Injection cooling of blunt bodies flying at high Mach numbers

A study of transpiration cooling of blunt bodies such as a hemicylinder is made by solving Navier-Stokes equations. An upwind, implicit time-marching code is developed for this purpose. The study is conducted for both perfect-gas and real-gas (chemical equilibrium) flows. Investigations are carried out for a special wall condition that is referred to as no heat flow into the wall condition. The effects of air injection on wall temperature are analyzed. Analyses are carried out for Mach numbers ranging between 6-10 and Reynolds numbers ranging between 10(6)-10(7). Studies are made for spatially constant as well as spatially varying mass injection rate distributions, White cold air injection reduces the wall temperature substantially, transpiration cooling is relatively less effective when the gas is in chemical equilibrium.

Study of transpiration cooling over a flat plate at hypersonic Mach numbers

Transpiration cooling over a flat plate at hypersonic Mach numbers is analyzed using Navier-Stokes equations, without the assumption of an isothermal wall with a prescribed wall temperature. A new criterion is proposed for determining a relevant range of blowing rates, which is useful in the parametric analysis. The wall temperature is found to decrease with the increasing blowing rate, but this effect is not uniform along the plate. The effect is more pronounced away from the leading edge. The relative change in the wall temperature is affected stronger by blowing at high Reynolds numbers. (AIAA)

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

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.

Studies on the effects of test gas on the flow field around large angle blunt cone flying at hypersonic Mach number

The effect of the test gas on the flow field around a 120degrees apex angle blunt cone has been investigated in a shock tunnel at a nominal Mach number of 5.75. The shock standoff distance around the blunt cone was measured by an electrical discharge technique using both carbon dioxide and air as test gases. The forebody laminar convective heat transfer to the blunt cone was measured with platinum thin-film sensors in both air and carbon dioxide environments. An increase of 10 to 15% in the measured heat transfer values was observed with carbon dioxide as the test gas in comparison to air. The measured thickness of the shock layer along the stagnation streamline was 3.57 +/- 0.17 mm in air and 3.29 +/- 0.26 mm in carbon dioxide. The computed thickness of the shock layer for air and carbon dioxide were 3.98 mm and 3.02 mm, respectively. The observed increase in the measured heat transfer rates in carbon dioxide compared to air was due to the higher density ratio across the bow shock wave and the reduced shock layer thickness.

Direct measurement of phase of foreward-scattered light using polarization heterodyne interferometer

We describe direct measurement of phase of ballistic photons transmitted through objects hidden in a turbid medium using a polarization interferometer employing a rotating analyzer. The unwrapped phase difference measurements from interferometry was possible for medium levels of turbidity and accurate phase measurement from the sinusoidal intensity was not detectable when l/l* is increased beyond 4.3. The measured phase on reconstruction using standard tomographic algorithms resulted in the recovery of the refractive index profile of the object hidden in the turbid medium.

Measurement of Heat Transfer Rate on Backward-Facing Steps at Hypersonic Mach Number

Two backward-facing models with step heights of 2 and 3 mm are used to measure the convective surface heat transfer rates by using platinum thin-film gauges, deposited on Macor inserts. Heat transfer rates have been theoretically calculated along the flat plate portion of a model using the Eckert reference temperature method. The experimentally determined surface heat transfer rate distributions are compared with theoretical and numerical estimations. Experimental heat flux distribution over a flat plate model showed good agreement with the reference temperature method at stagnation enthalpy range of 0.8-2 MJ/kg. Theoretical analysis has been used for downstream of a backward-facing step using Gai's nondimensional analysis. It has been found from the present study that approximately 10 and 8 step heights are required for the flow to reattach for 2 and 3 mm step height backward-facing step models, respectively, at a nominal Mach number of 7.6.

Effect of gas injection on drag and surface heat transfer rates for a 30 degrees semi-apex angle blunt body flying at Mach 5.75

Effect of coolant gas injection in the stagnation region on the surface heat transfer rates and aerodynamic drag for a large angle blunt body flying at hypersonic Mach number is reported for two stagnation enthalpies. A 60° apex-angle blunt cone model is employed for this purpose with air injection at the nose through a hole of 2mm diameter. The convective surface heating rates and aerodynamic drag are measured simultaneously using surface mounted platinum thin film sensors and internally mounted accelerometer balance system, respectively. About 35–40% reduction in surface heating rates is observed in the vicinity of stagnation region whereas 15–25% reduction in surface heating rates is felt beyond the stagnation region at stagnation enthalpy of 1.6MJ/kg. The aerodynamic drag expressed in terms of drag coefficient is found to increase by 20% due to the air injection.

Numerical analysis of transpiration cooling in carbon dioxide atmosphere at hypersonic Mach numbers

The numerical solutions are obtained for skin friction, heat transfer to the wall and growth of boundary layer along the flat plate by employing two dimensional Navier-Stokes equations governing the hypersonic flow coupled with species continuity equations. Flow fields have been computed along the flat plate in CO2 atmosphere in the presence of transpiration cooling using air and carbon dioxide.

Penetration Characteristics of Air, Carbon Dioxide and Helium Transverse Sonic Jets in Mach 5 Cross Flow

An experimental investigation of sonic air, CO2 and Helium transverse jets in Mach 5 cross flow was carried out over a flat plate. The jet to freestream momentum flux ratio, J, was kept the same for all gases. The unsteady flow topology was examined using high speed schlieren visualisation and PIV. Schlieren visualisation provided information regarding oscillating jet shear layer structures and bow shock, Mach disc and barrel shocks. Two-component PIV measurements at the centreline, provided information regarding jet penetration trajectories. Barrel shocks and Mach disc forming the jet boundary were visualised/quantified also jet penetration boundaries were determined. Even though J is kept the same for all gases, the penetration patterns were found to be remarkably different both at the nearfield and the farfield. Air and CO2 jet resulted similar nearfield and farfield penetration pattern whereas Helium jet spread minimal in the nearfield.

LISA as a dark energy probe

Recently, it has been shown that the inclusion of higher signal harmonics in the inspiral signals of binary supermassive black holes (SMBH) leads to dramatic improvements in the parameter estimation with Laser Interferometer Space Antenna (LISA). In particular, the angular resolution becomes good enough to identify the host galaxy or galaxy cluster, in which case the redshift can be determined by electromagnetic means. The gravitational wave signal also provides the luminosity distance with high accuracy, and the relationship between this and the redshift depends sensitively on the cosmological parameters, such as the equation-of-state parameter w = p(DE)/rho(DE) of dark energy. Using binary SMBH events at z < 1 with appropriate masses and orientations, one would be able to constrain w to within a few per cent. We show that, if the measured sky location is folded into the error analysis, the uncertainty on w goes down by an additional factor of 2-3, leaving weak lensing as the only limiting factor in using LISA as a dark energy probe.

Force Measurements on a Reflex Cambered Delta Wing

AREFLEX spanwise cambered delta wing with a conical camber designed for M= 1.4, using the method of Ref. 1, was tested at the design Mach number as well as off-design Mach number M=0.15 and 2.3, respectively. The test results are compared with those of a plane wing and also with the available theoretical results at the design condition. At subsonic speed, the cambered wing has less lift at a given incidence and higher lift-to-drag ratio at a given lift than the plane wing, while at supersonic speeds, both of these quantities were less on the cambered wing. At supersonic speed, at the design incidence and Mach number, there is good agreement between results from theory and experiment. The center of pressure on the cambered wing is ahead of that on the plane wing at subsonic speed, while the reverse is true at supersonic speeds. Finally, it is found that over a useful range of lift the cambered wing is aerodynamically more efficient at subsonic speeds, and less so at supersonic speeds, than the plane wing.

Unsteady nonsimilar compressible laminar two-dimensional and axisymmetric boundary-layer flows

Unsteady nonsimilar laminar compressibletwo-dimensional and axisymmetric boundarylayer flows have been studied when external velocity varies arbitrarily with time and the flow is nonhomentropic. The governing nonlinear partial differential equations with three independent variables have been solved using an implicit finite difference scheme with quasilinearization technique from the origin to the point of zero skin-friction. The results have been obtained for (i) an accelerating stream and (ii) a fluctuating stream. The skin friction responds to the fluctuations in the free stream more compared to the heat transfer. It is observed that Mach number and hot wall cause the point of zero skin friction to occur earlier whereas cold wall delays it.