Experimental investigation of cowl shape and location on inlet characteristics at hypersonic Mach number

An experimental study is presented to show the effect of the cowl location and shape on the shock interaction phenomena in the inlet region for a 2D, planar scramjet inlet model. Investigations include schlieren visualization around the cowl region and heat transfer rate measurement inside the inlet chamber.Both regular and Mach reflections are observed when the forebody ramp shock reflects from the cowl plate. Mach stem heights of 3.3 mm and 4.1 mm are measured in 18.5 mm and 22.7 mm high inlet chambers respecively. Increased heat transfer rate is measured at the same location of chamber for cowls of longer lenghs is indicating additional mass flow recovery by the inlet.

Measurement of shock stand-off distance on a 120 degrees blunt cone model at hypersonic Mach number in Argon

The flow around a 120 degrees blunt cone model with a base radius of 60mm has been visualised at Mach 14.8 and 9.1 using argon as the test gas, at the newly established high speed schlieren facility in the IISc hypersonic shock tunnel HST2. The experimental shock stand off distance around the blunt cone is compared with that obtained using a commercial CFD package. The computed values of shock stand off distance of the blunt cone is found to agree reasonably well with the experimental data.

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.

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.

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

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.

Film cooling at hypersonic Mach numbers using forward facing array of micro-jets

Experimental investigations are carried out in the IISc hypersonic shock tunnel on film cooling effectiveness of a single jet (diameter 2 mm and 0.9 mm), and an array forward facing of micro-jets (diameter 300 mu m each) of same effective area (corresponding to the respective single jet). The single jet and the corresponding micro-jets are injected from the stagnation zone of a blunt cone model (58, apex angle and nose radius of 35 mm). Nitrogen and Helium are injected as coolant gases. Experiments are performed at freestream Mach number 5.9, at 0 degrees angle of attack, with a stagnation enthalpy of 1.84 MJ/kg, with and without injections. The ratios of the jet stagnation pressure to the freestream pitot pressure used in the present study are 1.2 and 1.45. Up to 50% reduction in surface heat transfer rate was observed with the array of micro-jets, compared to that of the respective single jet with nitrogen as the coolant, while the corresponding eduction was up to 37% for helium injection, with the schlieren flow visualizations showing no major change in the shock standoff distance, and thus no major changes in other aerodynamic aspects such as drag.

Shock tunnel testing of a Mach 6 hypersonic waverider

A waverider is a lifting body configuration whose upper surface is parallel to the free stream, and the lower surface aerodynamically so designed, that the resulting shock at the design Mach number, is always attached with the leading edge of the vehicle. This prevents spillage from high pressure (lower) surface to the low pressure (upper) surface.In the present study a conical waverider has been designed, fabricated and tested at Mach 6 in the IISc hypersonic shock tunnel HST2. The measurements show that the waverider has a lift to drag ratio of 4.28 at the designed Mach number. Exhaustive FEM and CFD studies are also carried out to complement the force measurements in the tunnel.

Schlieren visualization of shock wave phenomena over a missile-shaped body at hypersonic Mach numbers

The flow over a missile-shaped configuration is investigated by means of Schlieren visualization in short-duration facility producing free stream Mach numbers of 5.75 and 8. This visualization technique is demonstrated with a 41 degrees full apex angle blunt cone missile-shaped body mounted with and without cavity. Experiments are carried out with air as the test gas to visualize the flow field. The experimental results show a strong intensity variation in the deflection of light in a flow field, due to the flow compressibility. Shock stand-off distance measured with the Schlieren method is in good agreement with theory and computational fluid dynamic study for both the configurations. Magnitude of the shock oscillation for a cavity model may be greater than the case of a model without cavity. The picture of visualization shows that there is an outgoing and incoming flow closer to the cavity. Cavity flow oscillation was found to subside to steady flow with a decrease in the free stream Mach number.

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.

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.

Behaviour of turbulent Prandtl/Schmidt number in compressible mixing layer

The behaviour of turbulent Prandtl/Schmidt number is explored through the model-free simulation results. It has been observed that compressibility affects the Reynolds scalar flux vectors. Reduced peak values are also observed for compressible convective Mach number mixing layer as compared with the incompressible convective Mach number counterpart, indicating a reduction in the mixing of enthalpy and species. Pr-t and Sc-t variations also indicate a reduction in mixing. It is observed that unlike the incompressible case, it is difficult to assign a constant value to these numbers due to their continuous variation in space. Modelling of Pr-t and Sc-t would be necessary to cater for this continuous spatial variation. However, the turbulent Lewis number is evaluated to be near unity for the compressible case, making it necessary to model only one of the Pr-t and Sc-t..

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.