June 2012 

NASA/TM–2012-217583 

Measuring Global Surface Pressures on a 

Circulation Control Concept Using Pressure 

Sensitive Paint 

Anthony Neal Watkins, William E. Lipford, and Bradley D. Leighty 

Langley Research Center, Hampton, Virginia 

Kyle Z. Goodman 

Analytical Mechanics Associates, Inc., Hampton, Virginia 

William K. Goad 

Jacobs Technology, Inc., Hampton, Virginia 
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June 2012 

NASA/TM–2012-217583 

Measuring Global Surface Pressures on a 

Circulation Control Concept Using Pressure 

Sensitive Paint 

Anthony Neal Watkins, William E. Lipford, and Bradley D. Leighty 

Langley Research Center, Hampton, Virginia 

Kyle Z. Goodman 

Analytical Mechanics Associates, Inc., Hampton, Virginia 

William K. Goad 

Jacobs Technology, Inc., Hampton, Virginia 
Available from: 

NASA Center for AeroSpace Information 

7115 Standard Drive 

Hanover, MD 21076-1320 

443-757-5802 

Acknowledgments 

The authors would like to thank the personnel at the National Transonic Facility for their 

assistance in this work, especially Scott Goodliff, who served as Lead Test Engineer. 

Funding for this work was provided by the Subsonic Fixed Wing project under the 

Fundamental Aeronautics Program. 

The use of trademarks or names of manufacturers in this report is for accurate reporting and does not constitute an 

official endorsement, either expressed or implied, of such products or manufacturers by the National Aeronautics 

and Space Administration.
3

Abstract

This report will present the results obtained from the Pressure Sensitive Paint (PSP) technique on a 

circulation control concept model. This test was conducted at the National Transonic Facility (NTF) at the 

NASA Langley Research Center. PSP was collected on the upper wing surface while the facility was 

operating in cryogenic mode at 227 K (-50 oF). The test envelope for the PSP portion included Mach 

numbers from 0.7 to 0.8 with angle of attack varying between 0 and 8 degrees and a total pressure of 

approximately 168 kPa (24.4 psi), resulting in a chord Reynolds number of approximately 15 million.

While the PSP results did exhibit high levels of noise in certain conditions (where the oxygen content of 

the flow was very small), some conditions provided good correlation between the PSP and pressure taps, 

showing the ability of the PSP technique. This work also served as a risk reduction opportunity for future 

testing in cryogenic conditions at the NTF.

I. Introduction

The accurate determination of spatially continuous pressure and temperature distributions on 

aerodynamic surfaces is critical for the understanding of complex flow mechanisms and for the 

comparison with computational fluid dynamics (CFD) predictions. Conventional pressure measurements 

are based on pressure taps and electronically scanned pressure transducers. While these approaches 

provide accurate pressure information, pressure taps are limited to providing data at discrete points.

Moreover, the integration of a sufficient number of pressure taps on a surface can be time and labor 

intensive and add significant expense to the model.

Using pressure sensitive paint (PSP) allows for the accurate determination of pressure distributions 

over an aerodynamic surface and is based on an emitted optical signal from a luminescent coating. As 

originally developed,1-3this technique was primarily useful for mean pressure measurements at transonic 

and higher flows, but has since been adapted to lower speed flows4-6as well as measurements of 

fluctuating pressures.7,8 A number of review articles cover the topic in detail.9-11 PSP measurements 

exploit the oxygen (O2) sensitivity of luminescent probe molecules suspended in gas-permeable binder 

materials. In wind tunnel applications, the PSP is applied to the model by conventional paint spraying 

techniques. Light sources such as UV LED arrays are mounted external to the test section to illuminate 

the painted model and effect luminescence emission from the entrapped oxygen-sensitive molecules. For 

the majority of pressure paints, PSP emission occurs in the red or orange region of the visible spectrum 

(~580 - 650 nm). The intensity is inversely proportional to the amount of oxygen present such that 

brighter regions in the paint emission indicate lower concentration of oxygen (and thus lower pressure) 

relative to the darker regions. Scientific-grade CCD cameras with spectral band-pass filters to 

discriminate between the excitation (blue) and emission (orange) signals, capture the intensity image of 

the PSP-coated model surface, providing a means to recover global surface pressure distributions on test 

articles of interest. PSP measurement systems all employ a ratio of image pairs to compensate for 

intensity non-uniformity due to sources other than oxygen concentration, the most significant of which 

are paint application and illumination heterogeneity. In the conventional approach, PSP images acquired 

either prior to or immediately following tunnel operation (wind-off) are ratioed with images acquired at 

each tunnel condition (wind-on). A companion technique, temperature sensitive paint (TSP), uses probe 

molecules whose brightness varies with temperature, but which are insensitive to oxygen concentration.

If the test surface under study is immersed in an atmosphere containing O2 (e.g. air), the recovered 

luminescence intensity can be described by the Stern-Volmer relationship12

I0/I = 1 + KSV(T)PO2 (1)
4

where I0 is the luminescence intensity in the absence of O2 (i.e. vacuum), I is the luminescnec intensity at 

some partial pressure of oxygen PO2, and KSV is the Stern-Volmer rate constant. The value of KSV depends 

on the properties of both the luminescent molecules and the binder, and is generally temperature 

dependent. Since it is a practical impossibility to measure I0 in a wind tunnel application, a modified form 

of the Stern-Volmer equation is typically used. This form replaces the vacuum calibration (I0) with a 

reference standard

IREF/I = A(T) + B(T)P/PREF (2)

where IREF is the recovered luminescence intensity at a reference pressure, PREF. A(T) and B(T) are 

temperature dependent constants for a given PSP formulaiton and are ususally determined a priori using 

laboratory calibration procedures. The calibration and interpretation of TSP data follows a similar 

procedure, using the TSP luminescence intensity recovered at a reference temperature.

PSP measurements are difficult to make under cryogenic conditions for two reasons. First, the test gas 

is typically nitrogen, refrigerant, or some other medium that usually contains little or no oxygen. Second, 

the diffusion of oxygen into the paint binder is highly temperature dependent, and at low temperatures, is 

practically nonexistent. As such, it is not surprising that initial cryogenic testing with luminescent paints 

used TSP.13 Indeed; probe molecules can be found which respond strongly to temperature even under 

cryogenic conditions.

For PSP, researchers developed two distinct methods to overcome the challenges caused by the low 

diffusion of oxygen into the paint binder. Initial cryogenic PSP tests were conducted using anodized 

aluminum.14 In these tests, the model was constructed from aluminum and anodized to coat the surface 

with a monolayer of aluminum oxide. Then the model was coated with an oxygen luminophore that 

chemically bonds with the aluminum oxide. This produces a coating that is essentially a few molecules 

thick, thus alleviating the issues with paint binders and oxygen diffusion. The disadvantages to this

technique are that the anodizing requirement makes it difficult to use with large and complex models, and 

it cannot be used with the stainless steel models that are typically employed in large-scale cryogenic 

testing. For this type of testing, a PSP formulation consisting of a binder that has a very large diffusion 

rate and can be applied with standard airbrush techniques and capable of adhering to many different types 

of surfaces was developed.15,16 These formulations have been used successfully at the NASA Langley 0.3-

Meter Transonic Cryogenic Tunnel15 as well as other facilities.16 A variation of the PSP technique has 

been performed at the NTF with this formulation to visualize surface flow patterns on a blended wing 

body model,17and preliminary testing of cryogenic PSP at the European Transonic Windtunnel has also 

been reported.18 Recently, a full cryogenic PSP test has also been reported at the NTF on a traditional 

commercial transport wing.19

The cryogenic PSP test on the traditional commercial transport wing at the NTF highlighted some 

potential challenges that needed to be overcome. These were confirmed in a follow on test that was 

performed on the Common Research Model.

20 In this test, severe contamination of the PSP was seen in 

the PSP that mitigated much if not all of its pressure sensitivity at cryogenic conditions. This 

contamination occurred if the tunnel required a significant amount of conditioning before cryogenic 

operations. To overcome this contamination, a two-pronged approach was developed and applied on this 

test. First, the base coat was cleaned thoroughly using toluene as a solvent to remove as much oil as 

possible before over-spraying with the PSP layer. Second, the PSP layer was applied in an access insert in 

the tunnel test section after the majority of the tunnel was conditioned for cryogenic operation. A full 

detailed procedure for the application and removal of the PSP is provided in the attached appendix.

This report will detail the results of these procedures as they were applied to a circulation control 

wing concept through a variety of conditions. Overall, the PSP performed much better than previous tests, 
5

though some issues with noise levels still exist, depending on oxygen content in the flow. Additional 

lessons were also learned, and they will be addressed as well. This test served as a risk reduction 

opportunity for follow-on tests at this facility.

II. Experimental

A. Paint Formulation and Calibration

The PSP formulation used for this 

study was made by dissolving an oxygen 

sensitive luminophore (platinum mesotetra(pentafluorophenyl) porphine, or 

Pt(TfPP) in plytrimethylsilylpropine, 

PTMSP.15,16,21 PTMSP was chosen as the 

binder because it is a glassy polymer with 

a large free volume, enabling it to have a 

high oxygen diffusion rate. This leads to 

high oxygen permeability even at 

cryogenic conditions.

All paint calibrations were performed 

separate from the wind tunnel in a 

laboratory calibration chamber that has 

been described previously.20 A typical 

PSP response at cryogenic conditions to 

increasing pressure of a 3000 ppm 

oxygen-nitrogen mix is shown in Figure1.

Oxygen-nitrogen mixtures are typically 

used to simulate expected NTF conditions with respect to O2 levels and were chosen based on previously 

experienced results.19,20

B. Facility

The NTF is one of the world’s leading facilities for providing high quality flight Reynolds number 

aeronautical data and has been operational since 1982. The tunnel has a test section of 2.5 m x 2.5 m (8.2 

ft x 8.2 ft) and is capable of operating at speeds from subsonic (M = 0.1) to transonic (M = 1.1) with 

Reynolds numbers from 13.1 x 106/m (4 x 106/ft) to 476 x 106/m (145 x106/ft) at transonic Mach numbers. 

The tunnel is capable of operating in either air mode at elevated pressure and temperature or at cryogenic 

conditions (to -156 oC or -249 oF) by injecting liquid nitrogen into the tunnel circuit. In the cryogenic 

operating mode, the NTF is capable of providing full-scale flight Reynolds numbers without an increase 

in model size. Several optical diagnostic techniques are available in the NTF, including Video Fluorescent 

Minituft Flow Visualization, Sharp Focusing Schlieren Flow Visualization, Video Model Deformation, 

and a newly designed and installed PSP/TSP system.

As mentioned in the introduction above, PSP requires the presence of O2 in the flow to operate 

correctly. In previous tests,19,20 this was accomplished by injecting dry air into the tunnel using a valve to 

control the injection. However, in this case, the air injection was only available from the model itself as 

part of the circulation control configuration. Oxygen monitors (described below) were interfaced into the 

tunnel and monitored the amount of O2 at any given time in the flow.

C. Model

The test article for this experiment was a semi-span representation of a generic transonic transport 

named the Fundamental Aerodynamic Subsonic Transonic Modular Active Control (FASTMAC) model. 

Figure 1. Typical calibration curve s for PSP under 

cryogenic conditions using a 3000 ppm O2 in N2 to simulate 

typical NTF conditions.
6

It was designed by Christopher Cagle at NASA LaRC, and was fabricated by the NASA LaRC machine 

shop as well as Triumph Aerosopace Systems in Newport News, VA. This model was specifically 

designed for use with the flow control system, and its modular design should provide for easier 

modifications or augmentation of its assembly. With this flexibility, the model is perfectly suited to act as 

a testbed for future circulation control studies at the NTF. As with any model at the NTF, all model 

hardware were designed and fabricated according to the model design criteria dictated in Langley 

Procedural Requirement (LPR) 1710.15. A planform sketch of the model as assembled is shown in Figure 

2 and a side view is shown in Figure 3 with pertinent dimensions in inches (cruise configuration shown).

The model fuselage and non-metric stand-off are composed of 2024-T351 aluminum allow. The 

wing-balance interface block and the wing assembly, including the flaps, are composed of 13-8Mo H1050 

stainless steel. All of the pressure tubing inside the model that directs the high-pressure air from the flow 

control system out of the trailing edge of the wind is composed of 304 stainless steel. It must be noted 

here that the model was fabricated with 1.9 cm (0.75 in) diameter 347 stainless steel tubing, but during 

assembly it was deemed that the 0.12 cm (0.049 in)wall thickness of this tubing was too thin, and 394

stainless steel tubing that had a wall thickness of 0.165 cm (0.065 in) was used inside the model. Fully 

assembled the model is 2.08 m (82 in) long (fuselage nose to tail) with a wingspan of about 1.29 m (50.72 

in) measured from test section wall to wingtip.

The model was designed and fabricated such that it could be configured for low-speed high-lift or for 

high-speed cruise. Three different flaps were fabricated including a 60° flap, a 30° flap, and a 0° or cruise 

flap. A leading edge slat was also fabricated that attaches to the leading edge with five slat brackets. 

Many sets of small shims were fabricated that can be used to change the trailing edge gap height. The gap 

height can be configured for a constant gap height across the span, or for a variable gap height across the 

span that is proportional to the chord length.

Figure 2. Planform view of the FASTMAC model.
7

The fuselage assembly consists of a non-metric stand-off and the metric fuselage sections. The nonmetric stand-off is approximately two inches wide and is bolted and pinned to the turntable. A springloaded Teflon seal is installed in a recessed perimeter around the back side of the standoff, and this serves 

as a flow blocker between the stand-off and test section wall as the model pitches. The metric upper and 

lower fuselage sections are bolted and pinned to the wing-balance interface block, and the aft fuselage 

section is bolted to the upper and lower fuselage sections after they are installed. A labyrinth seal exists 

between the fuselage sections and the stand-off to act as a flow blocker.

High-pressure air is directed through the model through a series of pipes, valves, and flow splitters. 

The wing-balance interface block attaches directly to the top hat and to the model interface plate on the 

bellows. A cap is installed on the wing-balance interface to prevent flow from entering the model from 

the low-flow side of the flow control system. Air from the high flow side of the flow control system is 

directed through the wing-balance interface block to fittings on the side of the block (two per side). Four 

1.9 cm (0.75 in) diameter tubes are then used to route the air to four independently controlled valves 

inside the fuselage. Another set of piping is then used to route the air from the four valves through four 

separate flow splitters, or manifolds, into four different plenums along the trailing edge of the wing spar. 

The flow then passes through four separate choke plates before entering the aft plenum of the wing and 

trailing edge flap. Air is ejected out of the wing through a gap between the flap and the upper flap cover.

The fuselage is instrumented with 42 static pressure orifices. It also has threaded holes for installing 

five pressure taps. The main wing spar and leading edge is instrumented with 111 static pressure orifices. 

Small blocks are also installed in the wing for the installation of eight pressure taps to measure the 

pressure inside the four plenums in the aft wing spar. The leading edge slat is instrumented with 32 static 

pressure orifices. Each of the three trailing edge flaps are also instrumented with 60 static pressure 

orifices. Four Endevco pressure taps are also permanently installed in each flap to measure the unsteady 

pressures near the trailing edge gap. With the leading edge installed, the model contains 245 static 

pressure orifices. All pressure orifices in the wing are arranged in four streamwise rows at four span 

locations on the upper and lower surfaces of the leading edge slat, wing, and flaps. It must be noted that 

the pressure rows on the flaps are actually arranged perpendicular to the trailing edge and are not arranged 

in the streamwise direction.

D. Instrumentation

Illumination

Illumination was achieved using several custom designed LED-based arrays capable of operating at 

either room temperature (for NTF air model operation) or at cryogenic conditions. The arrays consisted 

of 80 individual LED elements arranged on a 12.7 cm diameter 4.5 mm thick aluminum substrate. The 

aluminum substrate is also equipped with an RTD sensor to monitor temperature as well as resistive 

heaters on the back for cryogenic operation. The LED arrays are operated using a 1kW switching power 

supply and can be remotely operated using standard TTL (transistor-transistor logic) pulses. The LED 

array produces light centered at 400 nm (~20 nm full width at half maximum) at greater than 80 W under 

Figure 3. Side view of the FASTMAC model.
8

nominal forward current conditions. This can also be manually adjusted as needed by controlling the 

current applied to each array. 

Image Acquisition

Images were acquired from cameras placed directly over the wing of the model. The cameras 

employed were thermoelectrically-cooled monochrome interline transfer CCD cameras employing 11 

Mpixels (4000 x 2672) capable of acquiring images at up to 5 fps at full resolution. The cameras employ 

an electronic shutter alleviating various frame transfer issues that were seen previously when cameras 

using mechanical shutters were used. This allows for operation where the LED illuminators are on 

constantly, alleviating some electrical and timing issues seen previously. The cameras employ 12-bit 

digital resolution and were interfaced to the computer using custom designed armored fiber optic cables 

capable of operating at cryogenic temperatures.

Oxygen Monitoring

To facilitate calibration of the paint, the oxygen concentration in the flow must be measured 

accurately. This was done by interfacing an oxygen monitoring system into the tunnel control loop. Two 

oxygen monitors were attached to outlets of pressure ports, one from a static port in the floor and one 

from a dynamic port in a small pressure rake. Both were mounted approximately 4.5 m forward of the 

model. The oxygen sensor system employed a zirconium oxide sensor with a time response of less than 

10 s. The system was also equipped with an in situ calibration option allowing it to maintain linearity and 

repeatability of less than 2% of reading or 0.5 ppm O2 absolute. The unit is controlled by a personal 

computer via RS-232 protocol.

E. Data Acquisition

Data acquisition procedures for this test were modified from the previous cryogenic tests at NTF.20 In 

this case; no additional air was injected into the tunnel. All the oxygen was introduced into the tunnel 

during the normal operation of the circulation control (i.e. all the air was introduced from the blowing 

flap). As in the previous tests, all wind-off images were acquired using a “slow roll” with the tunnel 

operating at ~ Mach 0.09 (to ensure adequate circulation of the oxygen). The oxygen concentration was 

continuously monitored as described above. However, no attempts were made to stabilize the amount of 

oxygen present. As such, the oxygen concentration varied wildly throughout the tests from only a few 

hundred ppm to near 1%, depending on the blowing conditions at the time of image acquisition. Because 

of this, only an in situ calibration of the PSP using existing taps was possible.

F. Data Analysis

Data analysis for this work used typical procedures for the analysis of intensity-based data 

acquisition. The basic data analysis used the following protocol:

1. Background correction of all images

2. Registration of the wind-on images to a suitable slow roll image. The slow roll image chosen for a 

particular wind-on image matched the angle of attack and was as close in time to the wind-on image as 

possible.

3. Conversion of the images to pressure was accomplished by using an in situ technique9in which the 

pressure taps were used to calibrate the PSP. This was done due to the potentially significant oxygen 

differences between the flap and the upper surface.

III. Results and Discussion

A. General Considerations

The in situ method for the calibration of the paint was accomplished by comparing a subset of 

pressure taps on the upper surface of the wing with a corresponding PSP area. The general locations of 
9

the chord-wise rows used for the paint calibration are shown in Figure 4. The row numbering was based 

on the location titles from the file provided in the tunnel data acquisition system. From the figure, it is 

readily apparent that Row 1 could not be used as it is in a region of the wing that could not be imaged 

adequately by the camera. This is most likely due to vignetting that has been seen previously.20 

Additionally, from Figure 4, there is definite evidence that the flap can show large differences than the 

upper surface of the wing. In the case of Figure4, it is due to the blowing from the slot of the wing. This 

causes a significant increase in the local concentration of O2 this is directly behind the slot that is blowing 

air. Meanwhile the upper surface of the wing is essentially in a region of the facility that has a very low 

O2 concentration. The portions of the flap that seem to have no data have actually had the PSP 

completely quenched so that the camera can pick up no data from these areas. Thus, for this test, all data 

analysis is concentrated on the upper wing surface. For all of the data runs, PSP was acquired at angles of 

attack of 0o, 4o, and 8o.

Figure 4. Sample PSP image depicting the general locations of the chord-wise tap rows.

B. Mach 0.7 (Note: Figure 5-10 are placed at the end of the paper for clarity)

Figures 5-7 show the PSP results obtained at Mach 0.7 with no air blowing form the slots. As there 

was no external air introduced into the N2 atmosphere of the tunnel, the native O2 concentration was very 

low (typically less than 1000 ppm). This has resulted in very noisy data on the upper wing. In addition, 

the flap shows some strange inconsistencies that arise because some of the valves in the circulation 

control segment were stuck open in the slow-roll configuration, thus biasing the data. These are manifest 

as extremely low pressure regions on the flap. While the data on the upper wind surface is noisy and 

shows many features in the paint that arise from the application (due to very small intensity changes), the 

actual agreement with the taps is fairly good. These are readily apparent in Figure 5 (0o AOA), where 

there is little change chord-wise across the wing, and in Figure 7 (8o AOA), in which both the PSP and 

taps show lower pressure at the leading edge increasing toward the trailing edge. However, the PSP data 

for Figure 6 (4o AOA) shows significantly more noise in both the image and the tap agreement. It is 

unknown why this occurred, though it could be due to significant depletion of the O2 concentration in the 

facility at this point.
10

The corresponding cases where air was actually being blown from the slot are shown in Figures 8-10. 

For this case, the slow-roll condition was taken after the wind-on acquisition, and the issue with the 

valves had been resolved, alleviating the issue with the flap. Unfortunately, the noise in the PSP data for 

Figure 8 (0o AOA) is quite excessive. However, the other two conditions, shown in Figures 9 and 10 (4o

and 8o AOA, respectively), show much cleaner PSP data which agrees with the taps quite well. From the 

two sets of data collected at Mach 0.7, the 8o AOA case was the only one that can be directly compared as 

the PSP was the cleanest in each instance. This comparison, along with the tap comparison from Row 3 

is shown in Figure 11. The comparison shows that there is a slight effect from the blowing over the flap. 

However, there is another anomaly that seems to occur near the trailing edge in the PSP. As seen in the 

tap comparison at Row 3, the PSP begins to deviate at x/c ~ 0.8. This can be seen in the PSP image as 

higher pressure region that seems to be extending from the flap towards the leading edge. This will be 

explored in more detail in the next conditions.

Figure 11. Comparison of PSP data acquired at Mach 0.7, 8o AOA with and without blowing through 

CCW. The PSP comparison to the taps in both cases is taken at Row 3 from Figure 4.

C. Mach 0.8 (Note: Figure 12-17 are placed at the end of the paper for clarity)

The next set of runs was acquired at Mach 0.8. Figures 12 -14 show the PSP results for the case 

where air is not being blown through the slot. This used the same slow-roll condition as the Mach 0.7 

case, so the flap anomalies are present. Again, as in the Mach 0.7 condition, the PSP is very noisy due to 

low O2 concentration. In this case, the 0o AOA (Figure 12) case was run first and shows fairly good 

agreement with the taps. For the 4o AOA condition (Figure 13), the PSP data was much cleaner than the 

Mach 0.7 condition, and shows the existence of a separation bubble near the leading edge. The boundary 

of this bubble can be seen in the tap measurements as well as the PSP, which tracks very well. For the 8o
11

AOA (Figure 14) condition, the PSP is again very noisy. However, this was taken at the end of the data 

run, so the O2 concentrations were at their lowest. Still, the PSP does seem to be following the trend from 

the taps.

The PSP results are cleanest at Mach 0.8 when the slots are set for blowing air. This should be 

expected as this is the condition in which both the O2 concentration in the tunnel and the pressure changes 

over the wing should be the greatest. These results are detailed in Figures 15-17. At the 0o AOA (Figure 

15) condition, the wing shows relatively little pressure change across the surface, as seen in both the tap

readings as well as the PSP. The good agreement between the PSP and the pressure taps is also seen in 

the 4o(Figure 16) case. In this case, the separation bubble is present and has shifted rearward a large 

amount when compared to the 4ocase at Mach 0.7 (as seen in Figure 9). At Mach 0.7, the bubble if fairly 

small and reattaches at x/c ~ 0.15. However, at Mach 0.8, this reattachment occurs at x/c ~ 0.45. While 

the reattachment location is evident in both the tap data as well as the PSP, the PSP provides a means to 

view the entire separation bubble to ascertain shape over the wing.

For the 8ocase (Figure 17) at Mach 0.8 with blowing, there is a large deviation of the PSP from the 

taps at the trailing edge of the upper wing surface. Evidence for this could be seen beginning at the 8o

condition at Mach 0.7 (Figures 10 and 11), but is readily seen at Mach 0.8, Because the phenomenon is 

only seen in the PSP, it can be assumed that this is a significantly large temperature gradient from inside 

the model caused by the air being introduced through the bellows. Unfortunately, only the three angles of 

attack were investigated, though even from this limited data, it seems to follow that this heating event 

seems to only occur at higher angles. Ideally, with adequate run time, a full angle of attack sweep would 

be needed to confirm this, as well as identify when this begins to occur as it could indicate potential issues 

with the circulation control system at the facility.

IV. Conclusions

This report presented results from a PSP test performed at cryogenic conditions at the National 

Transonic Facility on a circulation control wing concept. This test built upon the lessons learned from a 

PSP test described previously,20 especially in regards to the paint application. The proper application of 

PSP is the most critical stage to ensure adequate performance in cryogenic testing, especially in an 

environment with greatly reduced O2 levels. The main result of this test was the development of these 

procedures that are presented in the appendix of this report. Following these procedures resulted in a 

viable coating that required no repainting and had much less contamination than the previous test.

As for the PSP results, data acquired while the circulation control was off (i.e. no blowing) produced 

noisy data which can be traced back to the fact that the only O2 introduced into the facility was by the 

circulation control itself. Thus, in these conditions, the O2 levels were very small. However, even in 

these cases, the PSP generally agreed with the tap readings, and certain conditions did provide some 

insight into the flow over the upper wing surface. However, when the circulation control was active (i.e. 

blowing), the O2 levels generally rose to acceptable levels, resulting in very clean PSP data that generally 

agreed well with the taps. The PSP was able to ascertain separation and reattachment regions (especially 

at Mach 0.8), but did show a probable temperature anomaly at the leading edge of the upper wing surface 

at the highest angle of attack (8o). This resulted in a significant deviation between the taps and PSP.

References

1. Kavandi, J., Callis, J., Gouterman, M., Khalil, G., Wright, D., Green, E., Burns, D., and McLachlan, 

B., “Luminescence Barometry in Wind Tunnels,” Review of Scientific Instruments, Vol. 61, No. 11, 

1990, pp. 3340-3347.

2. Morris, M.J., Benne, M.E., Crites, R.C., and Donovan, J.F., “Aerodynamic Measurements Based in 

Photoluminescence,” AIAA Paper 93-0175, 31st Aerospace Sciences Meeting and Exhibit, Reno, 

Nevada, January 11-14, 1993.
12

3. McLachlan, B. and Bell, J., “Pressure-Sensitive Paint in Aerodynamic Testing,” Experimental 

Thermal and Fluid Science, Vol. 10, No. 4, pp. 470-485.

4. Torgerson, S., Liu, T., and Sullivan, J., “Use of Pressure Sensitive Paints in Low Speed Flows,” AIAA 

Paper 96-2184, Advanced Measurement and Ground Testing Technology Conference, New Orleans, 

Louisiana, June 17-20, 1996.

5. Coleman, P.B., Wallis, S.B., Dale, G.A., Jordan, J.D., Watkins, A.N., Goss, L., Davis, C.P.N., and 

Walter, T.M., “A Comparison of Pressure Sensitive Paint (PSP) Techniques for Aerodynamic Testing 

at Slow Velocities,” Society of Automotive Engineers Paper 2002-01-255, SAE 2002 World Congress 

and Exhibition, Detroit, Michigan, March 2002.

6. Jordan, J.D., Watkins, A.N., Davis, J., Weaver, W.L., Dale, G.A., Navarra, K., Urban, J., Devoid, W., 

and Stange, R., “Pressure-Sensitive Paint Measurements in a Large Scale Commercial Engine Test 

Stand,” Proceedings of the 18th International Congress on Instrumentation in Aerospace Simulation 

Facilities, Toulouse, France, June 14-17, 1999.

7. Carroll, B., Abbott, J. Lukas, E., and Morris, M., “Pressure Sensitive Paint Response to a Step 

Pressure Change,” AIAA Paper 95-0483, 33rd Aerospace Sciences Meeting and Exhibit, Reno, 

Nevada, January 9-12, 1995.

8. Winslow, N., Carroll, B., and Setzer, F., “Frequency Response of Pressure Sensitive Paints,” AIAA 

Paper 96-1967, 27th Fluid Dynamics Conference, New Orleans, Louisiana, June 17-20, 1996.

9. Liu, T., Campbell, B., Burns, S., and Sullivan, J., “Temperature- and Pressure-Sensitive Luminescent 

Paints in Aerodynamics,” Applied Mechanics Review, Vol. 50, No. 4, 1997, pp. 227-246.

10.Bell, J.H., Schairer, E.T., Hand, L.A., and Mehta, R.D., “Surface Pressure measurements Using 

Luminescent Coatings,” Annu. Rev. Fluid Mech., Vol. 33, 2001, pp. 155-206.

11.Liu, T., and Sullivan, J.P., “Pressure and Temperature Sensitive Paints (Experimental Fluid 

Dynamics),” Springer-Verlag, Berlin, 2004.

12.Lackowicz, J., Principles of Fluorescence Spectroscopy, Second Edition, Plenum Press, New York, 

1999, Chapter 8.

13.Asai, K. Kanda, H., and Kunimasu, T., “Detection of Boundary-Layer Transition in a Cryogenic Wind 

Tunnel by Using Luminescent Paint,” AIAA Paper 96-2185, Advanced Measurement and Ground 

Testing Technology Conference, New Orleans, Louisiana, June 17-20, 1996.

14.Asai, K., Kanda, H., Cunningham, C.T., Erausquin, R., and Sullivan, J.P., “Surface Pressure 

Measurement in a Cryogenic Wind Tunnel Using Luminescent Coating,” Proceedings of the 17th

International Congress on Instrumentation in Aerospace Simulation Facilities, Pacific Grove, 

California, September 29-October 2, 1997, pp. 105-114.

15.Upchurch, B.T. and Oglesby, D.M., “New PSP Developments at NASA Langley Research Center –

Low Temperature PSP,” Proceeding of the 6th Annual Pressure Sensitive Paint Workshop, Seattle, 

Washington, October 6-8, 1998, pp. 10-1 – 10-24.

16.Asai, K., Amao, Y., Iijima, Y., Okura, I., and Nishide, H., “Novel Pressure-Sensitive Paint for 

Cryogenic and Unsteady Wind Tunnel Testing,” AIAA Paper 2000-2527, 21st AIAA Aerodynamic 

Measurement Technology and Ground Testing Conference, Denver, Colorado, June 19-22, 2000.

17.Watkins, A.N., Goad, W.K., Obara, C.J., Sprinkle, D.R., Campbell, R.L., Carter, M.B., Pendergraft, 

O.C., Bell, J.H., Ingram, J.L., Oglesby, D.M., Underwood, P.J., and Humber, L.R., “Flow 

Visualization at Cryogenic Conditions Using a Modified Pressure Sensitive Paint Approach,” AIAA 

Paper 2005-0456, 43rd Aerospace Sciences Meeting and Exhibit, Reno, Nevada, January 10-13, 2005.

18.Germain, E. and Quest, J., “The Development and Application of Optical Measurement Techniques 

for High Reynolds Number testing in Cryogenic Environment,” AIAA Paper 2005-0458, 43rd

Aerospace Sciences Meeting and Exhibit, Reno, Nevada, January 10-13, 2005.
13

19.Watkins, A. N., Leighty, B.D., Lipford, W.E., Oglesby, D.M., Goodman, K.Z., Goad, W.K., Goad, 

L.R., and Massey, E.A., “The Development and Implementaiton of a Cryogenic Pressure Sensitive 

Paint System in the National Transonic Facility,” AIAA Paper 2009-0421, 47th AIAA Aerospace 

Sciences Meeting including The New Gorizons Forum and Aerospace Exposition, Orlando, Florida, 

January 5-8, 2009.

20. Watkins, A.N., Lipford, W.E., Leighty, B.D., Goodman, K.Z., Goad, W.K., and Goad. L.R., “Results 

From a Pressure Sensitive Paint Test Conducted at the National Transonic Facility on Test 197: The 

Common Research Model,” NASA/TM-2011-217065, 2011, 18 pp.

21.Masuda, T., Isobe, E., and Higashimura, T., “Poly[1-(trimethylsilyl)-1-propyne]: A New Polymer 

Synthesized With Transition-Metal Catalysts and Characterized By Extremely High Gas 

Permeability,” Journal of the American Chemical Society, Vol. 105, No. 25, 1983, pp. 7473-7474.
14

Appendix: Procedure for Applying and Removing Pressure Sensitive Paint for 

Cryogenic Testing at the National Transonic Facility

This appendix will serve as a Standard Operating Procedure for applying and removing Pressure 

Sensitive Paint for cryogenic testing at the National Transonic Facility. This also serves as one of the 

main lessons learned during this testing and these practices were refined during this tunnel entry.

Preparation of Model

1. Mask off the model areas that will not be painted. Use 3M Scotch Brand 218 MQA striping tape or 

equivalent.

2. If taps are present, apply an air purge to the model taps to prevent particles or paint spray from 

entering the taps. Determine how many pressure tap openings will need to be purged and adjust the air 

flow such that no cratering occurs around the taps. Use your tongue to gauge the amount of pressure from 

each pressure tap opening. (For Example: If there are 10 pressure tap openings use approx. 7-12 psi of 

nitrogen or clean, dry air depending on the size of the pressure tap openings.)

3. If unable to purge pressure taps, use a hole punch, hammer and 3M Scotch Brand 218 MQA striping 

tape or equivalent to create a circular mask. MUST BE APPROVED PRIOR TO INSTALLATION 

AND/OR HAVE FACILITY PERSONNEL INSTALL CIRCULAR MASK! If circular masks are 

approved for installation, determine what the diameter of the pressure tap hole is and prepare a circular 

mask of the appropriate size to be placed over top of the pressure tap hole using fine tip tweezers. (For 

Example: If the pressure tap hole is 0.020” wide use a 1/8 inch circular mask.) 

NOTE: If used during painting, the circular mask may create a circular indention after painting is 

complete. Indentions on the surface may also cause varying results during data acquisition due to 

variations of the flow. 

4. Once pressure taps are protected, lightly spray solvent onto a paint towel and/or rag and gently wipe 

surface of the model to remove all traces of oil, grease, debris and surface impurities

5. Wipe the surface with a tack cloth just prior to spraying to remove any dust, dirt and/or debris.

Application of PSP

1. Paint the model with GBP 988 Self Etching Primer. The Self Etching Primer should air cure for 30 

minutes.

NOTE: Spray several light coats of self etching primer to make sure surface stays as smooth and 

clean of debris as possible when spraying inside the tunnel due to heat and humidity of tunnel conditions. 

When spraying in the test section and/or wind tunnel, make sure a fan and/or air filtering machine 

(portable paint booth) is used at all times to remove all paint and solvent fumes.

2. Then wet sand the model with 1500 grit sand paper to a roughness of 5 μin (Ra), as measured with the 

Mitutoyo Surftest 211.

3. Spray the surface with Spectra Prime. Cure the primer for 2 hours at 65o C. The Spectra Prime must 

be cured until it can be rubbed with a cloth soaked in THF and will not become gummy.
15

NOTE: Add 1 oz more activator/reducer to allow for better flow of Spectra Prime base coat when 

spraying inside the tunnel due to heat and humidity of tunnel conditions. Spray several heavy coats to 

ensure the primer stays wet during the application which will provide a smoother surface for wet sanding. 

When spraying in the test section and/or wind tunnel, make sure a fan and/or air filtering machine 

(portable paint booth) is used at all times to remove all paint and solvent fumes.

4. Wet sand the cured primer with 2000 grit wet sand paper until a roughness below 5μin (Ra) (as 

measured with the Mitutoyo Surftest 211 roughness meter) and the allowed thickness is achieved. 

5. Rub the primed, sanded surface with a soft cloth or tissue soaked with tetrahydrofuran (THF).

6. The entire surface should have a dull finish. Any glossy spots should be rubbed until dull.

7. Spray the binder/toluene solution on the THF treated Spectra Prime primer in 15 - 20 light coats, 

allowing each coat to dry (this should take only a second or two; otherwise the coats are too heavy).

NOTE: Use a cross pattern when spraying the binder on the model and a sweeping motion for each 

pass. The binder must be sprayed approx. 10-12 inches away in the test section and 6-10 inches away in 

the tunnel due to the heat and humidity of the tunnel conditions. When spraying in the test section and/or 

wind tunnel, make sure a fan and/or air filtering machine (portable paint booth) is used at all times to 

remove all paint and solvent fumes.

8. When the model has been adequately coated the surface should have a uniform sheen.

9. After coating the entire model surface with the low temperature binder, allow it to dry for 10 minutes 

at room temperature.

10. Clean the spray gun with toluene.

11. Spray the polymer coated surface with a 600 ppm solution of PtT(PFP)P in methyethylketone (MEK) 

(also called 2-butanone) to give a uniform medium shade of pink. Allow to dry for 10 minutes at room 

temperature.

NOTE: Use a sweeping motion when spraying the luminophore on the model and do not repeat the 

same pass until the previous coat has cured for 1 minute. The luminophore must also be sprayed approx. 

10-12 inches away in the test section and 6-10 inches away in the tunnel due to the heat and humidity of 

the tunnel conditions. When spraying in the test section and/or wind tunnel, make sure a fan and/or air 

filtering machine (portable paint booth) is used at all times to remove all paint and solvent fumes.

12. Clean the spray gun with toluene. Check the roughness. It should be less than 10 μin (Ra).

Removal of PSP

If the PSP has not cured for a long period of time (more than 12 hours) it may be removed with 

acetone; however, it will also remove some and/or most of the primer. If acetone will not remove the 

PSP, Tal Strip 2 aircraft coating remover can be used. The area must be well ventilated and the paint 

remover must be sprayed 10-12 inches away from the model in light fluid movements until the entire 

model has been covered. Wait 10-20 minutes and then the model must be checked to see if most or all of 

the paint has begun to bubble and/or strip away from the model. Once most of the paint has bubbled 

and/or stripped away, the model should be thoroughly sprayed with deionized water while being wiped 
16

with a KimwipeTM. The model can then be thoroughly cleaned with acetone until all the paint has been 

removed and the model is clean of grease, dust, debris, paint, etc.
17

Figure 5. PSP image and comparison of PSP data with taps for Mach 0.7, 0o AOA, and no blowing through 

CCW slot.
18

Figure 6. PSP image and comparison of PSP data with taps for Mach 0.7, 4o AOA, and no blowing through 

CCW slot.
19

Figure 7. PSP image and comparison of PSP data with taps for Mach 0.7, 8o AOA, and no blowing through 

CCW slot.
20

Figure 8. PSP image and comparison of PSP data with taps for Mach 0.7, 0o AOA, and air blowing through 

CCW slot.
21

Figure 9. PSP image and comparison of PSP data with taps for Mach 0.7, 4

o AOA, and air blowing through 

CCW slot.
22

Figure 10. PSP image and comparison of PSP data with taps for Mach 0.7, 8

o AOA, and air blowing through 

CCW slot.
23

Figure 12. PSP image and comparison of PSP data with taps for Mach 0.8, 0o AOA, and no blowing through 

CCW slot.
24

Figure 13. PSP image and comparison of PSP data with taps for Mach 0.8, 4o AOA, and no blowing through 

CCW slot.
25

Figure 14. PSP image and comparison of PSP data with taps for Mach 0.8, 8o AOA, and no blowing through 

CCW slot.
26

Figure 15. PSP image and comparison of PSP data with taps for Mach 0.8, 0o AOA, and air blowing through 

CCW slot.
27

Figure 16. PSP image and comparison of PSP data with taps for Mach 0.8, 4o AOA, and air blowing through 

CCW slot.
28

Figure 17. PSP image and comparison of PSP data with taps for Mach 0.8, 8o AOA, and air blowing through 

CCW slot.
REPORT DOCUMENTATION PAGE Form Approved

OMB No. 0704-0188

2. REPORT TYPE 

Technical Memorandum

 4. TITLE AND SUBTITLE

Measuring Global Surface Pressures on a Circulation Control Concept 

Using Pressure Sensitive Paint

5a. CONTRACT NUMBER

 6. AUTHOR(S)

Watkins, Anthony Neal; Lipford, William E.; Leighty, Bradley D.; 

Goodman, Kyle Z.; Goad, William K.

 7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

NASA Langley Research Center

Hampton, VA 23681-2199

 9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001

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 REPORT NUMBER

L-20153

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NASA

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Unclassified - Unlimited

Subject Category 35 

Availability: NASA CASI (443) 757-5802

19a. NAME OF RESPONSIBLE PERSON

STI Help Desk (email: help@sti.nasa.gov)

14. ABSTRACT

This report presents the results obtained from the Pressure Sensitive Paint (PSP) technique on a circulation control concept 

model. This test was conducted at the National Transonic Facility (NTF) at the NASA Langley Research Center. PSP was 

collected on the upper wing surface while the facility was operating in cryogenic mode at 227 K (-50 oF). The test envelope 

for the PSP portion included Mach numbers from 0.7 to 0.8 with angle of attack varying between 0 and 8 degrees and a total 

pressure of approximately 168 kPa (24.4 psi), resulting in a chord Reynolds number of approximately 15 million. While the 

PSP results did exhibit high levels of noise in certain conditions (where the oxygen content of the flow was very small), some 

conditions provided good correlation between the PSP and pressure taps, showing the ability of the PSP technique. This work 

also serves as a risk reduction opportunity for future testing in cryogenic conditions at the NTF.

15. SUBJECT TERMS

Cryogenic Testing; Global Pressure Measurement; Non-intrusive Sensing; Optical Pressure Measurement; Pressure Sensitive 

Paint

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April 1, 2011 to April 29, 2011

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