Hypersonic Flight Speeds Research Articles

Technological advancements in Pulse Detonation Engine Technology in the recent past: A Characterized Report

Pulse Detonation Engine (PDE), is an emerging and promising propulsive technology all over the world in the past few decades. A pulse detonation engine (PDE) is a type of propulsion system that uses detonation waves to combust the fuel and oxidizer mixture. Theoretically, a PDE can be operate from subsonic to hypersonic flight speeds. Pulsed detonation engines offer many advantages over conventional air-breathing engines and are regarded as potential replacements for air-breathing and rocket propulsion systems, for platforms ranging from subsonic unmanned vehicles, long-range transportation, high-speed vehicles, space launchers to space vehicles. This article highlights the operating cycle of PDE, starting with the fuel-oxidizer mixture, combustion and Deflagration to detonation transition (DDT) followed by purging. PDE combustion process, a unique process, leads to consistent and repeatable detonation waves. This pulsed detonation combustion process causes rapid burning of the fuel-oxidizer mixture, which cannot be seen in any other combustion process as it is a thousand times faster than any other mode of combustion. PDE not only holds the capability of running effectively up to Mach 5 but it also changes the technicalities in space propulsion. The present paper is the extension of the previous study which is also a well characterized status report of PDE in different areas. The present study deals with the categorization of the design approach, computations & simulations, flow visualization, DDT & Thrust enhancement, PDRE’s, experimental detonation engines with some of the experience and research undertaken in Punjab Engineering College under the complete supervision and guidance of Prof. Tejinder Kumar Jindal followed by applications of PDE technology.

Unsteady numerical analysis of drag on non-winged hypersonic vehicle during re-entry

Reentry vehicles, such as upper stages of ballistic missiles, experience severe aerodynamic loads during the reentry phase of trajectory. This phase is characterized by three aspects namely, the continuous change in atmospheric conditions, the hypersonic flight speed, and the unsteady flight of the vehicle itself. Understanding the aerodynamic features of the vehicle during reentry phase is crucial to the designers of such vehicles. The objective of the present study is to investigate the aerodynamic features of a generic bi-cone non-winged vehicle that accelerates during reentry through the upper layer of dense atmosphere. The investigation is conducted via numerically simulating the unsteady flow around the vehicle using a commercial transient laminar CFD solver. The unsteady simulation accounts for the temporal variation of freestream conditions and the linear acceleration of the vehicle during descent. The used CFD model is validated against available wind tunnel data of a similar vehicle. The results address the evolution of the flow field pattern as well as the temporal variation of drag acting on the vehicle during reentry.

A terahertz signal propagation model in hypersonic plasma sheath with different flight speed

A hypersonic plasma sheath could shield communication signals, which results in the so-called “blackout.” Hypersonic flights in near space used to suffer from the “blackout.” Terahertz (THz) communication is considered to be a potential solution to the blackout problem. On the other hand, it is still not clear that how the complex flight conditions impact the THz communication channel conditions in the plasma sheath. The relation between the occurrence of communication blackout and the flight conditions is unclear either. In the present study, the relation between the THz signal attenuation in plasma sheaths and the hypersonic flight speed was investigated using the numerical hypersonic hydrodynamic model. According to this study, the thickness of the plasma sheath decreases with the increasing flight speed. The electron density and the electron collision frequency increase with the flight speed. As a result, the THz signal attenuation increases with the flight speed. In addition, a new estimation model, which is able to predict the THz signal attenuation in the plasma sheath for arbitrary flight speed, was developed on the basis of the numerical study. The estimation model, in which the hypersonic hydrodynamics is absent, could help to significantly reduce the time consumption in predicting the THz communication channel conditions as well as the occurrence of the blackout.

Calculation analysis of the experimental data of HIFiRE-I using the computer code NERAT-2D

A computer simulation of the densities of convective heat fluxes to the surface of a blunted circular cone has been performed at hypersonic flight speeds. For comparison, the conditions implemented in the scientific program HIFiRE-I were used.Calculation studies were performed using the author's computer code NERAT-2D and the algebraic models of turbulent mixing of Baldwin – Lomax and Prandtl. A satisfactory agreement between the calculated and experimental data is shown.

Numerical simulation and performances evaluation of the pulse detonation engine

A pulse detonation engine (PDE) is a type of propulsion system that uses detonation waves to combust the fuel and oxidizer mixture. The engine is pulsed because the mixture must be renewed in the combustor between each detonation wave. Theoretically, a PDE can operate from subsonic up to hypersonic flight speed. Pulsed detonation engines offer many advantages over conventional propulsion systems and are regarded as potential replacements for air breathing and rocket propulsion systems, for platforms ranging from subsonic unmanned vehicles, long range transports, high-speed vehicles, space launchers to space vehicles. The article highlights elements of the current state of the art, but also theoretical and numerical aspects of these types of unconventional engines. This paper presents a numerical simulation of a PDE at h=10000 m with methane as working fluid for stoichiometric combustion, in order to find out the detonation conditions.

Experimental Investigation of Brazilian 14-X B Hypersonic Scramjet Aerospace Vehicle

The Brazilian hypersonic scramjet aerospace vehicle 14-X B is a technological demonstrator of a hypersonic airbreathing propulsion system based on the supersonic combustion (scramjet) to be tested in flight into the Earth’s atmosphere at an altitude of 30 km and Mach number 7. The 14-X B has been designed at the Prof. Henry T. Nagamatsu Laboratory of Aerothermodynamics and Hypersonics, Institute for Advanced Studies (IEAv), Brazil. The IEAv T3 Hypersonic Shock Tunnel is a ground-test facility able to produce high Mach number and high enthalpy flows in the test section close to those encountered during the flight of the 14-X B into the Earth’s atmosphere at hypersonic flight speeds. A 1 m long stainless steel 14-X B model was experimentally investigated at T3 Hypersonic Shock Tunnel, for freestream Mach numbers ranging from 7 to 8. Static pressure measurements along the lower surface of the 14-X B, as well as high-speed Schlieren photographs taken from the 5.5° leading edge and the 14.5° deflection compression ramp, provided experimental data. Experimental data was compared to the analytical theoretical solutions and the computational fluid dynamics (CFD) simulations, showing good qualitative agreement and in consequence demonstrating the importance of these methods in the project of the 14-X B hypersonic scramjet aerospace vehicle.

Research on Numerical Simulation of Support Interference for Dynamic Derivatives at Hypersonic Flight Speeds

The research on dynamic derivative under the dynamic unsteady condition is one of the most difficult aspects of the aircraft development process. For the special experiment of dynamic derivatives, no numerical simulation of support interference has yet been systematically studied. The numerical simulation of support interference for identification of dynamic derivatives with 7o blunted cone under the hypersonic condition was done in this paper. The 2-order Roe scheme and the dual-time-step method based on LU-SGS were respectively applied to discrete of the spatial and time derivative of the unsteady flow. A least square method based on the Etkin unsteady aerodynamic model was used for identification of dynamic derivatives. Hypersonic missile module was adopted as a verification example, and the numerical calculation results are consistent with the experimental results. For two different sting support forms with 7o blunted cone, the impact of sting support interference on dynamic derivatives was studied. Results show the moment coefficients of two support forms under static conditions are essentially the same, while there is a big difference in the pitch damping derivatives under dynamic conditions. The support interference of nonlinear aerodynamic loads resulted from the shock wave induced separation and other unsteady flow structures under dynamic conditions is far more complicated than that under static conditions, and the correction law of support interference under static conditions cannot be applied directly to the unsteady dynamic situations.

Analysis of non-linear aeroelastic response of a supersonic thick fin with plunging, pinching and flapping free-plays

The flutter of a 3-D rigid fin with double-wedge section and free-play in flapping, plunging and pitching degrees-of-freedom operating in supersonic and hypersonic flight speed regimes have been considered. Aerodynamic model is obtained by local usage of the piston theory behind the shock and expansion analysis, and structural model is obtained based on Lagrange equation of motion. Such model presents fast, accurate algorithm for studying the aeroelastic behavior of the thick supersonic fin in time domain. Dynamic behavior of the fin is considered over large number of parameters that characterize the aeroelastic system. Results show that the free-play in the pitching, plunging and flapping degrees-of-freedom has significant effects on the oscillation exhibited by the aeroelastic system in the supersonic/hypersonic flight speed regimes. The simulations also show that the aeroelastic system behavior is greatly affected by some parameters, such as the Mach number, thickness, angle of attack, hinge position and sweep angle.

Aeroelastic response of a 2-D airfoil in a compressible flow field and exposed to blast loading

This paper deals with the generation and use of proper aerodynamic indicial functions toward the aeroelastic formulation of two-dimensional lifting surfaces in the subsonic compressible, linearized transonic, supersonic and hypersonic flight speed regimes. The indicial function approach enables one to treat in an unified way (i.e. in the time and frequency domains) the subcritical aeroelastic response and the flutter instability of lifting surfaces. Validations of the aerodynamic model are documented and excellent agreements are reported. In addition, closed form solutions and aerodynamic derivatives for different flight speed regimes are obtained; comparisons, and results displaying the aeroelastic response to blast loads are presented, and pertinent conclusions are outlined.

Three-dimensional supersonic flow past blunt bodies with contour discontinuities, with account for equilibrium and frozen state of the gas in the shock layer

An algorithm is devised for calculating by the finite difference method the supersonic flow region for three-dimensional steady-state flow of a viscous gas past a blunted body with many contour discontinuities. The state of this gas at high hypersonic flight speeds can be characterized by equilibrium or frozen physicochemical processes.

Sublimation of graphite at hypersonic speeds

In this paper, a new theoretical model is presented for the sublimation of graphite at hypersonic flight speeds. The aerothermochemical interactions between dissociated air and graphite are treated by means of a nine-component model, including O, O2, N, N2, CO, CO2, C, C3, and CN. The mass-transfer rate, the heat-transfer rate, and the skin-friction coefficient are determined numerically and are then correlated by means of algebraic equations as a function of stagnation pressure, stagnation enthalpy, and wall temperature in the high-Reynoldsnumber stagnation-point flow regime.

Tail Shock Problem

Attention is drawn to unresolved difficulties concerning the mechanism of tail shock formation behind bodies of revolution at supersonic and hypersonic flight speeds. In the limit of infinite Reynolds number, the flow near the neck of the wake is assumed to approach a locally conical field with vertex at the point on the axis where the shock starts. A detailed treatment is given for the case where the flow downstream of the shock is axial and of large Mach number M, so that the semivertex angle τ of the shock is small, but Mτ is not, and the shock is not weak. The problem then involves a small entropy parameter, and a uniform first‐order approximation with respect to it is obtained. This may be the first purely analytical treatment of a conical shock that is not weak.The gas is found to suffer a precompression, upstream of the shock, such that the density vanishes at the inner boundary of the stream. It is shown that this result must be expected to be independent of all but the one assumption that the gas impinges on the axis at a nonzero angle. Accordingly, tail shock formation in two‐dimensional and axially symmetrical flows must be expected to differ in significant, still unknown ways.

NONEQUILIBRIUM STAGNATION POINT BOUNDARY LAYERS WITH ARBITRARY SURFACE CATALYCITY

An approximate theory of nonequilibrium stagnation point boundary layer flow on highly cooled blunt bodies at hypersonic flight speeds, including a finite atom recombination rate on the surface, is presented. The analysis is based on a simplified nonlinear representation of the reaction rate distribution across the boundary layer which is derived from the fact that recombination near the wall dominates the nonequilibrium behavior when the wall is highly cooled. Closed-form solutions for the conditions at the gas/solid interface, such as atom concentration at the wall and heat transfer, are obtained in terms of a universal function of a single, composite Damkohler number. This number represents the combined effects of homogeneous and heterogeneous reaction over a wide range of the various thermochemical parameters involved. The theory is found to be in good agreement with the numerical results for nonequilibrium flow reported by Fay and Riddell.

Ablation of Glassy Material Around Blunt Bodies of Revolution

The steady-state equations of mot ion for a th in layer of an incompressible glassy material on the surface of an ablating and radiating b lunt body are reduced to a first-order ordinary differential equation which is integrated numerical ly. This solut ion is coupled with the solut ion of the air boundary layer for both laminar and turbulent heat transfer with or wi thout mass vaporization of the ablating material . The distribution of the effective energy of ablation around the body is thus obtained for a cone cylinder with a hemispherical cap that re-enters the atmosphere at hypersonic flight speeds, and has quartz as the ablating material . It is found that the ablation process from turbulent heat ing is more efficient than from the laminar case because of increased vaporization. This solution of the equations of mot ion at the s tagnat ion point has been verified by arc wind tunnel experiments. The present state of development of the arc wind tunnel does not permit i ts use for experimental investigations of ablation around b lunt bodies under turbulent heat ing .

Laminar Heat Transfer Over Blunt-Nosed Bodies at Hypersonic Flight Speeds

This paper deals wi th two l imit ing cases of laminar heat transfer over blunt-nosed bodies at hypersonic flight speeds, or high s tagnat ion temperatures: (a) thermodynamic equil ibrium, in which the chemical reaction rates are regarded as fas compared to the rates of diffusion across streamlines; (b) diffusion as rate-governing, in which the volume recombination rates within the boundary layer are s low compared to diffusion across streamlines. In either case the gas density near the surface of a blunt-nosed body is m u c h higher than the density jus t outside the boundary layer, and the velocity and stagnation enthalpy profiles are m u c h less sensitive to pressure gradient than in the more familiar case of moderate temperature differences. In fact, in case (a), the nondimensionalized enthalpy gradient at the surface is represented very accurately by the classical zero pressure gradient value, and the surface heat-transfer rate distribution is obtained directly in terms of the surface pressure distribution. In order to i l lustrate the method , this solution is applied to the special cases of an unyawed hemisphere and an unyawed, b lunt cone capped by a spherical segment . In the opposite l imit ing case where diffusion is ratecontrolling the diffusion equation for each species is reduced to the same form as the low-speed energy equation, except that the Prandtl number is replaced by the Schmidt number . The simplifications introduced in case (a) are also applicable here, and the expression for surface heat transfer rate is similar; the maximum value of the ratio between the rate of heat transfer by diffusion alone and by heat conduction alone in the case of thermodynamic equil ibrium is given by: (Prandtl n o . / S c h m i d t no.)'. When the diffusion coefficient is es t imated by taking a reasonable value of a tom-molecule collision cross section this ratio is 1.30. Additional theoretical and (especially) experimental studies are clearly required before these s imple results are accepted.