Hypersonic Flight Regime Research Articles

Nonequilibrium Viscous Shock-Layer Heat Transfer with Arbitrary Surface Catalycity

An approximate yet accurate analysis of aerodynamic heating at a highly cooled stagnation region, which embraces the entire high-altitude hypersonic flight regime where nonequilibrium dissociation/recombination in the shock layer and finite surface catalysis effects are both important, is given. Closed-form relationships are given for the relative nonequilibrium effects that are convenient for cost-effective engineering studies of high-altitude hypervelocity vehicle heating.

Attitude Stability of Blunt-Body Capsules in Hypersonic Rarefied Regime

To assess the attitude stability of blunt-body reentry capsules in the hypersonic rarefied flight regime, a systematic numerical procedure has been developed and applied practically to the asteroid sample return capsule. The low-density flows around the capsule are solved by the direct simulation Monte Carlo (DSMC) method in the three-dimensional geometry to determine the low-density aerodynamic coefficients. The overall aerodynamic coefficients are defined by an empirical bridging formula from the results of the DSMC analysis and the continuum coefficients obtained by the modified Newtonian method. By the use of these overall coefficients, the six-degree-of-freedom dynamics of the capsule is analyzed along the reentry trajectory. It is shown that the static instability about the pitching axis occurs in the low-density environments when the center of gravity is located in the rear part of the capsule. However, the pitching motion is found to be stabilized safely by initiation of the spin around the capsule axis at appropriate frequencies.

In-Flight Visualization of Supersonic Flow Transition Using Infrared Imaging

Infrared visualization was used to obtain information on the state of the boundary layer of a natural laminar flow airfoil in supersonic flight. In addition to the laminar/turbulent transition boundary, the infrared camera was able to detect shock waves and present a time-dependent view of the flow field. Thermal and flow mechanisms on the surface in flight were modeled numerically to aid in the design of the flight experiment and facilitate analysis of recorded infrared footage. A commercially available infrared camera was adapted for airborne use in this application. Readily available infrared technology has the capability to provide detailed visualization of various flow phenomena in subsonic to hypersonic flight regimes

X-33 Hypersonic Aerodynamic Characteristics

X-33 Hypersonic Aerodynamic Characteristics

Path-attitude decoupling and flying qualities implications in hypersonic flight

The hypersonic flight regime shows a decoupling of pitch attitude and flight path responses to pitch control inputs. The primary factors causing this decoupling are identified, and it is shown that it is a combined effect of the atmospheric density gradient and the high speed of hypersonic flight. Furthermore, it is shown which are the resulting unique path-attitude response characteristics. Flying qualities implications and metrics used in existing criteria and requirements are considered. It is shown which effects result from the path-attitude decoupling in hypersonic on these flying qualities metrics.

Thrust Enhancement in Hypervelocity Nozzles by Chemical Catalysis

In the hypersonic flight regime, the air-breathing supersonic combustion ramjet (scramjet) has been shown to be a viable propulsion system. The current designs of scramjet engines provide performance benefits only up to a Mach number of 14. Performance losses increase rapidly as the Mach number increases. To extend the applicability of scram'jets beyond Mach 14, research is being conducted in the area of inlet and wave drag reduction, skin-friction and heat-transfer reduction, nozzle loss minimization, low-loss mixing, and combustion enhancement. For high Mach number applications, hydrogen is the obvious fuel choice because of its high energy content per unit mass in comparison with conventional fuels. These flight conditions require engines to operate at supersonic internal velocities, high combustor temperatures, and low static pressures. The high static temperature condition enhances the production of radicals such as H and OH, and the low-pressure condition slows the reaction rates, particularly the recombination reactions. High-temperature and low-pressure constraints, in combination with a small residence time, result in a radical-rich exhaust gas mixture exiting the combustor. At high Mach number conditions (due to low residence time), H and OH do not have enough time to recombine ; thus, a significant amount of energy is lost as these high-energy free radical are exhausted. The objective of the present study is to conduct a flowfield analysis for a typical nozzle geometry for NASP-type vehicle to assess for thrust enhancement in hypervelocity nozzles by substituting small amount of phosphine for hydrogen.

Nonequilibrium boundary-layer effects on the aerodynamic heating of hypersonic waverider vehicles

The aerodynamic heating of hypersonic waverider vehicle configurations is examined in the high-altitude hypersonic flight regime where nonequilibrium dissociation-recombination in the boundary-layer and finite surface catalysis effects are both important. Analytical treatment of these effects is given for highly cooled surfaces in three important heating regions: 1) the nose, 2) the swept-wing leading-edge attachment line, and 3) the windward body center line. Relationships for the relative nonequilibrium effects are developed for each region that enable both computer code-prediction validation and cost-effective parametric engineering studies.

Nonequilibrium-dissociated stagnation boundary-layer flow on an arbitrarily catalytic swept wing

The aerodynamic heating of a swept stagnation line is examined in the high-altitude hypersonic flight regime where nonequilibrium dissociation/recombination in the boundary layer and finite surface catalysis effects are both important. Assuming a highly cooled body (T w /T s <<1), we extend Fay and Riddell's well-known self-similar solution for nonequilibrium-dissociated flow to include the effects of sweepback and arbitrary surface catalycity. Specific results for heat transfer are presented for the same zero sweep baseline case that was used in Fay and Riddell's work. In the frozen gas-phase limit, it is found that the combined net effects of sweep angle A wipe out the dissociation effects altogether beyond about Λ ≃ 65 deg and thus eliminate the usual favorable effect of a noncatalytic wall in reducing nonequilibrium heat transfer. Regarding the gas-phase chemistry aspect, results for a completely noncatalytic wall show that notwithstanding the significant spanwise viscous dissipation heating effect on the dissociation rate the net sweep effect diminishes the gas-phase reaction effect to zero beyond Λ ≃ 60 deg, analogously to the surface reaction case.

Non-equilibrium boundary layer effects on the aerodynamic heating of hypersonic vehicles

The aerodynamic heating of a generic lifting hypersonic vehicle configuration is examined in that high altitude hypersonic flight regime where non-equilibrium dissociation/recombination in the boundary layer and finite surface catalysis effects are both important. Analytical treatment of these effects is given for highly-cooled surfaces in three important heating regions: the nose, the swept wing leading edge attachment line and the windward body centerline. Relationships for the relative non-equilibrium effects are developed for each region that enable both computer code-prediction validation and cost-effective parametric engineering studies.

Navier-Stokes simulations of Orbiter aerodynamic characteristics including pitch trim and bodyflap

An analysis of the longitudinal aerodynamics of the shuttle orbiter in the hypersonic flight regime is made through the use of computational fluid dynamics. Particular attention is given to establishing the cause of the 'pitching moment anomaly,' which occurred on the orbiter's first flight, and to computing the aerodynamics of a complete orbiter configuration at flight conditions. Data from ground-based facilities as well as orbiter flight data are used to validate the computed results. Analysis shows that the pitching moment anomaly is a real-gas chemistry effect that was not simulated in ground-based facilities, which used air as a test gas. Computed flight aerodynamics for the orbiter are within 5% of the measured flight values and trim bodyflap deflections are predicted to within 10%.

Morphology of standing oblique detonation waves

Proposals have been made to utilize stabilized oblique detonation waves (ODWs) for the propulsion of hypersonic air-breathing vehicles and hypervelocity mass launchers. There exists hypersonic flight regimes where premixing of fuel and air may be desirable or unavoidable due to finite chemical induction times. Consequently, it is essential to understand under what conditions detonations may occur in order to design supersonic combustors to either avoid or utilize them efficiently. A theoretical analysis is made of supersonic flow of a combustible gas mixture past a wedge or an inclined wall, which shows that, for approach velocities roughly 25% or more greater than the Chapman-Jouguet velocity of the reactant mixture, there exists a usefully wide range of turning angles within which ODWs may be attached or stabilized. For smaller wedge angles, either an incomplete ODW, shock-induced combustion, or no combustion at all may ensue. For larger wedge angles, the wave will detach and form an overdriven normal detonation or normal shock-induced combustion wave immediately upstream of the stagnation point, decaying off axis to either a single oblique Chapman-Jouguet wave, or bifurcating to form an oblique shock followed by a shock-induced deflagration.

Periodic optical cruise of an atmospheric vehicle

Since the steady-state cruise path of an idealized point mass model of an atmospheric vehicle operating in the hypersonic flight regime is dynamically not fuel minimizing, closed periodic paths are numerically determined. By application of second-order conditions for local optimality, a periodic extremal path for a flat Earth is shown to be locally minimizing and produces an improvement in fuel usage of 4.2% over the steady-state cruise path. Application of these second variational conditions to extremal paths for the spherical Earth failed. Nevertheless, these paths produce improved fuel performance over the associated steady-state cruise path. UEL efficient cruise trajectories for aircraft have been a subject of continuous theoretical interest and are becoming one of practical interest as well. The analysis of aircraft trajectories for fuel minimization was first performed in a reduced state space.13 By neglecting the altitude and flight path angle dynamics, the equations of motion of a point mass representation of the aircraft motion reduces to the energy-state approximation where energy and fuel mass are the state variables, thrust and velocity (or altitude) are considered the control variables, and range is the independent variable. The first-order dynamics or rates are represented by the rate of change in the energy and fuel with respect to a change in range. The hodograph3 is formed by determining the boundary of reachable rates for admisible values of the control variables at a given energy value (the rates are independent of the fuel). The steady-state cruise fuel performance is given by the value of the fuel mass rate where the hodograph crosses the zero energy rate axis. If the hodograph is not convex so that a straight line tangent to two points on the hodograph (called the convex hull3) crosses the zero energy rate axis at a smaller value of fuel mass rate than does the hodograph, then the control variables at the points of tangency are used to form a chattering control sequence which theoretically will improve fuel performance over the steadystate cruise path. This chattering sequence, first discussed in Ref. 1, is an unrealizable infinite frequency control sequence between two thrust levels and two altitude and velocity points on the energy manifold where the aircraft is aerodynamic or propulsion efficient. This chattering cruise is also referred to as the relaxed steady state cruise in Ref. 2. Since velocity and altitude chattering is unrealizable, altitude has been added as a state variable in Ref. 4 and thrust and flight path angle are considered the control variables. In Ref. 4 the small angle approximation applied to the flight path angle results in flight path angle and thrust appearing linearly in the aircraft dynamic model and cost function. These control variables which lie interior to their admissible control sets along the extremizing steady-state cruise path form what is called a doubly singular arc in the calculus of variations.5'6 By applications of the matrix generalized Legendre-Clebsch condition due to Robbins, 7 it is demonstrated in Ref. 8 that the steady-state cruise path is not minimizing..

Inlet flowfield analysis for high Mach number applications.

Flow fields of inlets for real or perfect gases determined through four methods for application to supersonic and hypersonic flight regimes

Thermal Degradation of a Char-Forming Plastic During Hypsersonic Flight

A detailed theoretical aerothermostructural analysis is presented of the problem of determining the rate of ablation of a typical thermosetting hydrocarbon plastic tha t is subjected to the severe heat transfer rates encountered at the stagnation point of a hypersonic re-entry vehicle. The thermal degradation of the plastic is treated as a chemical reaction in depth by means of a chemical kinetic equation for the pyrolysis process. The analysis includes a mathematical description of the flow of the chemically reacting gases through the porous char. Hypersonic boundary layer mass transfer effects including gas phase combustion are considered in the overall energy balance. Numerical results are obtained for the reaction zone thickness, the rate of gasification of the plastic, the pressure drop through the char, and the temperature distribution within the material. The thermal stresses and the mechanical stresses due to the gas pressure drop through the porous char are calculated, and the mechanism by which the char layer builds up and collapses with time is demonstrated theoretically for the first t ime. Finally, correlated results are obtained for the maximum surface temperature, the period of spallation, and the maximum char thickness as a function of the stagnation pressure and stagnation enthalpy for the entire hypersonic suborbital flight regime.

Recent Advances in Nonequilibrium Dissociating Gasdynamics

T PURPOSE of this paper is to review some recent advances in the study of gasdynamic problems including effects of chemical reactions. To provide a background for the study the general concepts shall be outlined briefly. The discussions of the recent developments are restricted to inviscid flow problems only, neglecting viscosity, heat conduction and diffusion. Particular attention is directed to recent advances in analyses of nonequilibrium dissociating gas flows. In the hypersonic flight regime, high stagnation enthalpies sufficient to cause dissociation are realized. When the time to reach equilibrium is comparable with the time it takes for a fluid particle to pass through the flow, then there exist regions of the flow field where nonequilibrium states are encountered. A brief survey of both the linear and the nonlinear methods of treatment of these nonequilibrium flows, including some new developments that have not appeared elsewhere, will be presented.