Base pressure and heat transfer on planetary entry type configurations

An extraneous shock impinging on a blunt body in hypersonic flow is observed to alter the flow around the body and increase the local heat-transfer rate near the impingement point. A novel, quasi-static technique is developed to study this phenomenon. A hemispherical, glass model, equipped with platinum thin-film thermometers, is injected into a hypersonic tunnel through a slot in a variable-incidence flat plate. Analog networks provide graphs of the heat-transfer rates at various points on the model as a function of the model's position relative to the extraneous shock. Peaks in local heat-transfer rates up to 10 times the local, unperturbed, freestream values are recorded as the model traverses the shock. The peak heating is severest on the side of the model nearest the plate and increases with increasing shock strength, occurring over a narrow region where a shear layer or jet, originating at the intersection of the bow shock and the impinging shock, meets the model surface. A physical model is set up which predicts variations in shock interference patterns, in surface pressure distributions, and in the intensity and extent of the peak heating in accordance with experiment. Nomenclature M = Mach number P = pressure normalized with respect to freestream pressure q = heat-transfer rate, cal-cm~2-sec~1 r = radius of hemisphere, mm t = time, sec T = temperature, °C U = velocity normalized with respect to freestream velocity V = injection speed, m-sec1 x = distance along surface measured from stagnation point, mm

Heat loads on spacecraft traveling at hypersonic speed are of major interest for their designers. Several tests using temperature-sensitive paints (TSP) have been carried out in long duration shock tunnels to determine these heat loads; generally paint layers were thin, so that certain assumptions could be invoked to enable a good estimate of the thermal parameter ρck (a material property) to be obtained—the value of this parameter is needed to determine heat loads from the TSP. Very few measurements have been carried out in impulse facilities [viz. shock tunnels such as the High Enthalpy Shock Tunnel Gottingen (HEG)], where test times are much shorter. Presented here are TSP temperature measurements and subsequently derived heat loads on a ramp model placed in a hypersonic flow in HEG (specific enthalpy h 0 = ~3.3 MJ kg−1, Mach number M = 7.4, temperature T ∞ = 277 K, density ρ ∞ = 11 g m−3). A number of fluorescence intensity images were acquired, from which, with the help of calibration data, temperature field data on the model surface were determined. From these the heat load into the surface was calculated, using an assumption of a 1D, semi-infinite heat transfer model. ρck for the paint was determined using an insitu calibration with a Medtherm coaxial thermocouple mounted on the model; Medtherm ρck is known. Finally presented are sources of various measurement uncertainties, arising from: (1) estimation of ρck; (2) intensity measurement in the chosen interrogation area; (3) paint time response.

At hypersonic Mach numbers, the separated boundary layer from the model base develops as a shear layer and separates the outer in-viscid hypersonic flow from low subsonic flow in the base region. Characterizing such a complex mixed hypersonic flow of the base region in a short duration facility needs adequate runtime since base flow itself takes few milliseconds to establish as well as fast response sensors and instrumentation. In the present study fast response sensors are developed and time taken to establish base flow on a typical blunt body is calculated. These experiments were carried out in a shock tunnel at free stream Mach number of 6 and free stream Reynolds number of 0.2 million. To gain additional confidence on the base heat transfer distribution, steady pressure measurements were also carried out in a conventional hypersonic wind tunnel. The studies indicate that for the above test conditions, the heat transfer and pressure distribution on the blunt body base plane is nearly uniform. The time required for the base flow establishment is 1.27 milli seconds.

Numerical simulation of compression corner flows at Mach number 9

A program of experimental research and analysis was conducted to examine the heat transfer and pressure distributions in regions of shock/shock interaction over smooth and transpiration-cooled hemispherical noseshapes. The objective of this investigation was to determine whether the large heat transfer generated in regions of shock/shock interaction can be reduced by transpiration cooling. The experimental program was conducted at Mach numbers of 12 to 16 in the Calspan 48-Inch Shock Tunnel. Type 3 and type 4 interaction regions were generated for a range of freestream unit Reynolds numbers to provide shear layer Reynolds numbers from 10 exp 4 to 10 exp 6 to enable laminar and turbulent interaction regions to be studied. Shock/shock interactions were investigated on a smooth hemispherical nosetip and a similar transpiration-cooled nosetip, with the latter configuration being examined for a range of surface blowing rates up to one-third of the freestream mass flux. While the heat transfer measurements on the smooth hemisphere without shock/shock interaction were in good agreement with Fay-Riddell predictions, those on the transpiration-cooled nosetip indicated that its intrinsic roughness caused heating-enhancement factors of over 1.5. In the shock/shock interaction studies on the smooth nosetip, detailed heat transfer and pressure measurements were obtained to map the variation of the distributions with shock-impingement position for a range of type 3 and type 4 interactions. Such sets of measurements were obtained for a range of unit Reynolds numbers and Mach numbers to obtain both laminar and turbulent interactions. The measurements indicated that shear layer transition has a significant influence on the heating rates for the type 4 interaction as well as the anticipated large effects on type 3 interaction heating. In the absence of blowing, the peak heating in the type 3 and type 4 interaction regions, over the transpiration-cooled model, did not appear to be influenced by the model's rough surface characteristics. The studies of the effects of the transpiration cooling on type 3 and type 4 shock/shock interaction regions demonstrated that large surface blowing rates had significant effect on the structure of the flowfield, enlarging the shock layer and moving the region of peak-heating interaction around the body.

Waveriderwhich with its geometry designed for a specific Mach number optimally exploits the lift generated due to shock wavesis a promising aerodynamic configuration for hypersonic flight. A conically derived waverider is designed for a nominal Mach number of 6. Detailed aerodynamic tests on the designed wave rider are initiated. Since hypersonic flows are essentially characterized by high flow total enthalpies, their simulations require impulse facilities like shock tunnels. Experiments are performed in IISc Hypersonic Shock tunnel HST-2 at the design Mach number as well as at an offdesign Mach number of 8, with a flow total enthalpy of 2 MJ/kg. Heat transfer rates are important concern at such high total enthalpies. The surface heat transfer distribution on the waverider at design and off-design conditions, measured using platinum thin film sensors, are presented in this paper. Despite the same flow total enthalpy, higher heat transfer rates are observed at Mach 6 than at Mach 8.

This numerical study investigates the effects of adding a small amount of ignition promoters for controlling the wedge-induced oblique shock wave (OSW) to oblique detonation wave (ODW) transition in a premixed hydrogen–air mixture at hypersonic speeds. The time-dependent two-dimensional compressible Euler equations for multiple thermally perfect species with a reactive source term are solved using adaptive mesh refinement and detailed chemical kinetics. The wedge with a fixed angle of 26° exhibits abrupt to smooth transitions for freestream Mach numbers 7–9 (speeds 2.8–3.2 km/s) at a pressure of 20 kPa and a temperature of 300 K. The small amount (1000 PPM by vol.) of H2O2 and O3 is found to be effective at significantly reducing the initiation length for the oblique detonation transition for all Mach numbers, which suggests a practical approach to increase the operating flight range for oblique detonation wave engine with a finite length wedge. At Mach number 8, the abrupt OSW to ODW transition turns towards a smooth transition with a small amount of H2O2 and O3 addition. Comparatively, O3 addition was found to be effective in reducing the ODW initiation length by promoting reactivity behind even a weaker oblique shock at low Mach number 7, for abrupt transition, while H2O2 addition was more effective than O3 at high Mach numbers 8 and 9, during a smooth transition. The maximum 73% and 80% reduction in initiation length of ODW was observed with 10,000 PPM H2O2 and O3 addition, respectively, during an abrupt OSW to ODW transition at Mach 7.

The use of a monorail ro (; kE::t sled system is detailed and recomme nded as a method of obtaining aerodyna mic data in a regime of high Mach and Reynolds numbers .The aerodyna mic and mechanical design of the Sandia rocket sleds is described , and the data c ollection system discussed in detail.The system has been successfu lly used by Sandia Laborato ry t o a cquire pressure data on conecylinders .These data are compared with data from free-fligh t r ockP.ts, wind-tunn el tests, anrl computer calcula tions.A principal advantage of the sled testing method is the high degree of accuracy possible in determini ng actual test condition s, which include Mach number, static pressure , and static temperat ure .Explorato ry tests indicate that sleds may be used to collect other types of experiment al data such as heat-tran sfer and boundary -layer transition .The high-spee d sled testing technique appears to be the most promis ing method of simulatin g flight condition s at low altitudes at Mach numbers up to M = 8.To date, excellent data have been obtained at a velocity of 7430 ft/sec (M = 6.53) which to our knowledg e is the highest velocity ever attained by a ro cket sled.At this speed, and the Holloman Air Force Base track altitude, the Reynolds number was 42 x 106 per ft, and the -dynamic pressure, 55,000 lbs/ft2. I. Introduct ionDuring wind tunnel and f:;:-ee-fiigh t rocket tests of slightly blunted, slender ( 5 to 2 5 degree half-angle ) cone-cyli nder bodies, it was noted that at Mach numbers greater than 4, the surface pressure ratio, Pm/P cn (measure d surface pressure divided by free-stre am static pressure) , in the region just aft of the cone-cyli nder junction was a strong function of free-stre am Reynolds number.Attempts to gather wind tunnel data at Reynolds numbers over 107 per ft and Mach numbers over M = 4. 5 were thwarted by inadequat e high-pressure facilities and/ or inadequat e pressure measurin g capabiliti es to meet the accuracy requirem e nt of 2 percent of the pressure ratio .

This study presents flow simulation over the reentry capsule at supersonic and hypersonic speeds. Numerical algorithms solve for the CFD method, which is produced using help ANSYS Fluent 19.2. The using GPU core to get a solution faster. The main purpose – flow simulation and numerical analysis reentry capsule; understand the behavior of supersonic and hypersonic flow and its effect on the reentry capsule; compare temperature results for the range Mach numbers equals 2–6. This study showed results on velocity counters, on temperature counters and vector of velocity for range Mach numbers equals 2–6. This study demonstrates the importance of understanding the effects of shock waves and illustrates how the shock wave changes as the Mach number increases. For every solves, the mesh had adapted for pressure gradient and velocity gradient to get the exact solution. As a result of the obtained solution, it is found that a curved shock wave appears in front of the reentry capsule. The central part of which is a forward shock. An angular expansion process is observed, which is a modified picture of the Prandtl- Mayer flow that occurs in a supersonic flow near the sharp edge of the expanding region. It is revealed that with an increase in the Mach number, the shock wave approaches the bottom of the reentry capsule, and there is also a slope of the shock to the flow direction, with an increase in the Mach number. The relevance and significance of this problem for the design of new and modernization of old reentry capsules.

A thorough experimental investigation of shock/shock interferences in high Mach number flows

This paper is a continuation of the studies dealt with numerical simulations of supersonic turbulent separated flows. 2-D turbulent supersonic flows in the vicinity of the plane configuration were explored in [1], [2]. The results of numerical modeling of axisymmetric flows in vicinity of a cylinder-flare configuration at various Mach numbers (M =3,5,7) are presented in the paper. All the calculations were performed in conditions of the real physical experiments proposed in [3] as the test cases for AGARD Working Group 18 (1997). The Table shows the condition of the experiments investigated in the computations. Here α stands for flare angle, r is the cylinder radius, M∞ and Re 1 are the freestream Mach and unit Reynolds numbers, P 0 and T 0 are the freestream static pressure and stagnation temperature; δ and C f0 are the boundary layer thickness and the skin friction before the interaction, L denotes the laminar case and T is for the turbulent case.

An experimental investigation was conducted to determine the characteristics of the near wake of a blunt body in hypersonic flow. Base-pressure and body-surface pressure measurements are combined with hot-wire and pitot-pressure data to describe the free shear layer, the base-flow region, the recompression zone, and the initial development of a laminar wake. It is found that the postulate of thin shear layers is valid for Reynolds numbers greater than about 104, if the Mach number Me external to the shear layer is not large. Detailed measurements with a steady-state hot-wire behind a two-dimensiona l circular cylinder indicate that the maximum pressure rise occurs downstream of the stagnation point formed by the merging shear layers. The hot-wire measurements are corroborated by base-pressure data obtained with cylinders and wedges. Comparison between the experimental and theoretical results points out the importance of the base-flow temperature and the initial shear-layer profile in determining the observable characteristics of the near wake.

An investigation of effectiveness and heat transfer on a cylinder model with showerhead cooling has been conducted in a free jet test facility, using the transient liquid crystal technique. A three- and a four-row configuration, covering a region of ±21°, were chosen to study the cooling behaviour for zero and off-design incidences. Typical engine airfoil leading edge conditions were maintained for the freestream Reynolds and Mach numbers to 1.55e5 and 0.26 for the three-row, and 1.84e5 and 0.30 for the four-row configuration, at a turbulence intensity of 7%. The blowing and momentum rates ware varied from 0.4 to 1.8 and 0.1 to 1.9, at a density ratio of 1.65. At zero incidence it could be observed for both configurations that the highest effectiveness of about 0.3 was achieved at a blowing rate of 0.4. Negative incidence for the four-row configuration resulted in much higher effectiveness, being as high as 0.5 at a blowing rate of 0.9, with an associated high heat transfer. Much lower effectiveness was achieved at positive incidence, where the blowing rate of 0.9 showed the best cooling behaviour. Higher effectiveness — but also increased heat transfer — was in general observed for the four-row configuration, with a 60% higher massflow.

Corresponding to Oswatitsch’s Mach number independence principle the Mach number of hypersonic inviscid flows, , does not affect components of various non-dimensional formulations such as velocity and density, pressure coefficients and Mach number behind a strong shock. On this account, the principle is significant in the development process for hypersonic vehicles. Oswatitsch deduced a system of partial differential equations which describes hypersonic flow. These equations are the basic gasdynamic equations as well as Crocco’s theorem which are reduced for the case of very high Mach numbers, . Their numerical solution can not only result in simplified algorithms prospectively utilized to describe the flow around bodies flying mainly in the lower stratosphere with very high Mach numbers. It also offers a deeper understanding of similarity effects for hypersonic flows. In this paper, a solution method for Oswatisch’s equations for perfect gas, based on a 4-step Runge-Kutta-algorithm, is presented including a fast shock-fitting procedure. An analysis of numerical stability is followed by a detailed comparison of results for different Mach numbers and ratios of the specific heats.

An experimental investigation has been conducted to study the effectiveness of film cooling on radiative heat transfer at the stagnation point of a body traveling at hypersonic flow speeds. A simple flat-faced cylinder is chosen as the model for the present campaign. Thermocouples were utilized to assess convective heat transfer rates, and emission spectroscopic technique is employed to collect radiation data at the stagnation point of the model. Nitrogen is used as the coolant, and it is injected through a tiny hole at the body's stagnation point. It is observed that the coolant injection resulted in a decrease of around 20% in convective heat flux. An increase in the intensity of radiation is observed with the injection of nitrogen as a coolant gas.