aerospace Engineer, Configuration Aerodynamics Branch, Mail Stop 499, AIAA Associate Fellow. 5 Aerospace Engineer, Configuration Aerodynamics Branch, Mail Stop 499.

29th AIAA Applied Aerodynamics Conference27 - 30 June 2011, Honolulu, Hawaii

AIAA 2011-3331

Copyright © 2011 by Gulfstream Aerospace Corporation. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.


2

I. Introduction ulfstream Aerospace Corporation has, for a number of years, pursued a goal of developing the technologies necessary to facilitate the design of viable, commercial, over-land supersonic business-class aircraft. Both

aircraft-level and system-level computational and experimental investigations have been conducted and continue to be conducted in support of this activity.1,2,3,4,5

In 2008 a series of wind tunnel investigations were conducted to support computational tool development. The first investigation was a follow-on near field acoustic signature (NFAS) test conducted at the NASA Langley 4- by 4-foot Unitary Plan Wind Tunnel involving a simple wing-body-vertical configuration, while the former testing, also at Langley in 2002, was conducting on a sheared, faceted body only design. Its purpose was to also validate our computational tools by capturing near-field pressure (acoustic) signatures from a variety of area distributions including Gulfstream patented Quiet Spike technology. Also testing in 2008, the wing-body-vertical model was run in the NASA Ames 9- by 7-foot Unitary Plan Wind Tunnel as part of a larger experimental effort to evaluate sonic measurement techniques.

With a primary goal to design a quiet (relative to sonic boom characteristics) configuration, efforts have been made to design a set of experiments to test and measure the off-body acoustic field in order to support and validate our computational methods and tools. To that end, both computational and experimental analysis is being conducted to support quiet aircraft technologies.

6,7

II. Experimental Approach

With the larger sized tunnel, 9- by 7-foot vs. 4- by 4-foot, it was possible with the NASA Ames testing to obtain NFAS data beyond the near-field region of the model.

A. Wind Tunnel The wind tunnel test was conducted in the

Langley Unitary Plan Wind Tunnel (UPWT). The UPWT is a closed–circuit, continuous-flow, variable-density supersonic wind tunnel with two 4- by 4- by 7-foot test sections.8

Figure 1

Each test section is set up to cover a specific range of Mach numbers with an overlap between Mach 2.3 and 2.9. Test section no. 1, which was used for the testing described in this paper, has a Mach number range of 1.5 to 2.9. Using an asymmetric sliding-block-type nozzle for varying the ratio of nozzle throat to test section areas the tunnel is able to provide a continuous variation in Mach number ( ). Test section no. 1 has a total pressure capability up to 56 psia. The test gas is dried air with a temperature range of 125 to 175 degrees F. Tunnel dew point is maintained at sufficient levels to minimize vapor condensation effects. Table 1 summarizes the average free-stream conditions used during the testing in test section no. 1.

Table 1. Tunnel operating conditions.

Mach Re x 106 p

per ft psfa o T

°F o q

psfa p

psfa Dewpt °F

1.5 3.5 1839 125 789 501 -23.5 1.6 3.5 1888 125 796 444 -23.1 1.8 3.5 2020 125 797 352 -21.9

B. Model Description The Mach 1.6 wing-body-vertical model developed for this testing is designated as the NFAS model. It was

developed to provide validation for Gulfstream’s supersonic computational design methods and tools on lift generating three-dimensional aircraft-like configurations (Figure 2 and Figure 3). With a reference length of 13.2 inches, this model was sized to allow for near field acoustic signatures up to two (2) body lengths away from the model (h/L < 2.0) given the size of the unitary facility (4- by 4-foot).

G

Figure 1. NASA Langley 4- by 4-foot Unitary Plan Wind Tunnel – Test Section No. 1.


3

The model was designed and built from three separate parts: (1) Quiet Spike, (2) wing-body-vertical, and (3)

sting. Each part was designed to minimize the outer-mold line (OML) variances due to steps, gaps, and/or fasteners. As such the forward spike and sting were fastened from the top of the model minimizing potential pressure signature sources due to the model interface points. This approach provided a very ‘clean’ OML on the bottom of the model. Other than routing for static pressure tubing all model parts were cut from solid steel, 15-5PH H1015 alloy steel for all parts except the sting which utilized Vascomax 300-C alloy steel for strength and stability.

Like the Quiet Spike, the body is sheared to direct much of the shock structure off to the side and top of the model. This generates a keel line that is nearly flat over the length of the model. Both the vertical tail and wing leading edges attempt to tangentially blend into the body, with the wing planform taking on a swallow-like shape to reduce the acoustic signature at the trailing edge intersection with the body (Figures 2 and 3).

Like most efficient supersonic configurations abrupt changes in model cross sectional area are controlled in the vicinity of the aft body by necking down the body as the wing and vertical tail areas are introduced. Further necking results from the equivalent area generated by section lift. The result is a fairly narrow body not suited for more classic model mounting methods like a body sting or lower swept strut. Further complications result because of the need to measure both the forward and aft acoustic signatures on the NFAS model. As such a novel approach using an upper swept strut/sting not integrated into the vertical tail surface was introduced to mount the model and minimize signature inference.

The sting secured directly to the top of the body immediately forward of the vertical is swept at 65.5°. At the body, the sting interface consists of a contoured rectangular block functioning as a structural pad for model loads and as a cover plate for the static pressure tubing cavity. Flexible tubing inside the cavity provided the means of connecting the hard tubing inside the model and the hard tubing running up the sting. At the opposite end of the sting, before the interface with the sonic boom angle of attack mechanism (Figure 4), an integrated fixture was

Figure 3. Model as installed in 4- by 4-foot NASA Langley UPWT.

Figure 2. Wind tunnel model rendering with reference values.

L = 13.2”4.4”


4

included to allow for force (lift) and (pitching) moment measurement. Consisting of two rectangular shaped cross sections, set 5.5-inches apart, strain gages were applied on each fixture (top and bottom) to create a simple 2-component balance. In order to reduce interference to the expected sonic boom signature the sting cross sectional areas were tailored to correct for the loss of lift created by the interference between the sting and the model.9

Both free and forced boundary layer transition were allowed during this testing with free transition being used for the majority of testing. If boundary transition was forced, grit was placed on the model’s forward fuselage and wings. A #60 grit was used on the fuselage and a #80 grit was used on the wings. Grit was applied circumferentially on the fuselage 5-inches back from the leading edge, while the wing grit was placed 0.4-inches off of the wing leading edge.

At the side of body the sting initially has a chord of 2-inches and a leading edge half-angle of 5° on a modified biconvex cross section. Transitioning to an ovalized cross section, except at the gaged fixtures and the circular cross section where it interfaced with the sonic boom angle of attack mechanism, the sting progressively gets thicker once it no longer influences the model pressure signature.

C. Model and Instrumentation Installation Test section no. 1 was configured for either sonic boom signature measurements or schlieren flow visualization

by replacement of the west test section wall (Figure 4). Configured for sonic boom measurements the west test section wall was replaced by a solid steel door. For schlieren flow visualization the west test section wall was replaced by schlieren windows. The east test section wall remained unchanged with schlieren windows for the entire testing period.

The model support system provided gross level lateral displacement of the model relative to the west wall where

four (4) static pressure probes were installed (Figure 5). With the model rolled 90° toward the west wall the test set-up permitted pressure signature measurements up to 1.7 times the model reference length. The static pressure probes, one (1) reference and three (3) survey, were installed on a blade and tee-type struts respectively. The tee-type strut, mounted into a traversing mechanism, was free to move both axially (north and south) and laterally (east and west). Lateral movement was restricted to wind-off operation, while the axial motion was permitted during wind-on operation and remotely controlled from the facility control room. For each run the static pressure probes were fixed while the model support system traversed the model and associated pressure signature over the probes.

The reference probe is mounted above and slightly forward of the survey probes in order to avoid interference with the model shock structure (Figure 5). All of the probes were identical 2° cones with four (4) orifices drilled into a common line. The cross coupled probe pressures provided a level of insensitivity to probe misalignment,

Figure 4. Model general arrange in Test Section No. 1.

Angle of Attack Mechanism Roll Coupling

Model Support System

Survey Probes

Reference Probe Traverse

(West Wall)

(East Wall)

Top View

Figure 5. Survey probe layout.

Reference Probe

Survey Probe 1

Survey Probe 2

Survey Probe 3

Survey Probe Mechanism

(Test Section No. 1, west wall – solid steel door )


5

however because of the probes diminished diameter they were not robust to accidental contact. The survey probes remained in one position during an entire pressure signature run and would only be moved to

allow more or less separation distance as required between the model and the survey probes. The survey probes, with the model wing parallel to the west wall (φ = -90), are aligned either to measure on-track or off-track pressure signatures. Probe 2 (P02) lies on the model centerline (0°) and probes 1 (y/L = 0.223) and 3 (y/L = 0.481) falling off-track counter clockwise and clockwise respectively (looking upstream).

The model is attached to the sonic boom angle of attack mechanism which in turn is attached to a short sting. The short sting allows for mating to the facility roll coupling (Figure 4). The roll coupling is then attached to the tunnel model support system using a taper and draw nut arrangement. With a capability to move both laterally and axially, the tunnel support system was fixed during testing to hold the model level relative to the vertical plane. All model angle of attack movements were controlled in the sonic boom angle of attack mechanism, whereas roll movement was controlled by the roll coupling. The model was allowed to roll up to 90° to allow for a range of off track signature measurements and to allow for schlieren flow visualization.

D. Instrumentation and Measurements Force and moment data were measured using a 2-component strain-gage balance that was built into the sting

forward of the coupling with the sonic boom angle of attack mechanism. The gaged fixtures are shrouded to protect both the strain gages and the flexible surface pressure tubing used to gap the span of the fixture from the tunnel flow environment. Balance temperature was also measured at each of the gage sites on the sting, but the balance data was not corrected for variation in temperature. The thermocouples were used for monitoring purposes only.

Absolute pressure measurement of the reference probe utilized a 5 psia Druck pressure transducer. Differential pressure measurement of the survey probes utilized a Setra 0.09 psid transducer for each probe. Model surface pressure measurements also used 5 psia Druck pressure transducers. The reference probe pressure was connected to the reference sides of the Setra pressure transducers. The reference probe pressure was also connected via a bypass system to the sense side of the Setra pressure transducers. During tunnel startup and shutdown a shut-off valve was opened to prevent over pressurization in the bypass lines, then closed during data acquisition.

Data were scanned at a rate of 30 frames per second over a two-second period, and then averaged before data reduction. All the data described in this paper were obtained in a movement/pause mode of operation with an axial step size between data points of 0.125”. Allowing for movement, settling time, and data collection each test point took approximately 20 seconds and each signature approximately 45 minutes.

Data were collected to match pretest predictions on model angle of attack rather than model coefficient of lift. As detailed in a paper by Waithe9

E. Corrections

, the sonic boom signature is dependent on both changes in model cross sectional area and cross sectional lift coefficient. Initial test runs were used to characterize the relationship between model angle of attack and lift coefficient for the expected Mach numbers and Reynolds number. Based on this relationship angle of attack set points were used for the remainder of the testing to position and maintain the model for sonic boom signature measurements.

No corrections were made to the balance data. Pressure signatures were corrected to account for pressure offsets due to changing tunnel conditions.


6

III. Results and Discussion

A. Tunnel Stability Tunnel control and stability are important to any kind of wind tunnel testing, but with near field acoustic

signature tests control and stability become even more problematic. Static pressure measurements are typically less than ±1% of the freestream static pressure level, and as previously described take upwards of three-quarters of an hour to complete an entire signature. Stability of Mach number is a given, as would be expected, but circuit dew point also becomes a stability issue. This is complicated by the fact that most supersonic facilities rely on a limited amount of externally generated ‘dry’ air which in turn is depended on local atmospheric conditions. Meaning temporary and seasonal variation greatly affect the quality and availability of dry air. Fortunate for this test series was the fact that the testing was conducted during the winter months which are typically drier than other times of the year in Virginia.

During the course of each day of testing care was taken to control dew point, as result variation across a single signature were typically less than 5° degrees but from signature to signature the variation could be as large as 20 to 30° degrees F (Figure 6).

B. Force and Moment Designed for a specific coefficient of lift, the model was tested at over a range of lift coefficients to capture the

acoustic signature behaviors about the design point. Configured with the two-component (lift and pitching moment) balance, data were taken initially to characterize the sonic boom angle of attack mechanism’s ability to accurately hold and repeat angle of attack values consistent with the design point. This was accomplished by running two force and moment runs (Figure 7) with the model rolled level to the floor (φ = 0°) of test section no. 1 and then rolled parallel to the west wall (φ = -90°) of test section no. 1. All the acoustic signature testing would be performed with the model rolled -90°, while all the schlieren runs were performed with the model rolled 0°. The results show good agreement between the two roll conditions and provide a direct means of setting angle of attack to achieve the desired coefficient of lift.

Figure 6. Typical humidity stability over several acoustic signature runs.

-50

-25

0

10 15 20 25 30 35 40

Dew

Poi

nt [°F

]

Axial Probe Position [INCHES]

R134, H/L 1.20, M 1.6, Re 3.5E06, CL -0.36, phi -90R133, H/L 1.20, M 1.6, Re 3.5E06, CL 0.26, phi -90R135, H/L 1.20, M 1.6, Re 3.5E06, CL 0.68, phi -90


7

C. Schlieren Imagery Typical on-design schlieren imagery (Figures 8, 9, and 10) clearly shows the expected shock structures. The

multiple compressions and expansions can be seen from the Quiet Spike, with the effect of the sheared geometry highlighted in the stronger (lighter and darker) lines above the spike. Features like the body close out and the vertical tail closeout are also captured.

Figure 8. Typical acoustic signature physical associations.

-0.015

-0.01

-0.005

0

0.005

0.01

0.015

0.02

10 15 20 25 30 35 40

∆P 0

2/P r

ef


R133, H/L 1.20, M 1.6, Re 3.5E06, CL 0.26, phi -90

Figure 7. Configuration coefficient of lift.

-0.15

-0.1

-0.05

0

0.05

0.1

0.15

-2 -1.5 -1 -0.5 0 0.5 1 1.5 2

Coef

ficie

nt o

f Lift

[---

]

Model Angle of Attack [DEGREES]

R038, h/L ---, M 1.6, Re 3.5E06, phi 0

R039, h/L ---, M 1.6, Re 3.5E06, phi -90


8

Figure 10. Schlieren imagery, design point (M = 1.6, CL = 0.23) φ = -90°.

Figure 9. Schlieren imagery, design point (M = 1.6, CL = 0.23) φ = 0° .


9

D. On and Off Track Pressure Sonic boom pressure signatures to be presented capture variations in angle of attack (i.e., coefficient of lift),

model roll angle (φ), Mach number, and signature height. All the data presented will be for a single configuration and constant Reynolds number. Table 2 describes the extent of the data presented.

Table 2. Sonic Boom Signature Test Matrix.

Block Run Type Mach h/L AoA (°) φ (°) BL grit 1 AoA Sweep 1.6 --- -1.5 to 1.5 [-90, 0] off 2 Signature CL sweep 1.6 [1.2,1.7] [-0.36, 0.26, 0.68] -90 off 3 Signature Mach sweep 1.8 1.2 [-0.36, 0.26, 0.68] -90 off 4 Phi sweep 1.6 1.7 0.26 [-90, -69, -61, -45, -16.5] off 5 Grit 1.6 1.7 0.26 -90 [on, off]

An additional h/L condition (h/l = 0.5) was measured during the testing and is available, but the h/l = 0.5 data

will not be presented. The test set up for this NFAS measurement did not allow for a complete capture of the entire signature so that the available data did not capture the forward portion of the signature.

Typical acoustic signatures are presented in five (5) blocks: (a) On-design Mach number and h/L = 1.2 (Figures 11, 12, and 13), (b) On-design Mach number and h/L = 1.7 (Figures 14, 15, and 16), (c) Off-design Mach number and h/L = 1.7 (Figures 17, 18, and 19), (d) On-design Mach number and various roll angles (Figures 20, 21, 22, and 23), and (e) On-design Mach number with and without grit (Figures 24, 25, and 26). Consistent to each signature are a couple features that should be addressed because they play a role in how the plots were developed.

Distinct to the forward portion of the signature is a series of five compressions and five expansions that are generated by Gulfstream’s Quiet Spike. This feature is present in all the acoustic signatures and is independent of Mach number, angle of attack and roll angle. Following the initial series of compressions and expansion there is a region of compression generated by the area change in the fuselage and the lift generated by the wing. The eventual magnitude of the signature’s overall compression is controlled by the amount of lift being generated by the model and can be clearly seen in the signatures that span a range of lift coefficients. This is followed by a very large expansion coming off of the fuselage/wing intersection that quickly re-compresses, expands, and then compression back to the freestream conditions (Figure 8).

Because of the variability in controlling tunnel conditions (i.e., dew point), the length of time to acquire a single signature, and the relatively small magnitude of the pressure change it was necessary to offset each of the pressure signatures by a certain amount so that the ∆p prior to the signature would be zero. This was accomplished by taking an average of the first twenty (20) data points, then offsetting the whole signature by that amount. No attempt was made to adjust the signatures by also shifting the axial placement. This would be required to make comparisons of different h/L measurements in order to understand the changes in the signature (i.e., coalescing of the shock features) as it propagates away from the model.


10

Figure 12. Off-track (y/L=0.223), on-design (M=1.6), 1.2 Near Field Signature.

-0.015

-0.010

-0.005

0.000

0.005

0.010

0.015

0.020

10 15 20 25 30 35 40

∆P 0

1/P r

ef


R134, H/L 1.20, M 1.6, Re 3.5E06, CL -0.36, phi -90

R133, H/L 1.20, M 1.6, Re 3.5E06, CL 0.26, phi -90

R135, H/L 1.20, M 1.6, Re 3.5E06, CL 0.68, phi -90

Figure 11. On-track (y/L=0), on-design (M=1.6), 1.2 Near Field Signature.

-0.015

-0.01

-0.005

0

0.005

0.01

0.015

0.02

10 15 20 25 30 35 40

∆P 0

2/P r

ef


R134, H/L 1.20, M 1.6, Re 3.5E06, CL -0.36, phi -90

R133, H/L 1.20, M 1.6, Re 3.5E06, CL 0.26, phi -90

R135, H/L 1.20, M 1.6, Re 3.5E06, CL 0.68, phi -90


11

Figure 14. On-track (y/L=0), on-design (M=1.6), 1.7 Near Field Signature.

-0.015

-0.010

-0.005

0.000

0.005

0.010

0.015

0.020

20 25 30 35 40 45 50

∆P 0

2/P r

ef


R136, H/L 1.69, M 1.6, Re 3.5E06, CL -0.23, phi -90

R137, H/L 1.69, M 1.6, Re 3.5E06, CL 0.26, phi -90

R138, H/L 1.69, M 1.6, Re 3.5E06, CL 0.65, phi -90


-0.015

-0.010

-0.005

0.000

0.005

0.010

0.015

0.020

10 15 20 25 30 35 40

∆P 0

3/P r

ef


R134, H/L 1.20, M 1.6, Re 3.5E06, CL -0.36, phi -90

R133, H/L 1.20, M 1.6, Re 3.5E06, CL 0.26, phi -90

R135, H/L 1.20, M 1.6, Re 3.5E06, CL 0.68, phi -90


12


-0.015

-0.010

-0.005

0.000

0.005

0.010

0.015

0.020

20 25 30 35 40 45 50

∆P 0

3/P r

ef




-0.015

-0.010

-0.005

0.000

0.005

0.010

0.015

0.020

20 25 30 35 40 45 50

∆P 0

1/P r

ef


R136, H/L 1.69, M 1.6, Re 3.5E06, CL -0.23, phi -90

R137, H/L 1.69, M 1.6, Re 3.5E06, CL 0.26, phi -90

R138, H/L 1.69, M 1.6, Re 3.5E06, CL 0.65, phi -90


13

Figure 18. Off-track (y/L=0.223), off-design (M=1.8), 1.2 Near Field Signature.

-0.015

-0.010

-0.005

0.000

0.005

0.010

0.015

0.020

15 20 25 30 35 40 45

∆P 0

1/P r

ef



Figure 17. On-track (y/L=0), off-design (M=1.8), 1.2 Near Field Signature.

-0.015

-0.010

-0.005

0.000

0.005

0.010

0.015

0.020

15 20 25 30 35 40 45

∆P 0

2/P r

ef




14

Figure 20. Probe positions for various roll angles.

-2.0

-1.5

-1.0

-0.5

0.0

0.5

-1.0 -0.5 0.0 0.5 1.0 1.5 2.0 2.5

Non

-Dim

ensi

onal

Pro

be E

leva

tion

Pos

ition

Non-dimensional Probe Lateral Postion

R160, phi = -90.0R156, phi = -69.0R157, phi = -61.0R158, phi = -45.0R159, phi = -16.5h/L = 1.7

L


-0.015

-0.010

-0.005

0.000

0.005

0.010

0.015

0.020

15 20 25 30 35 40 45

∆P 0

3/P r

ef




15

Figure 22. Off-track (y/L=0.223), on-design (M=1.6), rolled near field signatures.

-0.020

-0.015

-0.010

-0.005

0.000

0.005

0.010

20 25 30 35 40 45 50

∆P 0

1/P r

ef


R160, H/L 1.70, M 1.6, Re 3.5E06, CL 0.27, phi -90

R156, H/L 1.70, M 1.6, Re 3.5E06, CL 0.27, phi -69

R157, H/L 1.70, M 1.6, Re 3.5E06, CL 0.24, phi -61

R158, H/L 1.70, M 1.6, Re 3.5E06, CL 0.22, phi -45

R159, H/L 1.70, M 1.6, Re 3.5E06, CL 0.27, phi -16.5

Figure 21. On-track (y/L=0), on-design (M=1.6), Rolled Near Field Signatures.

-0.020

-0.015

-0.010

-0.005

0.000

0.005

0.010

20 25 30 35 40 45 50

∆P 0

2/P r

ef


R160, H/L 1.70, M 1.6, Re 3.5E06, CL 0.27, phi -90

R156, H/L 1.70, M 1.6, Re 3.5E06, CL 0.27, phi -69

R157, H/L 1.70, M 1.6, Re 3.5E06, CL 0.24, phi -61

R158, H/L 1.70, M 1.6, Re 3.5E06, CL 0.22, phi -45

R159, H/L 1.70, M 1.6, Re 3.5E06, CL 0.27, phi -16.5


16

Figure 24. On-track (y/L=0), on-design (M=1.6), free verses forced boundary layer transition, CL = -0.23.

-0.010

-0.008

-0.006

-0.004

-0.002

0.000

0.002

0.004

0.006

0.008

0.010

20 25 30 35 40 45 50

∆P 0

2/P r

ef


R136, H/L 1.69, FREE transition, Re 3.5E06, CL -0.23, phi -90

R196, H/L 1.68, FORCED transition, Re 3.5E06, CL -0.27, phi -90

Figure 23. Off-track (y/L=0.481), on-design (M=1.6), rolled near field signatures.

-0.020

-0.015

-0.010

-0.005

0.000

0.005

0.010

20 25 30 35 40 45 50

∆P 0

3/P r

ef


R160, H/L 1.70, M 1.6, Re 3.5E06, CL 0.27, phi -90R156, H/L 1.70, M 1.6, Re 3.5E06, CL 0.27, phi -69R157, H/L 1.70, M 1.6, Re 3.5E06, CL 0.24, phi -61R158, H/L 1.70, M 1.6, Re 3.5E06, CL 0.22, phi -45R159, H/L 1.70, M 1.6, Re 3.5E06, CL 0.27, phi -16.5


17

Figure 26. On-track (y/L=0), on-design (M=1.6), free verses forced boundary layer transition, CL = 0.65.

-0.010

-0.008

-0.006

-0.004

-0.002

0.000

0.002

0.004

0.006

0.008

0.010

20 25 30 35 40 45 50

∆P 0

2/P r

ef


R138, H/L 1.69, FREE Transition, Re 3.5E06, CL 0.65, phi -90

R198, H/L 1.70, FORCED Transition, Re 3.5E06, CL 0.66, phi -90

Figure 25. On-track (y/L=0), on-design (M=1.6), free verses forced boundary layer transition, CL = 0.26.

-0.010

-0.008

-0.006

-0.004

-0.002

0.000

0.002

0.004

0.006

0.008

0.010

20 25 30 35 40 45 50

∆P 0

2/P r

ef


R137, H/L 1.69, FREE Transition, Re 3.5E06, CL 0.26, phi -90

R197, H/L 1.71, FORCED Transition, Re 3.5E06, CL 0.27, phi -90


18

IV. Summary A series of successful testing of a wing-body-vertical sonic boom aircraft model has been completed in the

NASA Langley 4- by 4-foot Unitary Plan Wind Tunnel. Data have been obtained at a Reynolds number 3.5x106

1. Mach number increase aggravated the signature levels by increasing absolute compression and expansion values by 30-40%.

per foot on a single configuration. Force and moment, surface pressure, schlieren imagery, and near field acoustic signature data were obtained but only force, schlieren imagery, and near field acoustic signatures are presented herein. Mach number, signature height, and boundary transition grit effects have been assessed.

2. Signature height change follows the expected trends of the Quiet Spike signature not coalescing but decreasing in magnitude, and the body closeout signature coalescing.

3. Addition of boundary layer trip, particularly on the wing, can be seen in the compression portion of the signature. It also can be seen to diminish the large expansion in the aft part of the signature.

Acknowledgments The authors would like to gratefully acknowledge the efforts of both the Gulfstream and NASA teams who

supported the computational and experimental campaigns. The Unitary Plan Wind Tunnel staff, particularly Rick Hall, is recognized for their group and individual efforts in overcoming a significant number of technical and mechanical hurdles during test preparation and conduct.

References 1 Henne, P., “The Case for Small Supersonic Civil Aircraft,” AIAA 2003-2555, 2003. 2 Wolz, R., “A Summary of Recent Supersonic Vehicle Studies at Gulfstream Aerospace,” AIAA 2003-0558,

2003. 3 Henne, P.A., Howe, D.C., Wolz, R.R., and Hancock, J.L., Gulfstream Aerospace Corporation, Savannah, GA,

U.S. Patent for a “Supersonic Aircraft with Spike for Controlling and Reducing Sonic Boom,” Patent No. US 6,698,684 B1, March 2, 2004.

4 Howe, D.C., "Improved Sonic Boom Minimization with Extendable Nose Spike", AIAA-2005-1014. 5 Howe, D.C., Simmons III, F., and Freund, D., “Development of the Gulfstream Quiet Spike for Sonic Boom

Minimization,” AIAA 2008-0124, 2008. 6 Durston, D. A., Cliff, S. E., Wayman, T. R., Merret, J. M., Elmiligui, A. A., and Bangert,L. S., “Near-Field

Acoustic Test on Two Low-Boom Configurations Using Multiple Measurement Techniques at NASA Ames,” AIAA 29th Applied Aerodynamics Conference, June 27-30, 2011 (to be published).

7 Cliff, S. E., Elmiligui, A. A., and Cambell, R., “Evaluation of Refined Tetrahedral Meshes with Projected, Stretched, and Sheared Prism Layers for Sonic Boom Analysis,” AIAA 29th Applied Aerodynamics Conference, June 27-30, 2011 (to be published).

8 Jackson, C. M., Corlett, W. A., and Monta, W. J., “Description and Calibration of the Langley Unitary Plan Wind Tunnel,” NASA Technical Paper 1905, November 1981. 9

Waithe, K. A., “Design of a Wind Tunnel Mound for a Low Boom Test,” AIAA 29th Applied Aerodynamics Conference, June 27-30, 2011 (to be published).

Date post:	14-Dec-2016
Category:	Documents
Upload:	floyd
View:	213 times
Download:	0 times

[American Institute of Aeronautics and Astronautics 29th AIAA Applied Aerodynamics Conference -...

Documents