PSP-Based Experimental Investigation of a Blended Wing ... PSP Blended Wing.pdf · The aerodynamic characteristics of a blended wing body aircraft, constructed using selective laser

AIAA 2005-4719 This material is a work of the U.S. Government and is not subject to copyright protection in the United States.

American Institute of Aeronautics and Astronautics

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PSP-Based Experimental Investigation of a Blended Wing Body Aircraft

David A. Gebbie * and Mark F. Reeder† Air Force Institute of Technology, Wright-Patterson AFB, OH, 45433

Charles Tyler,‡ Air Force Research Laboratory (AFRL/VAAC), Wright-Patterson AFB, OH, 45433

and

Vladimir Fonov# & Jim Crafton§, Innovative Scientific Solutions, Inc., Dayton, OH, 45440

The aerodynamic characteristics of a blended wing body aircraft, constructed using selective laser sintering (SLS), were assessed in the AFIT low-speed wind tunnel. The scaled-down model (Rec ~ 105 and M = 0.10 to 0.20) of a strike tanker consisted of a shaped fuselage and sweptback wings. The model evaluation and analysis process included force and moment measurements acquired from a wind tunnel balance, pressure sensitive paint (PSP) measurements, and a comparison to computational fluid dynamics (CFD) solutions. One of the most interesting results was the striking difference in the force and moment measurements before and after the paint was applied to the surface. The average surface roughness, Ra, was measured with and was found to have increased from approximately 0.3µm to 0.7µm when the paint was applied. Although this value is below the threshold suggested by 2-D boundary layer theory, the effect was clear and repeatable. Force and moment coefficient data suggest that the onset of wing stall was sudden across the entire wing for sufficiently smooth cases but occurs gradually for rougher surfaces. Interestingly, the CFD results compared well with the experimental data corresponding to the measurements of the rougher, painted model. Both the PSP data and the CFD results indicated that the primary mechanism of lift transitioned to vortex lift at sufficiently high angles of attack. Taken as a whole, the data suggests that, for the range of conditions tested, submicron roughness has a pronounced effect on the transition to the vortex lift mechanism for increasing angle of attack.

Nomenclature b = wingspan CD = drag coefficient Cp = pressure coefficient CL = lift coefficient Cm = pitching moment coefficient CN = normal force coefficient c = chord Re = Reynolds number α = angle of attack

* Second Lieutenant, USAF, currently with AFRL/PRTT, WPAFB, OH, AIAA Member. † Assistant Professor, Department of Aeronautics and Astronautics, AFIT/ENY, AIAA Member ‡ Research Engineer, AIAA Member # Research Engineer, AIAA Member § Research Engineer, AIAA Member.



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I. Introduction n order to characterize and improve aircraft designs, it is imperative to quickly assess their aerodynamics. Traditionally, this has been accomplished by test article fabrication, wind tunnel testing, and prediction of

performance using scale-up rules. The challenges inherent to the accomplishment of this goal have often resulted in lengthy time requirements. In the past few decades, technological advancements have enabled faster and more accurate assessment using three specific methodologies: 1) rapid prototyping, 2) computational fluid dynamic (CFD) modeling, and 3) global measurements. One goal of the Air Force Research Laboratory Air Vehicle Directorate (AFRL/VA) is to integrate each of these areas with one another so that they function in a complementary manner.

One goal of the current study was to apply each of these rapid development technologies for a modern

unconventional, rather than a simplified, air frame model. To this end, an industrial air framer provided AFRL with detailed electronic model information to fully define the external strike tanker geometry, which is shown in Figure 1. A rapid prototype design methodology was developed and used to construct a stainless steel model. The model was constructed using selective laser sintering (SLS) of stainless steel, and several features were incorporated in order to improve the ease with which wind tunnel measurements could be completed. For example, built-in pressure taps improved the speed and accuracy of pressure sensitive paint (PSP) measurements. Additional details of the model construction procedure are given in references 1 and 2. This electronic model was also used as a starting point for constructing the computational grid, and part of project was devoted to comparing experimental and computational results.

The experimental portion of the program has taken place in the Air Force Institute of Technology (AFIT) low

speed wind tunnel while the initial computational results have been processed using resources at AFRL/VAAC. Force and moment coefficient measurements were obtained from a six component wind tunnel balance, and PSP measurements were acquired for the upper surface of the test article. Each of these measurements was carried out for a variety of airspeeds and angles of attack. Preliminary results for this model were reported in reference 1. Among the more interesting findings disclosed in that reference was that while the experimentally measured and computationally predicted lift slopes and drag coefficients generally correspond well for some conditions, there was an experimentally observed stall at angles of attack approaching 20 degrees whereas no stall was found in the initial CFD model. As a result, a portion of the current investigation was devoted to investigating this finding regarding stall.

II. Background The tested model is a variation of a BWB design in that the centerbody of the aircraft is a lifting surface. It is

envisioned that this type of aircraft could serve in military applications as a tanker with increased capacity, which could be launched with a strike force, eliminating the need to get tanker assets in place before launching bombers.3 There have been a number of studies performed on the blended wing body concept, though many of these are proprietary. Two exceptions are given in references 4 and 5, and each of these is pertinent to the current investigation.

Katz et al.4 (1999) summarized experimentally obtained force and moment coefficient measurements acquired

for a specific BWB design at an approximately Rec = 4*105. Tests included both wings-on and wings-off data. Their data indicated that the CL-α curve was linear up to approximately 5 degrees, but that a reduction in the slope at that angle existed due to wing stall. It is notable that the CL-α curve continued to increase throughout the range of the tests, which are shown up to 18 degrees. The L/D curve exhibited a peak of approximately 15 in their study, near the break in the CL-α curve. The pitching moment coefficient data with respect to alpha (Cm-α) showed a slight increase in the slope, consistent with loss of lift aft of the point where the moment coefficient was determined. At larger angles of about 10 degrees, there was a reduction in the slope of the Cm-α plot, consistent with a rearward shift in the center of pressure. The results presented correspond to one air speed.

A very recent study incorporating both CFD and experimental results for a BWB design are summarized in

reference 5. Pao et al. (2005) reported results acquired in the NASA Langley 14 by 22-Foot Subsonic Tunnel for a 33:1 scale model. The tests most pertinent to the current study were of a Mach number range similar to the current set of experiments, but the Reynolds number was much larger – in their case, Rec ranged from 2.6*106 to 107. An interesting finding was that the experimentally determined normal force coefficient, CN, for the 2.6*106 case exhibited a break near 22° while the 3.7*106 case showed a milder break in slope at a smaller α, approximately 20°.

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In a mild contrast to the results given by Katz, the Cm-α curve exhibited a sharp increase in slope near 10° followed by a decrease near 25°. Notably, the result indicating the downward shift in the apparent stall angle, indicated by the break in the CN-α curve as Reynolds number increases is counter to the result expected for a prototypical lifting body.6 Another important finding disclosed in this report is that the computational results for the CN-α curve closely matched the results until α increased above the break, at which point the computational results departed significantly from the experimental data. Those results were obtained using a one-equation Spalart-Allmaras turbulence model.

Surface pressure maps are often used to gain more understanding of the dynamics of the flow field and were

sought in the current study. Pressure sensitive paint (PSP) has complemented traditional tests using pressure taps, and the procedure for using PSP is well documented in the literature.7-8 There are a number of challenges associated with testing in a low speed environment where the sensitivity to the intensity change can be quite low.9

Several experiments which are of special pertinence to the current study were dedicated to the study of paint

intrusiveness in the testing environment. One of the beneficial aspects of PSP is that no additional tubes or wires are required for the tests. Rather, a paint coating is applied, and it typically very thin. While the coating thickness is generally negligible in comparison to the thickness of an aircraft component, the paint could conceivably be a significant source of intrusion if the surface roughness of the model is affected.

A comprehensive study of this effect was documented by Schairer et al.10, who tested a swept wing at high and

low Reynolds numbers. They found that at high Reynolds numbers, the roughness due to the paint coating changed the pressure distributions near stall and post-stall. However, at low Reynolds numbers the paint intrusiveness was quite small and indistinct. In each case, the pre-stall lift slope was generally unaffected by the roughness. This set of results was consistent with the classical admissible roughness criteria, which is based on two-dimensional boundary layer flow, as documented by Schlicting.11

Vanhoutte et al. conducted a low and high subsonic speed study on swept wings. The low-speed results were

conducted for Reynolds numbers were between 105 and 106. Their results indicated that on a high lift application at low Reynolds numbers, a rough application of PSP influenced the flow around the leading edge. In particular, Vanhouette et al. indicated that the paint roughness potentially influenced the behavior of the separation bubble, possibly removing it all together, and speculated that the roughness may have also induced cross-flow transition, which is a predominant issue in swept wing applications.12

Another study of paint intrusiveness was documented by Amer et al.,13 who included an investigation of a delta

wing geometry at Mach 0.2. Their results indicated that PSP had little influence on measured lift and drag values when the paint roughness corresponded closely to the roughness of the original model for the geometry and flow conditions of their tests, which were conducted for Reynolds numbers on the order of 107.

The more general topic of the effect of surface roughness on the transition to turbulence in swept-wing flows is

itself an area of current interest in the field of fluid dynamics. A comprehensive review of this subject is given in reference 14. One matter of significance to the current investigation is brought out by Radeztsky et al., who found through flow visualization that even at Reynolds numbers over 106, there were distinguishable differences in the transition characteristics of the flow depending on whether the measured roughness was 0.25 or 0.5 microns. This value is substantially lower than the admissible roughness criteria based on two-dimensional flow.15

III. Procedures and Methodology

A. Experimental Methods The rapid prototyping process of the strike tanker used in this experimental study began with a detailed

electronic model, as seen in below, provided to the University of Dayton Research Institute (UDRI), which was the lead organization for producing the prototype. The combined restrictions of the wind tunnel dimensions and balance range limited the wingspan of the model to 20.125 inches, which left the model at 1/72 scale. However, the limitations of current SLS equipment capped the maximum part size at 10 inches by 10 inches by 10 inches. The strike tanker model was subsequently fabricated in 6 pieces: forward and aft fuselage, left and right wings, and left and right tail fins. Ultimately, the model consisted of an assembly of the stainless steel parts shown in Figure 1.



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Care was taken to build in aft clearance for the wind tunnel mounting fixtures, and pressure taps with connecting tubes were built into each wing.

Once it was fully assembled, the surface was polished, and silicone rubber was used to fill seams between the

parts. The completed model is shown mounted on the balance in the wind tunnel in Figure 2 below. The strike tanker model used had a 20.125 inch wingspan with a 25 degree rearward sweep angle while the spilt v-tail had an overall span of 9 inches with the same sweep angle as the wing. The body of the strike tanker was 4 inches wide with an overall length of 16 inches with wing and tail chord lengths of 1.69 inches and 2.14 inches, respectively.

The model was mounted in the AFIT low speed wind tunnel test section of 31.5” (height) by 44” (wide) by 72” (long). The strike tanker span-to-tunnel width ratio (b/w) of 0.45 falls well under the acceptable ratio according to ref. 6 (b/w < 0.8). The Plexiglas side doors and top panel provided optical access. The model support sting enters the test section through a slot in the traverse plate at the bottom of the test section. The sting traverse system allows for angle of attack measurements from -25 degrees to +25 degrees, as well as sideslip angles from -20 degrees to +20 degrees. The sting mounted balance used to collect the force and moment data for the strike tanker was a one-half inch diameter Able Corporation Series D, MKII nominal 100-lbf, six-component internal strain gage balance, accurate to 0.25% of full capacity. The capacity specifications of the strain gage rosettes are listed below in Table 1, and the calibration was checked and confirmed prior to its installation. Additional details of the wind tunnel and mounting configurations are given in references 16 and 17.

Figure 2. Model configuration mounted in the AFIT low speed wind tunnel.

Figure 1. Model configuration is given in its formative phase (left) and as the model assembly. Pressure tap locations and nomenclature are noted.



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The force and moment signals were conditioned and collected using a National Instruments LabView Virtual

Instruments interface throughout the test program, and the force and moment coefficients were determined after each set of tests. The data was subsequently processed with an in-house Matlab code, which incorporated a model tare and implemented small corrections due to the wall effects and solid model blockages. Details of the data processing methodology are contained in references 16 and 17.

The force and moment coefficient results reported here encompass only tests performed using angle of attack

sweeps at zero yaw, collected at speeds of 60 mph, 90 mph, 110 mph, 130 mph, and 145 mph (26.8, 40.2, 49.2, 58.1, and 64.8 m/s, respectively). In addition, two angle of attack sweeps with the same range and increment were performed at yaw angles of +10 degrees and +20 degrees at 110 mph.

Pressure maps of the upper surface of the strike tanker were collected using PSP. The PSP setup in this

experimental study utilized a Bi-Luminophore PSP from Innovative Scientific Solutions Inc. (ISSI). Notably, no base coat was applied to the model for the series of tests described here, though future testing may incorporate a base coat in order to reduce noise in the PSP signal. This paint was chosen due to its low temperature sensitivity and its ability to provide compensation for the model displacement relative to the excitation light source, as well as the instability of the light source itself. The Bi-Luminophore PSP is very similar to Uni-FIB, a single layer paint composed of PtTFPP that uses a FIB binder; however, it uses a reference probe in addition to the signal probe. The bi-luminophore paint exhibits a pressure sensitivity of 6% psig and a response time of 0.3 sec and its emission spectrum ranges from 500 nm to 800 nm with a peak at 650 nm with 460 nm illumination at 20 degrees Celsius. The light sources used to illuminate the model for the strike tanker test condition were in the form of an array of four two-inch blue LED light sources emitting the at a wavelength of 405 nm. A matrix of more than 30 marker points was utilized to account for the spatial deformation of the model during the data processing.

The intensity images of the PSP during testing were captured using one of two lenses mounted to a 14-bit Cooke

PCO Series 1600 CCD camera linked up with ISSI’s image acquisition software, OMS Acquire. The Bi-Luminophore paint requires that two images be taken at each test condition. The use of a filter wheel is therefore necessary to capture two separate images under the same illumination conditions separated in time. The filter wheel used contained a 645-nm long pass filter for the signal probe and a 550 ± 40-nm band-pass filter for the reference probe. In addition to all of the PSP equipment used for the last set of experimental tests, all of the previously discussed wind tunnel equipment including the force and moment balance and pressure tap system was used. The pressure at the taps were collected using End tap equipment of special importance as it was used as a reference for comparison against the PSP data. For each measurements, the analysis began with the reference image that was acquired at ambient pressures, also known as a “wind-off” image. In addition, a “dark”, or background, image was acquired with the lighting system turned off to account for any variations in background room light. The two categories of “wind on” images (one for each filter) collected during for test were processed with reference to the dark image and the wind off image” using ISSI’s OMS Lite Version 1.0 software in order to obtain surface pressure plots. The images were aligned using the marker points within the images, and the resulting pressure maps were filtered and smoothed using a Gaussian filter with the maximum half width of the filter set to 10 pixels in each of the two axes. More details of the PSP system and data collection procedure are given in reference 17.

Normal Force - Total 100 lbf ± 0.25% Side Force- Total 50 lbf ± 0.25% Axial Force 50 lbf ± 0.25% Pitching Moment 52.5 in-lb ± 0.25% Rolling moment 15 in-lb ± 0.25% Yawing moment 25.5 in-lb ± 0.25%

0.50 Able Mark VI Balance Specifications

Table 1. Wind tunnel balance specifications.



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B. Computational Procedure CFD analyses were performed to determine the pressure distribution over the wing surface. The three-

dimensional solutions were obtained with a full Navier-Stokes code implementing the Spalart-Allmaras turbulence model. The CFD code is an unstructured, cell-centered, finite volume, Godunov-type solver that has least-squares gradient reconstruction and limiting for second-order spatial accuracy and first-order, point implicit time interpretation. The grid consisted of 501,300 cells: 372,472 tetrahedral and 128,828 prisms to acquire viscous boundary layer effects.

IV. Results

A. Force and Moment Coefficients Experimentally measured force and moment coefficients are given here for a variety of test conditions. A

number of surprising results affected the choices made within the data collection process, and for that reason, some reference is made to the chronology of experiments.

The initial set of results obtained for the model is given in Figure 1. This series of tests was conducted with the

intent of checking the strike tanker structural integrity and wind tunnel measurement system, and as such it was conducted for an unpainted model. The results were unanticipated in several respects. First, the apparent reduction in CL, hereafter referred to as the apparent stall event, was not predicted by CFD, nor was it indicated in the work described in reference 4.

The second surprising feature of the lift curve was the strong dependence of the reduction in lift on airspeed.

Both the magnitude of the apparent stall event and the angle at which it occurred were strongly affected by the airspeed with the angle gradually reduced as airspeed increased. This effect is the opposite of what one would expect for a typical lifting wing if Reynolds number effects are taken into account. However, it is noteworthy that this result is consistent with the findings of Pao et al. although it should be noted that this comparison is based on the authors’ presented data of normal force coefficients for only two different air speeds and Reynolds numbers.5

Thirdly, the drag coefficient did not show a dramatic increase, generally associated with stall. Rather the polar

plot showed simply a shift in the drag bucket from a more efficient to a less efficient outcome as angle of attack increased. Furthermore, both the lift and drag coefficients continued to increase with angle of attack, even after this stall event had taken place.

As a consequence of this result, it was deemed appropriate to check the data for repeatability. The model was

removed, and several months later, a few sets of data were acquired under the similar test conditions. The outcome for this repeatability check was consistent, and one set of repeatability tests is given here in Figure 4.

Given the repeatability of the results, it was deemed appropriate to continue testing with PSP, with focus given

to identifying the surface pressure features surrounding the stall event. The PSP images will be described in detail in the next section, but as it turned out, the force and moment coefficient data were affected by the application of the paint. Therefore, those results are given first.

It should be pointed out that the initial effort in the course of testing was focused neither on the intrusiveness of

PSP nor on the effects of model roughness on forces and moments. For this reason, no roughness measurements were taken for the original model. In retrospect, that type of measurement might have offered some additional insight into these results.



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Lift Coefficient Vs. Angle of Attack - Unpainted Model

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Lift Coefficient Vs. Drag Coefficient - Unpainted Model

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Figure 3. Lift coefficient data as a function of angle ofattack.



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A set of force and moment coefficient results collected for the painted model are given for the 110 mph test runs

in Figure 5 and for the 145 mph test runs in Figure 6. The results are compared to those for the original, unpainted model and the 110 mph results are further compared to the original CFD results for both a laminar simulation and for the aforementioned turbulence model. Another unexpected turn of events took place as it appeared that the paint application had apparently affected not only the characteristics surrounding the apparent stall event, but had even had some effect on the lift slope. It was some comfort that the new measurements were more closely matched to the CFD results and that the results for the painted model had similar characteristics to that reported in reference 4. Specifically, Katz et al. had reported a mild decrease in slope, which they attributed to wing stall rather than an abrupt reduction in CL.

Lift and Drag Coefficients Vs. Angle of Attack130 mph

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Lift Coefficient Vs. Angle of Attack - 145 mph

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Drag Coefficient Vs. Angle of Attack - 145 mph

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Lift Coefficient Vs. Drag Coefficient - 145 mph

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Figure 6. Lift coefficient data as a function of angle of attack.


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Drag Coefficient Vs. Angle of Attack - 110 mph

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Lift Coefficient Vs. Drag Coefficient - 110 mph

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Notably, the pitching moment coefficient results for the painted model were characteristically similar to those of

Katz et al for the painted model. By contrast, the results for the unpainted model, which showed a sharp increase at a specific angle of attack, were characteristically similar to that reported by Pao et al. in reference 5 for a different BWB model. Despite this relation to archival data, the difference in the character of the two set of results was somewhat disconcerting, especially since the lift slope rather than simply the stall event was apparently affected by the application of the paint. This sort of effect of paint roughness on the lift slope outside of the stall region was not reported in references 10 and 12 which were dedicated to quantifying PSP intrusiveness effects.

B. Surface Pressure Via PSP and CFD Despite the difference in the set of results for the painted and the unpainted model, the surface pressure map

obtained from the PSP measurements lent some insight into the generally phenomena in the flow field. PSP results are given here for the 110 mph and the 145 mph test cases. Note the slight difference in the color scale for each figure. As one would expect, a wider pressure range was measured for the higher airspeed condition. Additional results including results acquired at 2-degree increments for the same set of test conditions are given in reference 17.

Consistent with expectations, there was a reduction in the pressure near the leading edge of each wing for both

speeds up to and including the 8 degree case. Ambient pressure was 14.18 psia. For the 110 mph case, this reduction was from approximately 13.9 psia to 13.6 psia while the 145 mph case showed a reduction from approximately 13.6 to 13.2. One common result for both airspeeds was that the low pressure region near the leading edge of the wings was diminished between the 8-degree and 12-degree angle of attack test cases. Arguably, this low pressure region was slightly more intact at 12 degrees for 110 mph case than for the higher speed case. As the angle of attack increases toward 20 degrees, it is apparent that a low pressure region forms near the wing-body junction. Presumably, one might anticipate that a streamwise delta wing-type vortex system may have developed at these higher angles of attack, and this could be the primary lift mechanism for the vehicle at higher angles of attack.

It should be noted that the pressure taps, which were all located inboard of the wing body junction indicated a

monotonic decrease of pressure as angle of attack increased. Due to model geometry constraints, no pressure taps were located along the wings themselves. Rather the inboard taps served as a reference, and the paint was used entire to map the pressure in these regions. These contour plots contain additional information which might be useful in addition to the results concerning the wing and wing-body region. However, a modified camera angle, facing the tail, would likely improve the fidelity of the measurements. For that reason, discussion here is limited to the wing and wing-body regions only.

These pressure contour plots provide a level of understanding to the force coefficient data. With reference to

Figures 5 and 6, it can be seen that the CL-α for the painted model exhibits three approximately linear regions. The first region extends to approximately 6 degrees, the second from about 6 to 15 degrees, and the third being above about 15 degrees. These regions are slightly more pronounced for the 145 mph case, but are clear at each speed. A

Pitching Moment Coefficient Vs. Angle of Attack - 110 mph

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Pitching Moment Coefficient Vs. Angle of Attack - 145 mph

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Figure 7. Pitching moment coefficient data as a function of angle of attack for the unpainted and paintedmodels at 110 mph (left) and 145 mph (right).



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plausible explanation for this behavior is that in the intermediate region (6 to 15 degrees) a transition from a traditional lifting mechanism to vortex lift mechanism takes place.

Figure 8. PSP results at 110 mph. Scale is in psia.



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A small portion of the CFD results are compared to the global PSP measurements on a qualitative basis in Figure

10. The plots of surface pressure based on the CFD simulation employing the Spalart-Allmaras turbulence model showed a similar trend as the angle of attack was increased from 8 degrees to 14 degrees. The low pressure region near the leading edge of the wing diminished as a broader low pressure region near the wing body junction grows. It should be noted that the mesh spacing near the surface of the CFD model was on the order of 10 microns. As a

Figure 9. PSP results at 145 mph. Scale is in psia.



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result, any effect due to roughness below 10 microns would not be precisely modeled. In light of the experimental results which show a substantial influence due to the roughness height, it stands to reason that a finer mesh would be required to thoroughly capture the physics of the flow. This is the subject of planned future work.

Figure 10. Comparison of CFD and PSP results for varied angle of attack.



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C. Force and Moment Measurements Revisited: The Effect of Paint Removal and Model Polishing The surprising nature of the force and moment coefficient measurements mandated further study, and this was

completed in the form of additional force and moment data. First however, it was deemed appropriate to characterize the surface roughness of the painted model. If the roughness had exceeded the admissible level, one might have then expected the painted model to exhibit a difference. A series of linear surface roughness measurements were executed. The average Rt measured on the body was 6.814 µm while the Ra was measured at 0.66µm. Similar results were seen for the roughness measured on the wing which returned a Ra of 0.62 µm and an Rt of 6.51µm. Several additional profiles returned average values for Ra and Rt of 0.68 µm and 6.98 µm, respectively.

This level of roughness was compared to the admissible roughness for the Reynolds number of the experimental

tests using the approach outlined by Schairer et al., wherein a value of 6.2 times the Ra was recommended, based on a survey of different values in the literature. This value was found to yielded a characteristic roughness of 4.1 microns, which nearly an order of magnitude lower than the value of 18 microns (for Rec = 1.8*105). This indicates that the anticipated effect of roughness would be expected to be small in this case, counter to the actual results.

A series of measurements were then acquired for the model under different stages of paint removal with acetone

for a single air speed, 110 mph. The initial test of this nature was performed for the model with paint removed from the wing body junction only. Subsequently, measurements were acquired, with the model held in place on the sting, for full paint removal from the wings. Then paint was removed from the forward portion of the body. Finally, paint was removed from the tail region. The lift curves corresponding to each of these cases is shown in Figure 11. Careful scrutiny of this plot reveals that removing the paint from the wing body junction region had a fairly strong effect on the CL-α plot.

Initially it was surprising, and somewhat disappointing that when the paint was removed, the lift curve did not

return to its original level. However, there is precedence in the literature for changes in model surface characteristics due to wind tunnel testing. Specifically, reference 18 notes that sandblasting can occur when the model is tested for prolonged periods, as was the case for the PSP runs. Nevertheless, the lack of a repeatable outcome was bothersome.

For this reason, another set of data was collected for a polished model. The model was polished using an off-the

shelf stainless steel polishing solution, and the characteristic roughness was measured for this polished model and is given in comparison to the painted model in Figure 12. It is notable that the measured difference in roughness was slightly more than a factor of two with the polished model being smoother. The measured Ra was approximately 0.3 microns.


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Figure 11. Comparison of lift curves for the model with different paint coverage.



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The data was acquired for the polished model, and the results are given in Figure 13. Remarkably, the polished

model yielded a lift curve which exceeded the slope of the original model. Furthermore, the apparent stall event diminished for each speed. The lift curves for each speed were considerably more closely aligned, though with scrutiny, one can still observe a slight departure at the angles noted above for the original unpainted model.

Summarizing this set of results, paint was removed from the model in stages. The initial removal of paint from

the wing-body junction region led to a reappearance of the apparent stall found for the original model. However, this apparent stall event was apparently gradual in character, in contrast to that for the original unpainted model. As paint was removed from other regions of the model, the CL-α curve migrated toward the data recorded for the original model. However, even when the paint for the entire model was removed, the CL-α curve did not repeat its original data. Afterward, the model was removed and polished, its roughness was measured, and it was affixed to the balance. The polished model also did not repeat the original data. Rather the lift curve for the polished model increased to the point where the original data fell between that of the polished model and that of the model with all its paint stripped. This sequence of events affirmed the conclusion that, in fact, roughness was affecting the forces acting on the model. Additional details of the paint removal process is given in reference 17.

Surface finish

-3.0

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Figure 12. Comparison of sample profilometer output for the painted model and for the polishedmodel.



16

V. Discussion and Conclusions An experimental test program was undertaken as a part of a project focused on rapid design and assessment of

aircraft. The plan for the tests performed on the 72:1 scaled down blended wing body, strike tanker model in the AFIT low-speed wind tunnel consisted of the following elements: acquire force and moment data using a six-component balance; compare the data to CFD simulations of the model under identical conditions; and use a global measurement technique (PSP) to aid in discerning reasons for differences between the experiment and computation.

The testing covered air speeds ranging from 60 mph to 145 mph as described in previous chapters with a focus

on the 110 mph and 145 mph test speeds, which correspond to Mach numbers of 0.15 and 0.19, and Reynolds numbers of 1.35x105 and 1.79x105, respectively. The full experimental test matrix is given in reference 17. The initial measurements conducted for the model prior to the application of PSP showed substantial differences between the balance measured and CFD computed lift slopes. Moreover, an unexpected stall event was measured, which curiously revealed a strong dependence on air speed, despite the limited range of Reynolds numbers in the test. Although a similar dependence on air speed was indicated by the data given by Pao et al., the nature of this dependence in the current study appeared to be larger.

Subsequently, PSP was applied to the model to enable global pressure measurements with the intent that this

data could be used for comparison to the CFD model. Force and moment data was taken along with the PSP, but – based on the literature survey – the effect of the paint was expected to be minimal. However, upon inspection of the results of this thesis study, dramatic differences were seen in the measured force and moment data before and after paint had been applied to the model. The evidence of the change in the flight characteristics is prominent in all of the lift and drag coefficient comparisons, as well as the pitching moment stability for both the 110 mph and 145 mph test cases. The test conditions between the unpainted and painted model test were constant with the exception of the application of the paint to the model. The repeatability of the experiments was affirmed two days separated by several months was verified by the data given in Figure 3. At this point, a portion of this project was refocused on the reason why the paint would have such a profound effect. This led to the determination that the variation seen between the data for the unpainted and painted model was likely an effect of the strike tanker model surface roughness.

Subsequent tests of the model with its paint stripped and after polishing it provided evidence that the measured

roughness increase was the cause of the changes in the aerodynamic performance of the model. A profilometer was

C_L-alpha for the Strike Tanker at 110 mph

-0.2

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0

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Figure 13. Comparison of lift curves for the polished model with different paint coverage to other measured values, as indicated.



17

used to characterize the surface roughness of the model, and it was found that the lift curve was dramatically affected when the Ra increased from 0.32µm to 0.66µm, effectively doubling.

The research of surface roughness changes and their effects on aerodynamic performance and associated data is

quite extensive, and this body of research is very active. The general two dimensional boundary layer theories as described by Schlicting do not apply directly without some modification due to the crossflow effects on boundary layer transition generally found in swept wing flows. Some current studies performed in this very active research area have that indeed the admissible roughness is a poor criteria in such a circumstance. One example of this is given by the experiments of Radeztsky et al., which determined that micron sized roughness effects in swept wing applications had an effect on boundary layer transition. Their results showed that aerodynamic performance decreased as the surface finish of the stabilizer was improved.

Among the studies devoted to measuring the effects of roughness caused by PSP, none to the authors’ collective

knowledge, showed as dramatic of an effect as that reported here, which may be due in part to the fact that this study involves a complex BWB model. The PSP data indicated that there was a transition in the lift mechanism from a traditional lifting mechanism to a vortex lift mechanism for the painted model. The weight of evidence suggests that this transition is sensitive to surface roughness.

Interestingly, there are contrasting results for the BWB geometry in the literature, which could provide favorable

comparison to either the painted or unpainted model results.4,5 In the painted model tests, the lift curve slope was shown to steadily increase through the entire angle of attack range, with a slight break in the slope seen only in the unpainted model stall region. If one takes into account the differences in geometry, this result corresponds well to that of Katz et al., which indicates that this continuation of a positive lift curve slope is due to the lift created by the fuselage of the blended-wing body strike tanker. On the other hand the apparent stall observed for the unpainted model corresponds reasonably well to the results given by Pao et al., though there are differences, plausibly due to the different airframe geometries under consideration.

Despite the unexpected turns throughout the course of this study, the outcome does have some positive

implications on the integration of rapid technology approaches. Specific lessons here include the importance of utilizing a model not only with the appropriate geometry but also with the appropriate surface conditions for testing. In this case involving a complex model, the surface condition was more significant than one would have expected, based on the literature. A second lesson, specific to this case, is that a fully accurate CFD model would need to mesh volumes extremely close to the model surface in order to accurately account for surface effects. Each of these results completes a portion of the feedback loop into the CFD and Rapid Prototyping areas.

Acknowledgments The first and second authors would like to acknowledge the support of AFRL/VA. The authors would like to

acknowledge the support of Bill Braisted and James Higgins of UDRI in their extremely helpful work supporting the model construction. The authors would like to acknowledge the assistance and diligence of Mr. Dwight Gehring of the AFIT in coordinating and assisting with the wind tunnel measurements.

The views expressed in this article are those of the author and do not reflect the official policy or position of the

United States Air Force, Department of Defense, or the U.S. Government.

References

1Tyler, C., Braisted, W., Higgins, J., “Evaluation of Rapid Prototyping Technologies for Use in Wind Tunnel Model Fabrication,” AIAA Paper 2005-1301.

2Tyler, C., Reeder, M.F., Braisted, W., Higgins, J., AIAA Paper 2004-6870. (complete)

3Liebeck, “Design of the Blended Wing Body Subsonic Transport,” AIAA Journal of Aircraft, Vol. 41, No. 1, pp.

10-25, Jan. 2004.



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4Katz, J., Byrne, S. & Hahl, R. “Stall Resistance Features of Lifting-Body Airplane Configurations,” AIAA Journal of Aircraft, Vol. 26, No. 2, pp. 471-474, 1999.

5Pao, S.P., Bierdon, R.T., Park, M.A., Fremaux, C.M., & Vicroy, D. D., “Navier-Stokes Computations of

Longitudinal Forces and Moments for a Blended Wing Body,” AIAA Paper 2005-0045, 43rd AIAA Aerospace Sciences Meeting and Exhibit, January 2005.

6Barlow, Rae, and Pope, Low Speed Wind Tunnel Testing, 3rd ed., Wiley and Sons, 1999. 7Bell, J.H., Schairer, E.T., Hand, L.A., & Mehta, R., “Surface Pressure Measurements Using Luminescent

Coatings,” Annual Review of Fluid Mechanics, 2001, Vol. 33, pp. 155-206. 8Liu, T., Campbell, B.T., Burns, S.P., & Sullivan, J.P., “Temperature- and Pressure-Sensitive Luminescent

Paints in Aerodynamics,” Appl. Mech Rev., vol. 50, No. 4, April 1997. 9Bell, J.H., “Applications of Pressure-Sensitive Paint to Testing at Very Low Flow Speeds”, AIAA Paper 2004-

0878, January 2004.

10 Schairer, E.T., Mehta, R.D., Olsen, M.E., “Effects of Pressure-Sensitive Paint on Experimentally Measured Wing Forces and Pressures”, AIAA Journal, 2002, Vol. 40, No. 9, pp. 1830-1838.

11Schlicting, H., Boundary Layer Theory, 7th ed., McGraw Hill, 1979. 12Vanhoutte, F.G., Ashill, P.R., Gary, K.P., “Intrusion Effects of Pressure Sensitive Paint in Wind-tunnel Tests

on Wings”, AIAA Paper 2000-2525, June 2000. 13Amer, T.R., Obara, C.J., & Liu, T., “Quantifying the Effect of Pressure Sensitive Paint on Aerodynamic Data,”

AIAA Paper 2003-3951, 21st AIAA Applied Aerodynamics Conference, 23-26 June 2003. 14Reed, H.L., Saric, W.S., & Arnal, D., “Linear Stability Theory Applied to Boundary Layers,” AAnnu. Rev.

Fluid. Mech., Vol. 28, 1996, pp. 389-428. 15Radeztsky, R.H., Reibert, M.S., Saric, W.S., “Effect of Isolated Micron-Sized Roughness on Transition in

Swept Wing Flows”, AIAA Journal, Vol. 37, No. 11, November 1999, pp. 1370-1377. 16DeLuca, Anthony M., Experimental Investigation into the Aerodynamic Performance of Both Rigid and

Flexible Wing Structured Micro-Air-Vehicles, MS Thesis, AFIT/GAE/ENY/04-M06, Department of Aeronautics and Astronautics, Air Force Institute of Technology (AU), Wright Patterson AFB, OH, March 2004.

17Gebbie, D.A., “Experimental Study Of The Subsonic Aerodynamics Of A Blended Wing Body Air Vehicle

With A Focus On Rapid Technology Assessment,” MS Thesis, AFIT/GAE/ENY/04-M06, Department of Aeronautics and Astronautics, Air Force Institute of Technology (AU), Wright Patterson AFB, OH,, March 2004.

Documents

PSP-Based Experimental Investigation of a Blended Wing ... PSP Blended Wing.pdf · The aerodynamic characteristics of a blended wing body aircraft, constructed using selective laser