Kazuyuki Nakakita

Image measurements of unsteady pressure fluctuation by a pressure-sensitive coating on porous anodized aluminium

Pressure-sensitive luminescent coating on porous anodized aluminium (AA-PSP) was applied to measure non-periodic unsteady pressure distribution on a wind-tunnel model. A high-speed digital video camera was used to capture the PSP signal. The pressure-sensitive dye was tris(4,7-diphenylphenanthroline) ruthenium(II) ([Ru(dpp)3]2+). The coating has a short response time of O(10 µs), although it exhibits temperature and humidity sensitivities. A hydrophobic coating was applied on the anodized aluminium surface to suppress the humidity sensitivity. A temperature sensitive paint was used to obtain the temperature distribution instantaneously with the pressure. The temperature data were used to correct the PSP response. An appropriate data acquisition procedure as well as digital image processing algorithm was established to compensate for the error from the temperature and humidity sensitivities. The present system was applied to measure the pressure distribution on a delta wing at a high angle of attack in transonic flow, whose flow is unsteady due to the interaction between shock waves and leading edge vortices. The non-periodic unsteady pressure distribution on the delta wing was successfully measured with the sampling rate of 1 kHz and within a few per cent error in absolute pressure level.

Recent topics in fast-responding pressure-sensitive paint technology at National Aerospace Laboratory

This paper reviews the recent advances in fast-responding Pressure-Sensitive Paint (PSP) technology at National Aerospace Laboratory in Japan. Since NAL started a research program on PSP in 1994, much effort has been focused on the development of PSP for unsteady pressure measurements. In recent study, we have made a substantial progress in the fast PSP formulations. This includes the development of a PSP using a high polymer called poly(TMSP) and chemisorptive PSPs based on anodized-aluminum. To investigate the response characteristics of these paints, we have made systematic tests using a pressure chamber with a fast-acting solenoid-type valve and a shock tube. A summary of these experimental investigations is given in this paper with an emphasis on the relationship between the response time and the physical properties of binder materials. In addition, the feasibility tests conducted in a short-duration hypersonic shock tunnel and a shock tube are present.

Pressure-Sensitive Paint Application to a Wing-Body Model in a Hypersonic Shock Tunnel

Pressure-sensitive paint (PSP) was applied to Mach 10 hypersonic wind tunnel tests. This experiment was conducted at a short-duration Shock Tunnel with the duration of 30ms. A fast response anodized aluminum pressure-sensitive paint (AA-PSP) was used as PSP. The response of AA -PSP is more than 10kHz. A chemical process to fabricate AA-PSP on models with complicated geometry was developed and applied to a 3-dimensional wing-body model. PSP measurement was conducted at two angles of attack, 0 degree and 20 deg ree, and images from various view angles and a close -up image were acquired for each case. A detailed pressure distribution pattern of complex flow phenomena including shock-wave/shock-wave interactions could be clearly measured with PSP. Effects of temperature rise during the test were also evaluated, and it was found that they were at negligible levels. PSP data was in good quantitative agreement with conventional pressure transducers even at the point where a strong interaction was observed.

Experimental Study of Slat Noise from 30P30N Three-Element High-Lift Airfoil in JAXA Hard-Wall Low-Speed Wind Tunnel

Aeroacoustic measurements associated with noise radiation from the leading edge slat of the canonical, unswept 30P30N three-element high-lift airfoil configuration have been obtained in a 2 m x 2 m hard-wall wind tunnel at the Japan Aerospace Exploration Agency (JAXA). Performed as part of a collaborative effort on airframe noise between JAXA and the National Aeronautics and Space Administration (NASA), the model geometry and majority of instrumentation details are identical to a NASA model with the exception of a larger span. For an angle of attack up to 10 degrees, the mean surface Cp distributions agree well with free-air computational fluid dynamics predictions corresponding to a corrected angle of attack. After employing suitable acoustic treatment for the brackets and end-wall effects, an approximately 2D noise source map is obtained from microphone array measurements, thus supporting the feasibility of generating a measurement database that can be used for comparison with free-air numerical simulations. Both surface pressure spectra obtained via KuliteTM transducers and the acoustic spectra derived from microphone array measurements display a mixture of a broad band component and narrow-band peaks (NBPs), both of which are most intense at the lower angles of attack and become progressively weaker as the angle of attack is increased. The NBPs exhibit a substantially higher spanwise coherence in comparison to the broadband portion of the spectrum and, hence, confirm the trends observed in previous numerical simulations. Somewhat surprisingly, measurements show that the presence of trip dots between the stagnation point and slat cusp enhances the NBP levels rather than mitigating them as found in a previous experiment.

Aerodynamic/Aeroacoustic testing in Anechoic Closed Test Sections of Low-speed Wind Tunnels

This paper describes new anechoic closed test sections in Japan Aerospace Exploration Agency (JAXA) and Virginia Tech (VT) not only for an aeroacoustic but also for an aerodynamic testing ability in Low-speed wind tunnels. The anechoic closed test section with Kevlar wall is an innovative concept and had been originally developed at VT. It was later applied to JAXA’s 2m x 2m wind tunnel. By using a high-lift device model, the JAXAs new test section was evaluated and validated acoustically by comparing to VT anechoic test results. Moreover, the aerodynamic characteristics in the new test section were also evaluated by comparing to results of the same model in JAXAs closed hard-wall test section. New wall interference correction procedure is proposed for the Kevlar wall test section, and it showed very good agreement with well-known corrected hard-wall results. This anechoic test section is useful and a promising tool for both aerodynamic and aeroacoustic testing.

Temperature Correction of Pressure-Sensitive Paint for Industrial Wind Tunnel Testing

Pressure-sensitive paint measurement can obtain a much more detailed surface pressure distribution than can be obtained using conventional pressure taps. However, the pressure-sensitive paint is sensitive not only to pressure but also to temperature, and where high accuracy is required, it is essential to compensate for this temperature dependency. This paper discusses data processing methods for pressure-sensitive paint measurement in transonic industrial wind tunnel testing, and proposes three methods to compensate for temperature dependency of the pressure-sensitive paint: an in situ method, an a priori method, and a hybrid of a priori and in situ methods. The pressure distributions from the pressure-sensitive paint data obtained by these proposed methods are compared with pressure tap data measured by conventional pressure transducers, and it is confirmed that the proposed methods are effective in compensating the temperature dependency of pressure-sensitive paint and improve the accuracy of the obtained data. It is also found that the hybrid of a priori and in situ methods is widely applicable to the industrial wind tunnel testing even if the pressure range of the pressure tap data is limited.

Effect of TSP Layer Thickness on Global Heat Transfer Measurement in Hypersonic Flow

made of glass ceramics. The thickness of TSP layer was varied from 0.2 to 3 µm. The models were tested at Mach 10 in the JAXA 0.44-m Hypersonic Shock Tunnel and heat flux caused by aerodynamic heating on the model surface was measured. A comparison between TSP data and conventional thermocouple data showed that the measurement error changed with the layer thickness in nonlinear fashion. In addition, the difference between TSP and thermocouple measurement was larger than theoretical prediction. For TSP layer thickness less than 0.5µm, the accuracy in heat flux measurement by TSP was within the uncertainty of conventional thin thermocouple measurement. To demonstrate the capability of TSP with the optimized thickness, tests using a 3-dimensional wing-body model have been conducted and the complicated heat flux pattern caused by shock-wave/shock-wave interaction has been observed.

Analysis of NASA Common Research Model Dynamic Data in JAXA Wind Tunnel Tests

The JAXA 2m x 2m Transonic Wind Tunnel (JTWT) conducted tests for 80% scaled NASA Common Research Model (CRM). The dynamic data including buffet measurement with strain gauges and dynamic pressure sensors were acquired at 50,000 Hz sampling rate. A prediction of buffet phenomenon is one of important factors to design aircraft. If buffet phenomenon occurs, dynamic bending and torsion moment are measured with wing-root strain gauges. Spectrum analyses are being executed for the strain gauges and dynamic pressure data. It is expected to observe dynamic flow separation at points around where the relation with lift coefficient data and pitching moment coefficient is non-linear. This paper describes overview of the tests and the analysis data.

80% Scaled NASA Common Research Model Wind Tunnel Test of JAXA at Relatively Low Reynolds Number

A wind tunnel test of a 80% scaled copy of the NASA Common Research Model (CRM) was performed in the 2m × 2m transonic wind tunnel of Japan Aerospace Exploration Agency (JAXA). The wind tunnel model was fabricated by JAXA consulting NASA Langley Research Center and the Drag Prediction Workshop committee members. The test was conducted at relatively low Reynolds number of 2.27 × 10 due to the limitation of the tunnel capability and boundary layer transition was simulated with optimized roughness. In the test campaign, static pressure distribution and aerodynamic forces were successfully acquired while the model main wings were deformed during the test due to the dynamic pressure. To make a fair comparison with the data from other sources in different circumstances, data normalization techniques were applied. Then, the data was compared with the data of the National Transonic Facility of NASA and CFD. The data normalization successfully realized fair comparisons for pressure distribution and lift coefficients while the tests were performed at the different circumstances such as the different Reynolds numbers.

Experimental Investigation of Vortex Generator Effect on Two- and Three-Dimensional NASA Common Research Models

Aerodynamic characteristics of twoand three-dimensional NASA common research model (2D-CRM and 3D-CRM) with co-rotating vortex generators (VGs) were investigated to clarify the influence of the three-dimensionality of the wings on the VGs effect. The base height of the VGs was 1.5 times of the boundary layer thickness at the VGs location. The direction of the VGs on the 3D-CRM was toe-out which meant the leading edge of the VGs turned to the wing tip. The Mach numbers in the 2Dand 3D-CRM experiment were 0.74 and 0.85 considering the sweepback angle of the 3D-CRM. The lift coefficient and the oil flow visualization showed that the effect of the VGs on the 3D-CRM was much larger than that on the 2D-CRM. From the comparison between the experiments and the CFD results, we concluded that the difference between 2Dand 3D-CRM was mainly caused by the crossflow due to the swept wing. The cross-flow enhances the effect of the co-rotating toe-out VGs on the swept wings. The installation drag of VGs was also investigated for the 3D-CRM and validated an empirical method to estimate the installation drag. At CL conditions below the design CL = 0.5, the VGs increased the total drag as expected, while at CL conditions above the design CL, the VGs decreased the total drag because the VGs suppressed the separation and the effect exceeded the installation drag of the VGs.