Cone In Hypersonic Flow Research Articles

Effect of compressibility on instability of boundary layer over a spinning cone in hypersonic flow

Effect of compressibility on instability of boundary layer over a spinning cone in hypersonic flow

Flow alteration technique to mitigate wave drag by hemispherical aero-spike spike for large angle blunted cone in hypersonic flow

Flow alteration drag reduction by hemispherical aero-spike for a larger angle blunt cone at hypersonic speed (Mach = 8.0) is investigated by using ANSYS Fluent. Numerical studies were carried out using finite volume (FV) based compressible flow solver to solve the two-dimensional Reynolds-averaged Navier-Stokes (RANS) equation. In this study, the k-ωtwo-equation turbulence model was used. Air has been taken as flowing fluid for analysis. In this analysis, aero-spike tip of hemispherical shape has been taken. The length of the aero-spike spike is taken on the basis of four different lengths to base diameter (L/D) ratios, in which these ratios are 0.2, 0.5, 0.7 and 1.0 respectively. The shortest length aero-spike is 14 mm and the longest length is 105 mm. The numerical investigation found a drag reduction of about 23–76 % using aero-spike in a full planner 600 apex angle blunt cone.

Experimental Study on Cooling Efficiency of Hypersonic Optical Window

The cooling effectiveness of optical window influences the imaging quality of hypersonic vehicles. This study focuses on the supersonic film cooling efficiency on the optical window of a blunt cone in hypersonic flow. The experiments were conducted in a gun tunnel equipped with a Mach 8 nozzle providing a total pressure and temperature of 9 MPa and 900 K respectively. Three tangential 2D nozzles with different combinations of slot heights and Mach numbers were designed to detect the film cooling length under different injection pressures. The heat flux on window surface was measured by Thin Film Gauges and the flow field was monitored by schlieren technique. When the jet pressure matched the mainstream, the window was completely cooled, and the minimum mass flow rate was achieved when the slot height was 5mm and the jet Mach number was 2.5. If the pressure ratio of jet continues to increase, the heat flux density could be furtherly reduced, but the cooling efficiency of unit coolant mass flow decreased significantly. The data correlation results showed that the cooling efficiency presented a nonlinear relationship of second order polynomial with (x/Sh)λ-0.8 , and the effective cooling length of film was positively correlated with the cooling mass flow ratio (λ) and slot height of the nozzle. Besides, the increase of jet pressure resulted in thickening the mixing layer, which enhanced the heat insulation effect and reduced the heat flux as a result.

Heat reduction using conterflowing jet for a nose cone with aerodisk in hypersonic flow

Numerical simulations of a 2D axisymmetric aerodisked nose cone in hypersonic flow are conducted, and innovative techniques involving forward injection of the gas from the stagnation point of the sphere are investigated; techniques include the injection of various counterflowing jets (helium or carbon dioxide) as a coolant jet from the nose cone behind the aerodisk. In this study, the characteristics of the various jet conditions of a counterflowing jet on a cone surface were investigated numerically to improve performance of the jet on heat reduction at the surface of a nose. The compressible, unsteady, axisymmetric Navier–Stokes equations are solved with SST turbulence model for free stream Mach number of 5.75 at 0° angle of attack with and without gas injection. According to the investigation of various conditions of jets, important phenomena of flow field and some effective jet conditions are found. Heat transfer results show a significant reduction in heat flux, even giving negative heat flux for some low pressure ratio, indicating that the flow wetting the model had a cooling effect which could significantly impact thermal protection system design. The findings suggest that high-speed vehicle design and performance can benefit from the application of counterflowing jets as a cooling system.

Hypersonic interference heating on cones with short three-dimensional protuberances

Heat fluxes around short, three-dimensional protuberances on sharp and blunt cones in hypersonic flow were experimentally measured using platinum thin-film sensors deposited on macor inserts. A parametric study of different protrusion geometries and flow conditions were conducted. Excessive heating was observed at locations near the protrusion where increased vorticity is expected, with the hottest spot being presented at the foot of the protuberance immediately upstream of it. If left unchecked, these hot spots could prove detrimental to hypersonic flight vehicles. Z-type schlieren technique was used to visualize the flow features qualitatively. New correlations to predict the heat flux at the hot spot have been proposed.

Boundary-layer stabilization by an ultrasonically absorptive material on a cone in hypersonic flow: numerical investigations

The damping of acoustic second mode instabilities by passive porous surfaces is investigated numerically and compared with experiments for different hypersonic boundary-layer flows. The used geometry is a blunt 7° half-angle cone model with exchangeable nose. The cone surface consists partly of an ultrasonically absorptive material, which has a natural, random porosity. For the analyses of the second modes, the DLR stability code NOLOT, NOnLocal Transition analysis code, is used. The influence of the nose radius and the unit Reynolds number on the second modes is investigated using a smooth surface. These results of the smooth surface are compared with those of the porous surface to study the damping effect on the second modes and the transition shift. The numerical results are compared with wind tunnel measurements, which were performed in the DLR High Enthalpy Shock Tunnel Gottingen.

Pressure-sensitive paint on a truncated cone in hypersonic flow at incidences

The flow over a truncated cone is a classical and fundamental problem for aerodynamic research due to its three-dimensional and complicated characteristics. The flow is made more complex when examining high angles of incidence. Recently these types of flows have drawn more attention for the purposes of drag reduction in supersonic/hypersonic flows. In the present study the flow over a truncated cone at various incidences was experimentally investigated in a Mach 5 flow with a unit Reynolds number of 13.5×106m−1. The cone semi-apex angle is 15° and the truncation ratio (truncated length/cone length) is 0.5. The incidence of the model varied from −12° to 12° with 3° intervals relative to the freestream direction. The external flow around the truncated cone was visualised by colour Schlieren photography, while the surface flow pattern was revealed using the oil flow method. The surface pressure distribution was measured using the anodized aluminium pressure-sensitive paint (AA-PSP) technique. Both top and sideviews of the pressure distribution on the model surface were acquired at various incidences. AA-PSP showed high pressure sensitivity and captured the complicated flow structures which correlated well with the colour Schlieren and oil flow visualisation results.

Transient analysis of counterflowing jet over highly blunt cone in hypersonic flow

Understanding the characteristics of various Counterflowing jets exiting from a nose cone is crucial for determining heat load reduction and usage of this device in various conditions. Such jets can undergo several flow regimes during venting, from initial supersonic flow, to transonic, to subsonic flow regimes as the pressure of jet decreases. A bow shock wave is a characteristic flow structure during the initial stage of the jet development, and this paper focuses on the development of the bow shock wave and the jet structure behind it. The transient behavior of a sonic counterflow jet is investigated using unsteady, axisymmetric Navier–Stokes solved with SST turbulence model at free stream Mach number of 5.75. The coolant gas (Carbon Dioxide and Helium) is chosen to inject into the hypersonic air flow at the nose of the model. The gases are considered to be ideal, and the computational domain is axisymmetric. The jet structure, including the shock wave and flow separation due to an adverse pressure gradient at the nose is investigated with a focus on the differences between high diffusivity coolant jet (Helium) and low diffusivity coolant jet (CO2) flow scenarios.

Numerical Study on Heat Reduction of Various Counterflowing Jets over Highly Blunt Cone in Hypersonic Flow

In this paper, the effectiveness of counterflowing jets as heat-reduction devices for large-angle blunt cones flying at hypersonic Mach numbers is numerically simulated with various coolant jets. Different jet conditions have been chosen to investigate the effect of the counterflow jet on the surrounding flow field of nose cone. The compressible, unsteady, axisymmetric Navier-Stokes equations are solved with SST turbulence model for free stream Mach number of 5.75 at 0° angle of attack with and without gas injection. The coolant gas (air, Carbon Dioxide, and helium) is chosen to inject into the hypersonic flow at the nose of the model. The numerical results presented the surface heat reduction for different coolant jets. According to the investigation of various conditions of opposing jets, important phenomena of flow field and some effective jet conditions are found.

Effectiveness of Aerospike for Drag Reduction on a Blunt Cone in Hypersonic Flow

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

Aerodynamic drag reduction by heat addition into the shock layer for a large angle blunt cone in hypersonic flow

Reduction in aerodynamic drag for a large angle blunt cone flying at hypersonic Mach number by heat addition into the shock layer is demonstrated in HST2 hypersonic shock tunnel. The heat addition is achieved by the exothermic reaction of chromium atoms ablated from the stagnation region of the chromium coated blunt cone with the atomic oxygen behind the shock wave. The measurements show about 47% reduction in the drag coefficient for a 60° apex angle blunt cone in a Mach 8 flow of 3.4MJ∕kg specific enthalpy. The reduction in drag is measured using the accelerometer based force balance system and the heat addition into the shock layer is identified by the surface mounted thin film heat flux gauges and the corresponding movement of the shock wave is visualized by schlieren pictures.

Laminar Base Flow Downstream of Slender Cones in Hypersonic Flow

The base flow downstream of slender cones in a stream of perfect gas at Mach numbers 8 and 10 and Reynolds numbers 104 and 105 is numerically investigated. The calculated heat fluxes to the rear face of the body are compared with experimental data. It is shown that the friction and heat transfer coefficients increase without bound as the corner point is approached from both the lateral surface and the rear face, the sign of the latter coefficient being dependent on the body surface temperature factor.

Experimental study of the hypersonic turbulent boundary layer on a cold slender cone

This paper presents measurements of both the mean and time-resolved properties of a naturally turbulent boundary layer on a highly cooled sharp cone in hypersonic flow. The experiments were carried out on a 5 deg half-angle cone at zero angle of attack in a Mach 9.26 flow in a gun tunnel. The unit Reynolds number was 4.4 x lOVm and the wall-to-stagnation temperature ratio was 0.273. Conventional surface and flow field instrumentation, including a new compact fast responding total temperature probe, was used to acquire time- averaged data. Measurements of the mean and the fluctuating density within the boundary layer are included that were obtained using the electron beam fluorescence technique. Comparisons are made between the measured data and several laws applicable to low-Mach-numb er zero-pressure -gradient boundary layer flows. It is shown that measured data are correlated by the van Driest extension to the Crocco relationship and agree with the velocity defect law modified by Fernholz.1 As expected, the mixing length concept is shown to be too simple to predict the density fluctuation correctly for this flow. Fluctuation spectra differ from the result from incompressible flows. The spatial correlation of the fluctuations emphasizes that there is considerable energy in the small eddies.

Static and dynamic pitch stability of a blunted cone with forward facing aerodynamic spike in hypersonic flow

This paper presents experimental data highlighting the effects of a forward facing aerodynamic spike on the pitch stability of a blunted cone in hypersonic flow. Spike length and angle of attack variations on stability derivatives are described. Suitable application of the inviscid embedded Newtonian pitch stability prediction technique shows how the influence of spike aerodynamics may be successfully modelled. The paper extends the knowledge of aerodynamic spike flow effects, which have to date concentrated on aerodynamic drag reduction qualities at zero degrees angle of attack.

Concerning Lateral Dynamics of Flight on a Great Circle

References 1 lannuzi, F. A. and Weddington, E. D., Free Pressure across the of Sharp and Blunt 10-degree Cones at Mach Numbers from 15 to 20, TR 66-223, Dec. 1966, Arnold Engineering Development Center, Tullahoma, Tenn. 2 Softley, E. J. and Graber, B. C, An Study of the Pressure and Heat Transfer on the of Cones in Hypersonic Flow, AGARD, CP 19, May 1967. 3 Murman, E. M., Experimental Studies of a Hypersonic Cone Wake, AIAA Journal, Vol. 7, No. 9, Sept. 1969, pp. 1724-1730. 4 Cassanto, J. M. and Hoyt, T. L., Flight Showing the Effect of Mass Addition on Pressure, AIAA Journal, Vol. 8, No. 9, Sept. 1970, pp. 1705-1707. 5 Cassanto, J. M., Flight Test Pressure at Hypersonic Mach Numbers in Turbulent Flow, AIAA Journal, Vol. 10, No. 3, March 1972, pp. 329-331. 6 Cassanto, J. M., Radial Pressure Gradients in Flow, AIAA Journal, Vol. 5, No. 12, Dec. 1967, pp. 2278-2279. 7 Cassanto, J. M., Ratio on Pressure, AIAA Journal, Vol. 3, No. 12, Dec. 1965, pp. 2351-2352. 8 Sherman, M. M. and Nakumura, T., Flight Test Measurements of Boundary Layer Transition on a Non-Ablating 22° Cone, Journal of Scientific Research, Vol. 7, No. 2, Feb. 1970, pp. 137-142. 9 U.S. Standard Atmospheric Supplements, 1966, 15° N. Annual. 10 Ohrenberger, J. T. and Baum, E., Laminar Near Wake Solutions under Atmospheric Entry Conditions, AIAA Paper 72-116, San Diego, Calif., 1972. 11 Ohrenberger, J. T. and Baum, E., A Theoretical Model of the Near Wake of a Slender Body in Supersonic Flow, AIAA Journal, Vol. 10, No. 9, Sept. 1972, pp. 1165-1172. 12 Van Hise, V., of Variation in Pressure over the Reynolds Number Range in which Wake Transition Occurs for Two-Dimensional Bodies at Mach Numbers from 1.95 to 2.92, TN D-167, Nov. 1959, NASA. 13 Kavanau, L. L., Results of Some Based Pressure Experiments at Intermediate Reynolds Numbers with M = 2.84, Journal of Aerospace Sciences, Vol. 21, April 1954; also Base Pressure Studies in Rarefied Supersonic Flows, Preprint 529, Jan. 1955, Institute of Aerospace Sciences, 14 Batt, R. G. and Kubota, T., Experimental Investigation of Near Wakes behind 20° Wedges at M^ = AIAA Journal, Vol. 6, No. 11, Nov. 1968, pp. 2077-2083. 15 Chapman, D. R., An Analysis of Pressure at Supersonic Velocities and Comparison with Experiment, Rept. 1051, May 1950, NACA.

Ablation surface cross-hatching on cones in hypersonic flow

Ablation surface cross-hatching on cones in hypersonic flow

Interaction of sonic transverse jets with supersonic external flows

11 Griffith, B. J. and Lewis, C. H., A Study of Laminar Heat Transfer to Spherically Blunted Cones and Hemisphere-Cylinders at Hypersonic Conditions, AEDC-TDR-63-102, June 1963, Arnold Engineering Development Center. 12 Wagner, R. D., Jr. and Watson, R., Induced Pressures and Shock Shapes on Blunt Cones in Hypersonic TND-2182, March 1964, NASA. 13 Friberg, E. G., Pressure Distribution on a 5.6° Blunt Cone at Mach 14, ARL 68-02,08, Nov. 1968, Aerospace Research Lab. 14 Ericsson, L. E., Separated Flow Effects on the Static and Stability of Blunt Nosed Cylinder Flare Bodies, LMSC-667991, Contract NAS 8-5338, Dec. 1965, Lockheed Missiles & Space Co., Sunny vale, Calif. 15 Ericsson, L. E. and Reding, J. P., Dynamic Stability Problems Associated with Flare Stabilizers and Flap Controls, Journal of Spacecraft and Rockets, Vol. 7, No. 2, Feb. 1970, pp. 132-137. 16 Walchner, O. and Clay, J. T., Bluntness Effects on the Stability Derivatives of Cones in Hypersonic Flow, Transactions of the Second Technical Workshop on Stability Testing, Vol. 1, Paper 8, April 1965, Arnold Engineering Development Center, Arnold Air Force Station, Tenn. 17 Ericsson, L. E., Effect of Boundary Layer Transition on Vehicle Dynamics, Journal of Spacecraft and Rockets, Vol. 6, No. 12, Dec. 1969, pp. 1404-1409. 18 Rainbird, W. J., Turbulent Boundary-Layer Growth and Separation on a Yawed Cone, AIAA Journal, Vol. 6, No. 12, Dec. 1968, pp. 2410-2416. 19 Clay, J. T., Bluntness, Cone Angle, and Mach Number Effects on the Stability Derivatives of Slender Cones, ARL 67-0185, Sept. 1967, Aerospace Research Lab. 20 Ward, L. K, Jr. and Uselton, B. L., Dynamic Stability Results for Sharp and Blunted 10-deg Cones at Hypersonic Speeds, Transactions of the 3rd Technical Workshop on Stability Problems, Vol. Ill, Paper 2, Nov. 1968, NASA. 21 Ericsson, L. E. and Scholnick, I. M., Effect of Nose Bluntness on the Hypersonic Unsteady Aerodynamics of Flared and Conical Bodies of Revolution, Journal of Spacecraft and Rockets, Vol. 6, No. 3, March 1969, pp. 321-324. 22 Friberg, E. G. and Walchner, O., Pressure Correlation for Blunted Slender Cones, AIAA Journal, Vol. 7, No. 8, Aug. 1969, pp. 1618-1619. 23 Ericsson, L. E., Universal Scaling Laws for Nose Bluntness Effects on Hypersonic Unsteady Aerodynamics, AIAA Journal, Vol. 7, No. 12, Dec. 1969, pp. 2222-2227.

Transpiration and film cooling effects for a slender cone in hypersonic flow

Transpiration and film cooling effects for a slender cone in hypersonic flow

Simple formulas for unsteady pressure on slender wedges and cones inhypersonic flow

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

Unusual boundary-layer transition results on cones in hypersonic flow.

Subliming ablation effects on boundary layer transition for cones in hypersonic flow, discussing Reynolds numbers measured in wind tunnel tests