▪ Abstract An optical technique is described that is often used nowadays to measure surface pressures on wind tunnel models and flight vehicles. The technique uses luminescent coatings, which are painted on the model surface, excited by light of appropriate wavelength, and imaged with digital cameras. The intensity of the emitted light is inversely proportional to the surface pressure. Hence, the surface pressures can be measured efficiently and affordably with a high spatial resolution. The theory and chemistry of how such coatings work and the parameters that affect them are presented. The required hardware and software are described, with emphasis on the different measurement systems and procedures. The various error sources are discussed, and correction schemes that can be used to minimize them are presented. Sample results, covering a wide range of conditions and applications, are presented and discussed.

Surface Pressure Measurements Using Luminescent Coatings

The current paper describes the development of pressure-sensitive paint (PSP) technology as an advanced measurement technique for unsteady flow fields and short-duration wind tunnels. Newly developed paint formulations have step response times approaching 1 μs, making them suitable for a wide range of unsteady testing. Developments in binder technology are discussed, which have resulted in new binder formulations such as anodized aluminium, thin-layer chromatography plate, polymer/ceramic, and poly(TMSP) PSP. The current paper also details modeling work done to describe the gas diffusion properties within the paint binder and understand the limitations of the paint response characteristics. Various dynamic calibration techniques for PSP are discussed, along with summaries of typical response times. A review of unsteady and high-speed PSP applications is presented, including experiments with shock tubes, hypersonic tunnels, unsteady delta wing aerodynamics, fluidic oscillator flows, Hartmann tube o...

https://journals.sagepub.com/doi/pdf/10.1243/09544100JAERO243

A review of pressure-sensitive paint for high-speed and unsteady aerodynamics

Unstart phenomena induced by flow choking in scramjet inlet-isolators

This paper discusses the performance enhancement of supersonic air intake model through the implementation of blunted leading edge to the cowl lip section of the model. A supersonic air intake model with sharp cowl leading edge is initially considered to numerically investigate its performance. Mach 3, supersonic intake flow through the base model has been simulated using commercial CFD package Ansys Fluent-15. Comparison of numerical predictions and experimental measurements is presented to demonstrate the correctness and accuracy of numerical frame work followed in the present study. Higher order spatial accuracy of the solver along with suitably refined mesh helped in accurate capturing of the flow field. Modification to the cowl lip is proposed as an effective method to improve the performance of the supersonic air intake. Two different blunted cowl leading edge geometries were investigated to identify the possible enhancement in performance parameters. Improvement in mass capture and combustion stability attained through the use of forward shifted blunt cowl leading edge is presented. It is also revealed through the present study that the blunt cowl leading edge can reduce the intensity of shock wave boundary layer interaction occurring at the isolator entry section. Deviation in total pressure recovery and flow distortion observed with different supersonic air intake models are also discussed with reasons for the same. This study demonstrates the scope of overall improvement in scramjet engine performance through the use of suitably positioned blunt cowl leading edge.

https://jafmonline.net/JournalArchive/download?file_ID=44250&issue_ID=246

Alterations of Cowl Lip for the Improvement of Supersonic-Intake Performance

Thedevelopmentandapplicationofhigh-speedimagingandluminescentcoatingtechniquestomeasurefull-e eld surface heat-transfer rates in short-duration hypersonic e ow is presented. Tests were performed on an indented cone model at the 48-in. shock tunnel and the LENS I tunnel facilities at Calspan— University of Buffalo Research Center. Nominal test conditions ranged between Mach numbers of 9.5 and 11.1 and Reynolds numbers of 1:4 £ 10 5 and 3 £ 10 5 m ii 1 with run times of less than 10 ms. Processed submillisecond images show the threedimensional, time-dependent development of the embedded separated e ow and shock/boundary-layer interaction into a steady axisymmetric structure bounded by regions of laminar e ow. Conversion from processed image data to full-e eld heat-transfer measurements were performed using both an in situ calibration with thin-e lm platinum heat-transfer gauges as well as an a priori temperature calibration and transient heat-transfer theory. In situ calibrations displayed excellent correlation with surface-mounted gauges, whereas a priori calibrations showed a larger susceptibility to bias errors.

Heat-Transfer Measurements in Hypersonic Flow Using Luminescent Coating Techniques

The influence of small cylindrical bluntness of the leading edge of a flat plate on formation of spatial structures in a nominally two-dimensional supersonic compression corner flow at the Mach number M∞ ≈ 8 and a laminar state of the undisturbed boundary layer is studied by the method of temperature-sensitive paints. Streamwise vortices are found in the region of reattachment of the separated flow in a wide range of Reynolds numbers (0.15 · 106–2.55 · 106) for various angles of flow deflection and plate lengths. It is demonstrated that the existence of these vortices induces spanwise oscillations of the heat transfer coefficient; the amplitude of these oscillations may reach 30%. The maximum deviations of the Stanton number reaching 80% are observed in the case with significant roughness of the leading edge of the flat plate. Both the maximum Stanton numbers in the reattachment region and the amplitude of spanwise oscillations of the Stanton number induced by streamwise vortices are found to decrease significantly in the case of small bluntness of the leading edge. Solutions of three-dimensional Navier–Stokes equations are obtained for some test conditions. The computed results are in good agreement with experimental data, which points to a significant stabilizing effect of small bluntness on the intensity of streamwise vortices.

Effect of Small Bluntness on Formation of Görtler Vortices in a Supersonic Compression Corner Flow

Pressure sensitive paint technology is a new optical pressure measurement method providing new opportunities for non-contact pressure measurements. Specific features of PSP-technology for shock wind tunnel are discussed. An optical pressure measurement system for instantaneous measurements is described. Results of pressure measurements on a cylinder at Mach 6 are presented.

http://ieeexplore.ieee.org/iel3/3950/11428/00519465.pdf

Pressure sensitive paint application in shock wind tunnel

Application of Temperature Senstive Paint (TSP) technology for investigation of boundary layer transition in short-duration wind tunnels is presented. Investigations were made on sharp and blunted flat plates in UT-1 wind tunnel of TsAGI operated in Ludwieg scheme at Mach numbers 5, 6, and 8 (Reynolds numbers from 5.5 · 106 to 26 · 106 ). Both natural and induced boundary layer transitions were investigated.

/pdf/temperature-sensitive-paint-application-for-investigation-of-2biynioxht.pdf

Temperature sensitive paint application for investigation of boundary layer transition in short-duration wind tunnels

A luminescent pressure sensitive composition for determining an air pressure pattern comprises a binder, preferably a polymer binder, with a luminophore, of which the intensity of luminescence depends on air pressure. The luminophore of which the intensity of luminesence depends on air pressure (active luminophore) is mixed together with an additional luminophore of which the intensity of luminescence has different dependence on air pressure than said active luminophore (reference luminophore).

Vladimir Nikolaevich Radchenko

Papers

Effect of Small Bluntness on Formation of Görtler Vortices in a Supersonic Compression Corner Flow

Pressure sensitive paint application in shock wind tunnel

Temperature sensitive paint application for investigation of boundary layer transition in short-duration wind tunnels

Luminescent pressure sensitive composition

Effect of Görtler-like vortices of various intensity on heat transfer in supersonic compression corner flows