Hypersonic Mach Numbers Research Articles

Interaction Length Analysis of Hypersonic Shock–Boundary-Layer Interaction Including High-Enthalpy Effects

Shock-induced flow separation plays an important role in practical applications of hypersonic flows, like intake unstart and unsteady aeroelastic phenomena. The size of the shock–boundary-layer interaction (SBLI) region determines the extent of choking in engine inlet ducts. Recent scaling studies have shown that interaction length is a function of shock-induced pressure jump, upstream Mach number, Reynolds number, wall heating/cooling, and flow deflection angle at the shock wave. In this work, we analyze the interaction length scaling for different hypersonic SBLI cases at varying flow enthalpies. We perform Reynolds-averaged Navier–Stokes (RANS) simulations, using a shock-strength-dependent turbulence model, of hypersonic SBLI flows over a range of Mach numbers and at flow enthalpies where vibrational nonequilibrium starts to be important. The numerical results closely match the experimental data, validating the RANS predictions. The size of the separation bubble in such high-enthalpy cases is found to not follow the known scaling relations. We propose a modified scaling that collapses a wide range of hypersonic Mach numbers and freestream enthalpy conditions. We also derive the high-Mach-number limit to demonstrate that wall cooling/heating and geometry are the main factors affecting the size of the SBLI region in hypersonic flows.

Simulation of Hypersonic Shock-Boundary Layer Interaction Using Shock-Strength Dependent Turbulence Model

Shock-induced flow separation is important in hypersonic intake unstart and for unsteady aero-thermal loads. This work presents a computational study of shock-boundary layer interaction (SBLI) at hypersonic Mach numbers. The objective is to accurately predict the separation bubble size for oblique shock impingement on a turbulent boundary layer and for two-dimensional axisymmetric cone-flare SBLI cases. We use the Reynolds-averaged Navier Stokes method with an advanced shock-strength-dependent turbulence model. A recently developed transport equation-based shock sensor is employed to identify the location and strength of shock waves. The computational results are found to match experimental measurements of surface pressure and skin-friction coefficient for a series of test cases, highlighting the effects of Mach number, shock generator angle, and wall cooling on the size of the separation bubble. Results are also found to conform to the scaling of SBLI separation size found in the literature. This is one of the few such corroborations for hypersonic axisymmetric SBLI.

Influence of Sweep Angle on the Surface Pressure of Delta Wing Along Pivot Positions at Hypersonic Mach Numbers

Influence of Sweep Angle on the Surface Pressure of Delta Wing Along Pivot Positions at Hypersonic Mach Numbers

Design and Analysis Aerospike Blunt Body at Hypersonic Speed

Hypersonic vehicles have the drag and heating problem when they operate at high speeds, to avoid or minimize this problem, aero spikes are utilized. This investigation finds a solution for reducing the drag and heating. Using ANSYS fluent designed an aero spike, which has flow around a blunt body at hypersonic Mach number 6 with angle of attack zero. It is analyses for L/D ratio of 1.5 and 2. The front of the body is replaced by an aero spike which has weak oblique shocks and well-built detached shock. This develops a circular region between shock and blunt body. This works as a streamlined profile to reduce heat as well as drag.

INVERTED BRAYTON CYCLE ENGINE OPTIMIZATION FOR HYPERSONIC FLIGHT

Optimization of inverted Brayton cycle engine is made by particle swarm optimization method in this study. The optimum specific thrust value is reached by staying within the optimization constraints. When the total temperature of the cooling section is examined, it is seen that a temperature above the freezing temperature of the air is obtained. A very high specific thrust value, 451 N.s/kg is obtained at the hypersonic flight Mach Number (5 Mach) as a result of optimization. Temperature decrease with increasing altitude effect performance dramatically as positive. Low preburner exit total temperature and turbine total pressure ratio values and high afterburner exit total temperature values are desirable in terms of performance but values of total temperature at cooling section effect performance conversely in terms of specific thrust and specific fuel consumption

Effectiveness of Cone Angle on Surface Pressure Distribution along Slant Length of a Cone at Hypersonic Mach Numbers

In the present study, the prime attention is to numerically simulate the surface pressure distribution over the slant length of the cone at the various Mach numbers and for the considerable range of semi-cone angles. The Computational Fluid Dynamics (CFD) analysis is used to simulate the surface pressure distribution numerically. The hypersonic Mach numbers, semi-cone angle, and various locations along the slant length of a cone are considered parameters for this research work. The Mach numbers (M) considered for the research work are 5, 7, 9, 11, 13, and 15. The cone angle (θ) from 5° to 25° is considered. The results of pressure (P/Pa) are recorded at various locations (x/L) along the slant length of the cone as 0.1 to 1. The pressure distribution results are obtained by CFD analysis and compared with the analytical results available in the literature for Mach number 5. The findings obtained by CFD analysis and analytical results show good agreement. In this investigation, it is observed that the Mach number, cone angle, and slant length of the cone are the highly influential parameters for the surface pressure distribution. The pressure distribution over the slant length of the cone increases with an increase in the Mach number and the semi-cone angle.

Analytical and Numerical Simulation of Surface Pressure of an Oscillating Wedge at Hypersonic Mach Numbers and Application of Taguchi's Method

This paper aims to estimate the surface pressure of a wedge at hypersonic Mach numbers at a considerable angle of incidence. The Ghosh similitude, corresponding strip theory, and piston theory are used to determine the pressure distribution analytically, and the results are compared to those of the CFD analysis. The theory is valid when the shock wave is attached to the leading edge of the nose of the wedge. Pressure on the windward surface was considered in the analysis. The pressure on the Lee surface is neglected. The condition for the validity of the theory is that the Mach number M2 behind the shock wave is greater than 2.5. The parameters taken into account for the study are the wedge angle and Mach number. The range of wedge angle considered is from 5 to 25 degrees and the Mach number considered is from 5 to 15. The analytical and the CFD results are in good agreement. The findings indicate that the parameters like wedge angle and Mach number are influential parameters that influence the wedge surface static pressure. The surface static pressure rises with an increase in Mach number and wedge angle.

Performance Characterization of Single Expansion Ramp Nozzle at Hypersonic Mach Number

Single Expansion Ramp Nozzle (SERN) is a direct extension nozzle where the gas pressure flows from one place to other works only on one side. The SERN is a hypersonic speed demonstrator vehicle due to its advantages and operational limitations. The demonstrator vehicle is designed to attain a cruise flight at a speed of Mach 6.2. The operating limits given as constraints are the atmosphere limitations, i.e., within the lower atmosphere wherein the vehicle has to run on air-breathing scramjet engines. The project emphasizes the effect of performance parameters, namely thrust, lift and moment, due to variation in jet Mach number and jet pressure through a nozzle. The nozzle is considered separately due to the experimental limitations. The CAD model designed for the nozzle is taken. Only the vehicle's nozzle section is dissected from the whole model for the numerical simulation. The GAMBIT software is used to mesh the model and tocreate the domain space. The model is a 3D structure aligned to the 3 axis coordinate system, where the body's length is aligned along the x axis and width of the body along the z-axis. The cowl is attached to the combustion chamber exit, providing pitching stability to the vehicle. The model analysis is done in FLUENT, where the model to be solved is exported from gambit and imported to FLUENT. The cowl arrangement affects the nozzle performance and the effect of performance due to change in cowl geometry is studied. This paper presents the study of the performance parameter’s interactions with jet pressure, Mach number and cowl deflection of the nozzle

Busemann diffuser for supersonic ramjet engine with detonation combustion of hydrogen-air mixture

The expediency of using a Busemann diffuser in a scramjet engine with detonation combustion of hydrogen-air mixtures is studied. A method is developed for calculating an air intake that provides spontaneous initiation of oblique detonation in a hydrogen-air mixture under conditions of rarefied atmosphere at hypersonic freestream Mach numbers. The geometry of the Laval nozzle (convergent-expanding) with the Busemann diffuser for the free flow Mach number 9 is determined. It is shown that at an altitude of about 40 km, the detonation combustion of the hydrogen-air mixture in this nozzle can provide more than 0.4 tons of thrust and more than 35% efficiency. The mathematical model is based on unsteady two-dimensional Euler equations for axisymmetric multicomponent reacting gas flow. To simulate chemical transformations, the detailed kinetic scheme is used that takes into account 33 non-equilibrium reactions for nine components of the mixture. The heat capacity, enthalpy, and entropy of a mixture are calculated using the reduced Gibbs energy of the gas components. Numerical modeling is carried out according to the modified Godunov scheme of the second order of approximation in spatial variables.

Laser-induced schliere anemometry in a Mach 6 flow with collinear light entry.

In this work, we demonstrate velocity, Mach number, and static temperature measurements in a Mach 6 flow using the recently developed laser-induced schliere anemometry (LISA) technique. To our best knowledge, this represents the first application of LISA for characterizing flow in the hypersonic Mach number regime, and a comparison with known tunnel values is provided. The laser-induced schliere in this work are written from a distance of roughly 76 cm away from the final lens, much further than in previous work. Furthermore, the schliere are created from a laser beam introduced parallel to the collimated schlieren light, which is a new arrangement that could be useful for facilities with limited optical access. A discussion of setup limitations is provided. The mean core flow velocity determined from LISA is within 1% of the expected value from isentropic theory, while the mean Mach number measurements are within 1.6% of the M=5.85 value used in literature for the facility. Furthermore, the determined mean static temperature of the core flow is within 2.5% of the value measured simultaneously in the facility.

Prediction of Aerothermal Environment and Heat Transfer for Hypersonic Vehicles with Different Aerodynamic Shapes Based on C++

This research paper discusses constructing a unified framework to develop a full-rate scheme for hypersonic heating calculations. The method uses a flow tracing technique with normal phase vector adjustment in a non-structured delineated grid combined with empirical formulations for convective heat transfer standing and non-standing heat flow engineering. This is done using dev-C++ programming in the C++ language environment. Comparisons of the aerodynamic thermal environment with wind tunnel experimental data for the Space Shuttle and Apollo return capsules and standing point heat transfer measurements for the Fire II return capsule was carried out in the hypersonic Mach number range of 6 - 35 Ma. The tests were carried out on an 11th Gen Intel(R) Core(TM) i5-1135G7 processor with a valuable test time of 45 mins. The agreement is good, but due to the complexity of the space shuttle tail, the measurements are still subject to large errors compared to wind tunnel experiments. A comparison of the measured Fire-II return capsule standing-point heat values with the theory for calculating standing-point heat fluxes simulated using Fay & Riddell and wind tunnel experiments is provided to verify the validity of this procedure for hypersonic vehicle heat transfer prediction. The heat fluxes assessed using this method for different aerodynamic profiles of hypersonic vehicles agree very well with the theoretical solution.

Thermochemical effects on hypersonic shock waves interacting with weak turbulence

The interaction between a weakly turbulent free stream and a hypersonic shock wave is investigated theoretically by using linear interaction analysis (LIA). The formulation is developed in the limit in which the thickness of the thermochemical nonequilibrium region downstream of the shock, where relaxation toward vibrational and chemical equilibrium occurs, is assumed to be much smaller than the characteristic size of the shock wrinkles caused by turbulence. Modified Rankine–Hugoniot jump conditions that account for dissociation and vibrational excitation are derived and employed in a Fourier analysis of a shock interacting with three-dimensional isotropic vortical disturbances. This provides the modal structure of the post-shock gas arising from the interaction, along with integral formulas for the amplification of enstrophy, concentration variance, turbulent kinetic energy (TKE), and turbulence intensity across the shock. In addition to confirming known endothermic effects of dissociation and vibrational excitation in decreasing the mean post-shock temperature and velocity, these LIA results indicate that the enstrophy, anisotropy, intensity, and TKE of the fluctuations are much more amplified through the shock than in the thermochemically frozen case. In addition, the turbulent Reynolds number is amplified across the shock at hypersonic Mach numbers in the presence of dissociation and vibrational excitation, as opposed to the attenuation observed in the thermochemically frozen case. These results suggest that turbulence may persist and get augmented across hypersonic shock waves despite the high post-shock temperatures.

Correction to: Experimental investigation of blunt cone model at hypersonic Mach number 7.25

Correction to: Experimental investigation of blunt cone model at hypersonic Mach number 7.25

Estimation of Stiffness derivative for Ogives at Constant specific heat Ratio γ = 1.666 at Hypersonic Mach Numbers

Specific heat ratios play an essential role in the design of an ogive of an aircraft, which in turn makes it a priority point in estimating the stiffness derivative, keeping in mind the stability of an airplane. One such attempt is made here where the study of stiffness derivative is done for specific heat ratio =1.666 for other than the air. In some cases, higher values of specific heat ratios become a priority to stimulate the flow. For δ angles from 50 to 300, results have been obtained. For pivot locations near the nose, the pitching moment assumes higher values as compared to the pivot position near the shoulder of the ogive. The stability derivatives attain negative values for h = 0.6, and its importance is maximum at h = 1.

Experimental investigation of blunt cone model at hypersonic Mach number 7.25

Experiments were carried out in hypersonic shock tunnel in Defence Research and Directorate Laboratory at hypersonic Mach number of 7.25 using an 11.37° apex-angle blunt cone model. Heat flux measurement was carried out on cone model at different angles of attack with angle of rotation ϕ = 0° to 360° in steps of 45° with vacuum sputtered platinum thin film sensors. The measured experimental value of heat transfer data at stagnation point was compared with theoretical value estimated Fay and Riddell correlation. As angle of rotation was increased from ϕ = 0° to ϕ = 180°, the shock wave became closer to model surface due to high density ratio across the shock wave and consequently heat transfer rate became higher.

Investigation of Slender Bodies at Hypersonic Mach Numbers

Tests were carried in a shock tunnel to determine the heating rate and the wall pressure on a test model flying at hypersonic velocity. The experiments were performed at Mach M = 6.5 and a total enthalpy of 1.2 MJ/kg. Helium was used as the driver gas and air as the driven gas. The effective test time during the tunnel testing was 3.5 ms. The vacuum sputtered gages were used to evaluate the heating rate on the test model. The evaluated heating rate agrees well with numerical simulation. The experimentally measured pressure also agrees with computational fluid dynamics

Standing window of oblique detonation with pathological behaviour

Standing of an Oblique Detonation Wave (ODW) on a wedge within combustor is the prerequisite of thrust generation for ODW engine which is regarded as a novel and conceptual propulsion device with hypersonic flight Mach number. Usually a standing window of ODW is defined as the wedge angle ranged from the ODW detached angle from wedge (upper limit) to the angle that a Chapman-Jouguet (CJ) detonation occurs (lower limit). For pathological detonation cases, however, the CJ detonation cannot be achieved, and thus the lower limit of the standing window of ODW should be revisited. In present study, two types of reactions in hypersonic incoming flow that include the behavior of pathological detonation, that is, the single-step irreversible reaction with mole variation and the two-step irreversible reactions with exothermic process followed by endothermic process, have been used for studying standing behavior of ODW. The steady detonation polar analysis of ODW is carried out for both reaction systems. The results reveal that the reaction with more mole decrement and the reactions with stronger endothermic process show the pathological detonation feature and therefore modify the lower limit of standing window of ODW. Three equivalent parameters are proposed to quantitatively measure the standing window range of ODW from points of view of thermodynamics, Mach number of incoming flow and heat effect of reactions. It is found that the standing window of ODW is determined by the specific heat ratio, the overdrive degree of detonation and the endothermic level of the hypersonic incoming flow, regardless of whether the detonation is pathological or not.

On Unsteady Flow Analysis of a Round Spike Blunt Nose Afterbody in Mach 6 Flow

An aerospike, in front of a blunt body, has largely been deemed as a key passive control device for effectively reducing the wave drag and aerodynamic heating associated with high-speed flows. In addition, it has been reported that the presence of a spike brings in unsteadiness in the form of oscillation and pulsation to the flow characteristics. Past researchers, having investigated mainly the aerothermodynamic coefficients, have hinted towards the suppressing of such oscillations with the use of a round nose spike over a sharp spike, though a thorough and a concrete result is yet to be established together with offering a detailed explanation of the flow physics and its dependence on the spike’s geometric characteristics (spike-length to afterbody-diameter ratio, L/D). Numerical investigation has been carried out using axisymmetric Navier-Stokes laminar flow solver at hypersonic Mach number of 6.0. A round-tip spike with a flat face cylindrical afterbody have been simulated for spike length ratios of L/D = 0.5 – 2.0, with spike diameter to-afterbody diameter (d/D) of 0.1. The behavior and subsequent control of flow pulsation for a round (hemispherical) spike with varying L/D-ratio has been established.

High mach number drag analysis of a modern lightweight launch vehicle

This paper presents predictions of pressure and frictional drag using on the individual components of the VEGA C Light Launch Vehicle propelled by a solid rocket motor over a hypersonic Mach number range from five to eight at nominal atmospheric conditions associated with the corresponding stage of the ascent trajectory. Both motor-off and motor-on conditions are simulated to isolate the motor-on effects. A series of simulations were also performed in off-nominal conditions by varying the Reynolds number and mass flow rate to examine the effects on drag under more extreme conditions. The results indicate that the differences between the motor-off and motor-on cases in nominal conditions are mainly due to base drag. With the motor-on the base pressure increases thereby reducing the total drag. In the off-nominal conditions, when the Reynolds number is increased there is a reduction in drag, and when the nozzle mass flow rate is decreased there is an increase in drag. Conversely, when the Reynolds number is decreased and the nozzle mass flow rate is increased, the opposite effect occurs. However, for very low Reynolds numbers, an incipient separation can occur on the first stage, which also influences the drag.

Aerodynamic Similarity of Winged and Lifting Space Vehicles

The aerodynamic data of winged and lifting space vehicles are considered. Recently, there was an investigation with respect to the longitudinal aerodynamics of these vehicles in the supersonic and hypersonic Mach number regimes with the astonishing result that for the lift and drag coefficients a similarity with regard to the geometrical configuration could be observed. It was shown that by introducing a homogenizing reference area with which the related quantities were made dimensionless, fully similar solutions of the lift coefficient and partly similar solutions of the drag coefficients are achieved. In the present paper this investigation is extended to the pitching moment coefficient. Further a generalized fit equation on the basis of exponential functions is constructed, which is able to approximate all the longitudinal aerodynamic coefficients for the supersonic and hypersonic Mach numbers. Finally the potential of these findings are demonstrated by some examples.