The development and capabilities of fast-responding pressure-sensitive paint (fast PSP) are reviewed within the context of recent applications to aerodynamic and acoustic investigations. PSP is an optical technique for determining surface pressure distributions by measuring changes in the intensity of emitted light, whereas fast PSP is an extension applicable to unsteady flows and acoustics. Most fast PSP formulations are based on the development of porous binders that allow for rapid oxygen diffusion and interaction with the chemical sensor. This article reviews the development of porous binders, the selection of luminophore molecules suitable for unsteady testing, dynamic calibrations of PSP, data-acquisition methods, and noteworthy applications for flow and acoustic diagnostics. Calibrations of the dynamic response of fast PSP show a flat frequency response to at least 6 kHz, with some paint formulations exceeding a response of 1 MHz. Various applications of fast PSP are discussed that highlight the capabilities of the technique, and concluding remarks highlight the need for the future development of fast PSP.

Fast Pressure-Sensitive Paint for Flow and Acoustic Diagnostics

Transverse jet in supersonic crossflows

Recent advances in the aerothermodynamics of spiked hypersonic vehicles

Review on active thermal protection and its heat transfer for airbreathing hypersonic vehicles

Numerical simulation of hypersonic flow over highly blunted cones with spike

Effectiveness of aerospikes/aerodisk assemblies as retractable drag-reduction devices for large-angle blunt cones flying at hypersonic Mach numbers is investigated experimentally in hypersonic shock tunnel HST2 using a 120-deg apex-angle blunt cone. An internally mounted accelerometer balance system has been used for measuring the aerodynamic drag on the blunt cone with and without forward-facing aerospikes at various angles of attack. The measurements indicate around 55% reduction in drag for the blunt cone with flat-disk spike at zero degree angle of attack for a freestream Mach number of 5.75. Surface convective heat-transfer rate measurements have been carried out on the blunt cone with a flat-disk tipped spike of varying length in order to locate the shock reattachment point on the blunt-cone surface. The measured heat-transfer rates fluctuate by about ±20% in the separated flow region as well as near the reattachment point indicating the unsteady flowfleld around the spiked blunt cone. The shock structure around the 120-deg apex-angle blunt cone with a 12-mm-long flat-tipped aerospike has also been visualized using the electric discharge technique. The visualized shock structure and the measured drag on the blunt cone with aerospikes agree well with the axisymmetric numerical simulations

Experimental Investigations of Hypersonic Flow over Highly Blunted Cones with Aerospikes

A blunt-nosed hypersonic missile mounted with a forward-facing cavity is a good alternative to reduce the stagnation heating rates. The effects of a forward-racing cavity on heat transfer and aerodynamic coefficients are addressed in this paper. Tests were carried out in
hypersonic shock tunnel HST2, at a hypersonic Mach number of 8 using a 41 deg apex-angle blunt cone. The aerodynamic forces on the test model with and without a forward-facing cavity at various angles of attack are measured by using an internally mountable accelerometer force balance system. Heat flux measurements have been carried out on the
test model with and without a forward-facing cavity of the entire surface at zero degree angle of attack with platinum sensors. A numerical simulation was also carried out using the computational fluid dynamics code (CFX-Ansys 5.7). An important result of this study is that the smaller cavity diameter has the highest lift-to-drag ratio, whereas the medium cavity has the highest heat flux reduction. Theshock structure around the test model has also been visualized using the Schlieren flow visualization technique. The visualized shock structure and the measured aerodynamic forces on the missile-shaped body with cavity configurations agree well with the axisymmetric numerical simulations.

Investigation of Missile-Shaped Body with Forward-Facing Cavity at Mach 8

A three-component accelerometer balance system is used to study the drag reduction effect of an aerodisc on large angle blunt cones flying at hypersonic Mach numbers. Measurements in a hypersonic shock tunnel at a freestream Mach number of 5.75 indicate more than 50% reduction in the drag coefficient for a 120degrees apex angle blunt cone with a forward facing aerospike having a flat faced aerodisc at moderate angles of attack. Enhancement of drag has been observed for higher angles of attack due to the impingement of the flow separation shock on the windward side of the cone. The flowfields around the large angle blunt cone with aerospike assembly flying at hypersonic Mach numbers are also simulated numerically using a commercial CFD code. The pressure and density levels on the model surface, which is under the aerodynamic shadow of the flat disc tipped spike, are found very low and a drag reduction of 64.34% has been deduced numerically.

Shock tunnel study of spiked aerodynamic bodies flying at hypersonic Mach numbers

Effectiveness of film cooling technique to reduce convective heating rates for a large angle blunt cone flying at hypersonic Mach number and its effect on the aerodynamic characteristics is investigated experimentally by measuring surface heat-transfer rates and aerodynamic drag coefficient simultaneously. The test model is a 60° apex-angle blunt cone with an internally mounted accelerometer balance system for measuring aerodynamic drag and an array of surface mounted platinum thin film gauges for measuring heat-transfer rates. The coolant gas (air, carbon dioxide, and/or helium) is injected into the hypersonic flow at the nose of the test model. The experiments are performed at a flow free stream Mach number of 5.75 and 0° angle of attack for stagnation enthalpies of 1.16 MJ/kg and 1.6 MJ/kg with and without gas injection. About 30%–45% overall reduction in heat-transfer rates is observed with helium as coolant gas except at stagnation regions. With all other coolants, the reduction in surface heat-transfer rate is between 10%–25%. The aerodynamic drag coefficient is found to increase by 12% with helium injection whereas with other gases this increase is about 27%.

Film cooling effectiveness on a large angle blunt cone flying at hypersonic speed

A new three-component accelerometer force balance has been designed, calibrated and tested in hypersonic shock tunnel (HST2) of Indian Institute of Science. The newly designed balance is able to measure aerodynamic forces (within test time of one millisecond) on test models at angles of attack from 0 to 12°. Two models, a blunt cone with after body and a blunt cone with after body and frustum are used to establish the accuracy of the force balance. The tests were conducted for the above two configurations with a constant Mach number of 8 and total enthalpy of 2.0 MJ/kg. The effectiveness of the balance is demonstrated by comparing the forces and moments of measured data with AGARD models. The flow fields around the test model are simulated using a 3D axisymmetric Navier–Stokes solver and the simulated results were compared with the measured values. Measured and computed force data are matched within ±10% for two different models tested here. The accuracy of the force balance is also estimated with the Newtonian theory and the values are approximately ±10% for the axial component and ±8% for the normal and pitching moment components.

https://link.springer.com/content/pdf/10.1007%2Fs00193-008-0172-8.pdf

S. Saravanan

Papers

Experimental Investigations of Hypersonic Flow over Highly Blunted Cones with Aerospikes

Investigation of Missile-Shaped Body with Forward-Facing Cavity at Mach 8

Shock tunnel study of spiked aerodynamic bodies flying at hypersonic Mach numbers

Film cooling effectiveness on a large angle blunt cone flying at hypersonic speed

Aerodynamic force measurement using 3-component accelerometer force balance system in a hypersonic shock tunnel