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.

Drag reduction by a forward facing aerospike for a large angle blunt cone in high enthalpy flows

Drag reduction studies are conducted using a flat disc tipped aerospike for a 120-degree apex angle blunt cone model in high enthalpy flows. Accelerometer based force balance is used for the drag force measurement in the newly established free piston driven shock tunnel, HST3. Drag reduction upto about 58 percent has been achieved for Mach 8 flow of 5 MJ/kg specific enthalpy at zero degree angle of attack.

Drag reduction by counterflow supersonic jet for a blunt cone in high enthalpy flows

Counterflow supersonic jet is used as a drag reduction device during the experiments in free piston driven shock tunnel, HST3. Accelerometer based force balance is employed to measure the drag force experienced by the 60-degree apex angle blunt cone model without and with the supersonic jet opposing the hypersonic flow. It is observed that the drag force decreases with increase in injection pressure ratio until the critical injection pressure is reached. Maximum reduction in drag force of 44 percent is recorded at the critical injection pressure ratio 22.36. Further increase in injection pressure ratio has reduced the percentage drag reduction. Change in nature of the flowfield around the model has also been observed across the critical injection pressure ratio.

Effect of electric arc discharge on hypersonic blunt body drag

Experimental results on the effect of energy deposition using an electric arc discharge, upstream of a 60° half angle blunt cone configuration in a hypersonic flow is reported.Investigations involving drag measurements and high speed schlieren flow visualization have been carried out in hypersonic shock tunnel using air and argon as the test gases; and an unsteady drag reduction of about 50% (maximum reduction) has been observed in the energy deposition experiments done in argon environment. These studies also show that the effect of discharge on the flow field is more pronounced in argon environment as compared to air, which confirms that thermal effects are mainly responsible for flow alteration in presence of the discharge.

Experimental and numerical investigation of jet injection in a wall bounded supersonic flow

Numerical and experimental studies of a supersonic jet (Helium) inclined at 45 degrees to a oncoming Mach 2 flow have been carried out. The numerical study has been used to arrive at a geometry that could reduce an oncoming Mach 5.75 flow to Mach 2 flow and in determining the jet parameters. Experiments are carried out in the IISc. hypersonic shock tunnel HST2 at similar conditions obtained from numerical studies. Flow visualization studies carried out using Schlieren technique clearly show the presence of the bow shock in front of the jet exposed to supersonic cross flow. The jet Mach number is experimentally found to be approximate to 3. Visual observations show that the jet has penetrated up to 60% of the total height of the chamber.

Experimental investigation of heat transfer reduction using forward facing cavity for missile shaped bodies flying at hypersonic speed

Forward facing circular nose cavity of 6 mm diameter in the nose portion of a generic missile shaped bodies is proposed to reduce the stagnation zone heat transfer. About 25% reduction in stagnation zone heat transfer is measured using platinum thin film sensors at Mach 8 in the IISc hypersonic shock tunnel. The presence of nose cavity does not alter the fundamental aerodynamic coefficients of the slender body. The experimental results along with the numerically predicted results is also discussed in this paper.

Convective heat-transfer rate distributions over a missile shaped body flying at hypersonic speeds

Experiments are carried out with air as the test gas to obtain the surface convective heating rate on a missile shaped body flying at hypersonic speeds. The effect of fins on the surface heating rates of missile frustum is also investigated. The tests are performed in a hypersonic shock tunnel at stagnation enthalpy of 2 MJ/kg and zero degree angle of attack. The experiments are conducted at flow Mach number of 5.75 and 8 with an effective test time of 1 ms. The measured stagnation-point heat-transfer data compares well with the theoretical value estimated using Fay and Riddell expression. The measured heat-transfer rate with fin configuration is slightly higher than that of model without fin. The normalized values of experimentally measured heat transfer rate and Stanton number compare well with the numerically estimated results. (C) 2009 Elsevier Inc. All rights reserved.

Accelerometer-Based Force Balance for High Enthalpy Facilities

A single component accelerometer-based force balance is developed, calibrated, and used for high enthalpy applications. Functionality of this force balance, for such applications, is demonstrated for the first time during high enthalpy tests in a newly established free piston-driven shock tunnel, HST3, using a 60 degrees apex angle blunt cone model at 0 degrees angle of incidence. Usefulness of this force balance is also confirmed, for much complicated high enthalpy flow situations, during the drag reduction studies with counterflow supersonic jet from the stagnation point.

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.

Hypersonic wave drag reduction in re-entry capsules using concentrated energy deposition

Experiments are carried out in a shock tunnel at a nominal Mach number of 5.75 in order to study the effect of concentrated energy deposition on the drag force experienced by a 120° blunt cone. Electrical energy was deposited along the stagnation streamline of the model using a high voltage DC discharge circuit (1.5 – 3.5KW) and the drag force was measured by a single component accelerometer balance. Numerical simulations were also carried complimenting the experiments. These simulations showed a substantial drag reduction (20% ~ 65%) whereas the experiments show no appreciable reduction in drag

A new 6-component accelerometer force balance for short duration ground testing facilities

A new six-component accelerometer force balance is developed and used in the HST2 shock tunnel of Indian Institute of Science. Aerodynamic forces and moments for a hypersonic slender body measured using this balance system at a free stream Mach number of 5.75 and Reynolds number of 1.5 million and stagnation enthalpy of 1.5 and 2 MJ/kg are presented. These measured values compare well with the theoretical values estimated using modified Newtonian theory.

Surface pressure and heat transfer measurements in the unsteady separated hypersonic flow field over double cones

Reliable bench mark experimental database in the separated hypersonic flow regime is necessary to validate high resolution CFD codes. In this paper we report the surface pressure and heat transfer measurements carried out on double cones (first cone semi-apex angle = 15, 25 deg.; second cone semi-apex angle= 35, 68 deg.) at hypersonic speeds that will be useful for CFD code validation studies. The surface pressure measurements are carried out at nominal Mach number of 8.35 in the IISc hypersonic wind tunnel. On the other hand the surface heat transfer measurements are carried out at a nominal Mach number of 5.75 in the IISc hypersonic shock tunnel. The flow separation point on the first cone, flow reattachment on the second cone and the wild fluctuation of the transmitted shock on the second cone surface (25/68 deg. double cone) in the presence of severe adverse pressure gradient are some of the flow features captured in the measurements. The results from the CFD studies indicate good agreement with experiments in the attached flow regime while considerable differences are noticeable in the separated flow regime.

High speed digital holographic interferometry for hypersonic flow visualization

Optical imaging techniques have played a major role in understanding the flow dynamics of varieties of fluid flows, particularly in the study of hypersonic flows. Schlieren and shadowgraph techniques have been the flow diagnostic tools for the investigation of compressible flows since more than a century. However these techniques provide only the qualitative information about the flow field. Other optical techniques such as holographic interferometry and laser induced fluorescence (LIF) have been used extensively for extracting quantitative information about the high speed flows. In this paper we present the application of digital holographic interferometry (DHI) technique integrated with short duration hypersonic shock tunnel facility having 1 ms test time, for quantitative flow visualization. Dynamics of the flow fields in hypersonic/supersonic speeds around different test models is visualized with DHI using a high-speed digital camera (0.2 million fps). These visualization results are compared with schlieren visualization and CFD simulation results. Fringe analysis is carried out to estimate the density of the flow field.

Experimental Investigation of Heat Fluxes in the Vicinity of Protuberances on a Flat Plate at Hypersonic Speeds

Heat transfer rates measured in front and to the side of a protrusion on an aluminum flat plate subjected to hypersonic flow at zero angle of attack are presented for two flow enthalpies of approximately 2 MJ/kg and 4.5 MJ/kg. Experiments were conducted in the hypersonic shock tunnel (HST2) and free piston driven HST3 at a freestream Mach number of 8. Heat transfer data was obtained for different geometries of the protrusion of a height of 4 mm, which is approximately the local boundary layer thickness. Comparatively high rates of heat transfer were obtained at regions of flow circulation in the separated region, with the hottest spot generally appearing in front of the protuberance. Experimental values showed moderate agreement with existing empirical correlations at higher enthalpy but not at all for the lower enthalpy condition, although the correlations were coined at enthalpy values nearer to the lower value. Schlieren visualization was also done to investigate the flow structures qualitatively.