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Recent molecular-beam experiments of high-velocity atomic oxygen (O), atomic nitrogen (N), and molecular oxygen (O2) impacting carbon material at high temperature produced detailed surface chemistr...

Hydrocarbon fuels offer optimal high energy per volume for scramjet applications for sustained hypersonic flight but require additional residence time due to slower ignition delays (compared with h...

Lateral–directional stability is a critical issue for the design of any kind of aircraft. For the hypersonic waverider, the nonaxisymmetric, flat, and slender geometric feature makes it susceptible...

Wall pressure fluctuations beneath a boundary layer are the main cause of structural vibration for hypersonic aircraft. To accurately predict the load environment of pressure fluctuations, a measur...

形態: カラー図版あり, Physical characteristics: Original contains color illustrations, Accepted: 2020-07-09, 資料番号: PA2110004000

This paper presents a methodology to estimate the discrepancy in the model output of untested coupled multiphysics systems, based on tests on related systems. Model predictions often exhibit discre...

The waverider has an extensive application prospect in the design of hypersonic vehicles due to its excellent aerodynamic efficiency. However, it is known that the original cone-derived waverider i...

The double swept waverider is a classic example of the planform-customized osculating-cone waveriders, delivering satisfactory performances over a wide range of velocities. The nonlinear increase i...

Shock-wave–boundary-layer interaction is one of the most important phenomena in hypersonic flights due to its ubiquitous presence and its extreme effects on the aerothermodynamic loads on the aircr...

There is a recent re-interest in hypersonic flight with special attention on the capability of numerical simulations. A recent set of experiments with equilibrium freestream conditions was performe...

Model discrepancy in coupled, time-dependent analyses is addressed with emphasis on aerothermoelastic response prediction for hypersonic aircraft. Time-dependent temperature data from historic hype...

This paper proposes a family of high-pressure capturing wing configurations that aim to improve the aerodynamic performance of hypersonic vehicles with large volumes. The predominant visual feature...

This paper introduces a local piston theory with viscous correction for the prediction of hypersonic unsteady aerodynamic loads at high attitudes and large Mach numbers. A semi-empirical relation accounting for the viscous interaction effects to determine the effective shape is proposed. The method is validated by applying to thin airfoils at various Mach numbers, angles of attack, and operating altitudes. Sample two- and three-dimensional aerodynamic forces calculations are conducted demonstrating this method. Furthermore, flutter boundary predictions for a two-dimensional airfoil and pitching-in-damping derivative evaluations for a three-dimensional waverider configuration are performed with this unsteady aerodynamic model. Compared with the Euler-based local piston theory, this model performs much better at high altitudes for a wide range of Mach numbers, angles of attack, and shapes. Results suggest the feasibility of using the effective shape of hypersonic vehicles to efficiently and accurately obtai...

The first-order multidimensional flux difference splitting method (Parent, B., “Multidimensional Flux Difference Splitting Schemes,” AIAA Journal, Vol. 53, No. 7, 2015, pp. 1936–1948) is here extended to second-order accuracy through the use of limiters. The proposed second-order flux functions are such that they collapse to the first-order multidimensional flux difference splitting stencil in the vicinity of discontinuities and tend toward second-order finite difference stencils in smooth parts of the solution. Various test cases of interest to hypersonic flight including supersonic ramp injectors, shock-induced boundary-layer separation, and hypersonic viscous layers reveal that the proposed method achieves as much as a four times increase in resolution per dimension over its first-order counterpart or over the second-order symmetric total variation diminishing schemes. When integrated through a block-implicit approach, this entails as much as a 40-fold increase in computational efficiency while not com...

The performance of a scramjet inlet sensor based on tunable diode laser absorption spectroscopy of oxygen in the A-band near 760 nm during a low-altitude, subsonic flight test is discussed. Though the scramjet did not reach the targeted hypersonic flight regime due to the malfunction of the first-stage booster, analysis of the spectral absorption data acquired during the descent flight below 3 km successfully determined pressure and temperature in the inlet and demonstrated that the system was capable of producing an adequate signal-to-noise ratio to allow temperature, pressure, and velocity measurements under the expected hypersonic flight conditions at altitudes up to at least 30 km. This is the first flight-tested tunable diode laser sensor for a scramjet inlet flow, and the robustness and stability of the sensor design in dynamic load conditions were made evident by the quality of the data recorded after a very unstable and high-g launch scenario.

The Hypersonic International Flight Research Experimentation (HIFiRE) program is a hypersonic flight test program executed by the Air Force Research Laboratory and Australian Defence Science and Technology Organisation. The principal goal of HIFiRE flight five is to measure hypersonic boundary-layer transition on a three-dimensional body. The HIFiRE flight tests are supported by a ground test campaign; this paper presents measurements of heat flux and boundary-layer transition in the Boeing/AFOSR Mach-6 Quiet Tunnel. This facility has been developed to provide quiet flow at high Reynolds number, with low noise levels comparable to flight. This tunnel’s good optical access enabled global measurement of the heat flux by means of temperature-sensitive paint. Two modes of transition were observed: transition along the centerline, suspected to arise from the amplification of second-mode waves in the boundary layer, and transition roughly halfway between the centerline and leading edges, probably due to the bre...

‡Slightly blunted cones with smooth tips and roughened frusta were flown in the NASA Ames hypersonic ballistic range through quiescent air environments. Global surface intensity (temperature) distributions were optically measured and analyzed to determine transition onset and progression over the roughened surface. Real-gas Navier-Stokes calculations of model flowfields, including laminar boundary layer development in these flowfields, were conducted to predict key dimensionless parameters used to correlate transition on such configurations in hypersonic flow. For distributed roughness elements totally immersed within the laminar boundary layer, the critical roughness Reynolds number correlating approach was found to well model transition onset and progression over the roughened conic frusta. The critical value of the roughness Reynolds number for transition was found to be 269  20%, in agreement with the critical value of 250  20% determined in earlier experiments for transition on rough, blunt bodies in hypersonic free flight.

The need for high-speed flight-control devices has surged with recent interest in hypersonic flight. Plasma-based devices offer actuation times that are orders of magnitude smaller than conventional mechanical actuators. The plasma jet, which uses energy deposition to generate a high-speed jet, is evaluated for flight control. The jet is created by pulsing a cavity with energy deposition. The gas expands through a converging nozzle, inducing a jet flow. This research focuses on characterizing the forces generated by a single pulse of the plasma jet, assuming energy is deposited uniformly throughout the cavity and the jet exits to quiescent flow. A one-dimensional analytical model is developed to predict the force and impulse generated by the jet as well as the temporal evolution of the pressure, density, and temperature in the cavity. A relation between the dimensionless energy deposition and the dimensionless impulse is developed and verifiedwith computational results. It is shown that the dimensionless impulse generated by the jet is essentially independent of the dimensionless geometric parameters of the cavity. Additionally, a simplified analysis shows that the force from an array of plasma jets is sufficient to replace a conventional aerodynamic flap.

The supersonic combustion ramjet, or scramjet, is the engine cycle most suitable for sustained hypersonic flight in the atmosphere. This paper examines a key decision in the design of the inlet or intake of these hypersonic airbreathing engines, namely, the level of compression to be performed. Too much compression can lead to onerous system level issues including the need for bleed or variable geometry, while too little compression can result in low cycle efficiency and poor combustion of fuel. An analysis of the important factors that affect the choice of scramjet inlet compression ratio has been performed for hydrogen-fueled scramjets at Mach 6, 8, 10, and 12. It was found that contrary to classical thermodynamic analyses, scramjet cycle efficiency reaches an optimum at a relatively low compression ratio between 50 and 100 for all Mach numbers. Practical constraints related to nonequilibrium flow effects, inlet starting, and boundary-layer separation were also shown to prompt a desire for low compression ratio. The lower limit on compression was found to be set by the need to complete the combustion reaction in the available engine length and is therefore dependent on engine scale. On the basis of these factors it is recommended that scramjet inlet compression ratio be set to the minimum that satisfies the robust combustion requirement, with the caveat that it not be below 50 in order to maintain high cycle efficiency. For typical wind-tunnel-scale engines, this results in a requirement for the inlet to compress airflow entering the combustor to a pressure of approximately 1/2 atm, regardless of the flight Mach number.

S HOCK wave/boundary-layer interactions (SWBLIs) are a common feature of supersonic/hypersonic flight, and the unsteadiness of strongly separated interactions can lead to rapid fatigue of structural panels as well as inlet instability and unstart. To mitigate these problems, there is interest in developing techniques for controlling the separatedflowunsteadiness by using both passive and active control techniques. Previous SWBLI control work has focused on reducing the size of the separated flow and/or shifting the frequency of the interaction unsteadiness to a band that does not coincide with the resonant frequency of structural panels. A detailed survey of the variousmeans used for controlling SWBLI until the late 1980s is given in [1]. Recently, plasma-based actuators have been used by researchers for active control of SWBLI, since these actuators have several inherent desirable features such as high bandwidth and no moving parts. For example, previous researchers have used surface-mounted arc discharges [2] and glow discharges with external magnetic fields [3,4] to achieve control of reflected SWBLI. Wang et al. [5] used surface-mounted arc discharges with external magnetic fields and demonstrated the weakening of the separation shock strength in front of a compression ramp. Recently, arc discharges have been employed by Grossman et al. [6] and subsequent researchers [7–11] to generate a pulsed synthetic jet, which they termed a spark jet. The spark-jet design was modified by Narayanaswamy et al. [12] to extend the pulsing frequency to the kilohertz range. They termed the actuator a pulsed-plasma jet since the term spark implies a thermal discharge, which was not the case at the pressures used in their study (and in the current work). Narayanaswamy et al. [12] performed a detailed parametric study of the velocity and temperature characteristics of the pulsed-plasma jets. They reported a jet-exit velocity of about 300 m=s and a bulk gas temperature in the range 600–1000 K for the range of discharge currents tested. The same pulsed-plasma-jet array actuator is used in the present work to control the separation shock of a SWBLI generated by a compression ramp in a Mach 3 flow.

DOI: 10.2514/1.J050617 The goal of the United States Air Force to field durable platforms capable of sustained hypersonic flight and responsive access to space depends on the ability to predict the response and the life of structures under combined aerothermal andaeropressure loading. However,current predictive capabilities are limitedfor these conditions due in part to the inability to seamlessly address fluid-thermal-structural interactions. This study aims to quantify the significance of a frequently neglected interaction, namely: the mutual coupling of structural deformation and aerodynamic heating, on response prediction. The quasi-static response of a carbon–carbon skin panel is investigated. It is found that the significance of this coupling depends largely on the in-plane boundary conditions, since increasing resistance to thermal expansion results in buckling and increasing deflections into the flow. Including these deformations in aerodynamic heating results in O10% increase in peak temperature and O100% increase in surface ply failure index for deflections O1% of panel length. In these cases, the locations of peaktemperaturesandstressesaresignificantlyaltered.Finally,neglectingdeformationsintheaeroheatinganalysis results in the prediction of snap-through for a gradual heating trajectory, whereas, inclusion leads to a higher mode dominated, dynamically stable response.

H YPERSONIC flight began in February 1949 when a WAC Corporal rocket was ignited from a U.S.-captured V-2 rocket [1]. In the six decades since this milestone, there have been significant investments in the development of hypersonic vehicle technologies. The NASA X-15 rocket plane in the early 1960s represents early research toward this goal [2,3]. After a lull in activity, the modern era of hypersonic research started in the mid-1980s with the National Aerospace Plane (NASP) program [4], aimed at developing a single-stage-to-orbit reusable launch vehicle (RLV) that used conventional runways. However, it was canceled due mainly to design requirements that exceeded the state of the art [1,5]. A more recent RLV project, the VentureStar program, failed during structural tests, again for lack of the required technology [5]. Despite these unsuccessful programs, the continued need for a low-cost RLV, as well as the desire of the U.S. Air Force (USAF) for unmanned hypersonic vehicles, has reinvigorated hypersonic flight research. An emergence of recent and current research programs [6] demonstrate this renewed interest. Consider, for example, the NASA Hyper-X experimental vehicle program [7], the University of Queensland HyShot program [8], the NASA Fundamental Aeronautics Hypersonics Project [9], the joint U.S. Defense Advanced Research Projects Administration (DARPA)/USAF Force Application andLaunch fromContinentalUnited States (FALCON) program [10], the X-51 Single Engine Demonstrator [11,12], the joint USAF Research Laboratory (AFRL)/Australian Defence Science and Technology Organisation Hypersonic International Flight Research Experimentation project [13], and ongoing basic hypersonic research at the AFRL (e.g., [14–20]). The conditions encountered in hypersonic flows, combined with the need to design hypersonic vehicles, have motivated research in the areas of hypersonic aeroelasticity and aerothermoelasticity. It is evident from Fig. 1 that hypersonic vehicle configurations will consist of long, slender lifting body designs. In general, the body, surface panels, and aerodynamic control surfaces are flexible due to minimum-weight restrictions. Furthermore, as shown in Fig. 2, these

In this work the dynamic thermoelastic response of a slab with finite thickness subjected to aerodynamic heating is investigated. The system of generalized governing equations with generalized initial and boundary conditions is solved by the method of finite integral transformation, and the analytic expressions of the transient temperature and dynamical stresses in the slab are obtained. The finite integral transformation method shows its advantage and convenience to deal with the complex boundary conditions of a dynamic thermoelastic problem. Moreover, by introducing the aerodynamic heating and taking into account a typical Mach number curve for hypersonic flight, the analytical solution is obtained to simulate the dynamic thermoelastic response of the slab in hypersonic flight environment. Calculations are carried out for zirconium diboride to obtain transient temperature and dynamical stresses induced by aerodynamic heating.

The testing of aeroelastically and aerothermoelastically scaled wind-tunnel models in hypersonic flow is not feasible; thus, computational aeroelasticity and aerothermoelasticity are essential to the development of hypersonic vehicles. Several fundamental issues in this area are examined by performing a systematic computational study of the hypersonic aeroelastic and aerothermoelastic behavior of a three-dimensional configuration. Specifically, the flutter boundary of a low-aspect-ratio wing, representative of a fin or control surface on a hypersonic vehicle, is studied over a range of altitudes using third-order piston theory and Euler and Navier-Stokes aerodynamics. The sensitivity of the computational-fluid-dynamics-based aeroelastic analysis to grid resolution and parameters governing temporal accuracy are considered. In general, good agreement at moderate-to-high altitudes was observed for the three aerodynamic models. However, the wing flutters at unrealistic Mach numbers in the absence of aerodynamic heating. Therefore, because aerodynamic heating is an inherent feature of hypersonic flight and the aeroelastic behavior of a vehicle is sensitive to structural variations caused by heating, an aerothermoelastic methodology is developed that incorporates the heat transfer between the fluid and structure based on computational-fluid-dynamics-generated aerodynamic heating. The aerothermoelastic solution procedure is then applied to the low-aspect-ratio wing operating on a representative hypersonic trajectory. In the latter study, the sensitivity of the flutter margin to perturbations in trajectory angle of attack and Mach number is considered. Significant reductions in the flutter boundary of the heated wing are observed. The wing is also found to be susceptible to thermal buckling.

Anewtypeofdirect-measuring skin-frictiongaugewasdevelopedforthehigh-speed,high-temperatureenvironmentofturbulentboundary-layere owsinsupersoniccombustionramjet (scramjet)engines, with tests progressing from ground tests to an actual hypersonic scramjet-integrated e ight vehicle. The skin friction gauge was specie cally developed for installation on the Hyper-X e ight vehicle (X-43A). The design was nonnulling, with a sensing head that was e ush with the model wall and surrounded by a small gap. Unlike previous skin-friction sensor designs, this gauge eliminated water cooling and gap-e lling oil, in addition to a radically different e exure design. The design was verie ed by repeatable tests in a well-documented Mach 2.4 cold e ow, with results within 10 ‐15% of Cf estimates from simple theory. The sensor was qualie ed for e ight through a rigorous series of environmental tests, including pressure, temperature, vibration, shock, acceleration, and heat e ux tests. Finally, the skin-friction gauge was tested in the Hyper-X engine model (HXEM), a full-scale ‐partial-width wind-tunnel model of the e ight vehicle engine, at Mach 6.5 enthalpy in a freejet facility at NASA Langley Research Center. Successful testing in the HXEM provided the e nal verie cation of the gauge before installation in the X-43A e ight vehicle engine.

Turbulent, viscous interaction theory is used to understand the hypersonic flow past an expansion corner. By assuming a pressure law, the boundary-layer properties of the flow are obtained through simultaneous solution of a displacement thickness relationship and a coupling equation relating the effects of incidence and displacement thickness to the effective body shape. The pressure law is obtained through examination of available experimental data. The form of the pressure law is found to depend on the incidence. The pressure decay is dependent on both viscous and incidence effects. For weak incidence, viscous effects are important throughout the interaction, and they produce a gentle pressure decay. For strong incidence, viscous effects are important only near the front of the corner. In this latter case, the pressure decays rapidly followed by a long asymptotic downstream region; the overall pressure distribution then appears more similar to the in viscid case. Nomenclature A B C^ = constant of proportionality in linear viscosity-temperature relationship K - hypersonic similarity parameter, M^cc Kv = modified hypersonic similarity parameter, M^ dy e(x)/ dx M = Mach number P(x) = nondimensional pressure, p^/p^ p = pressure Re = unit Reynolds number T = temperature U = velocity (x, y) = physical coordinates along and normal to the incoming flow a = expansion corner angle y = ratio of specific heats 8 = boundary-layer thickness 6* = boundary-layer displacement thickness ^ = dummy variable of integration p = density % = weak turbulent, viscous interaction parameter, (M9 C / Re*)115

A comparative numerical study of the coupled vs uncoupled thermal-structural analysis of viscoplastic structures is presented. Deformations are assumed to be infinitesimal, and the viscoplastic behavior is described by using a unified constitutive model. The Galerkin finite element method is utilized along with a time-marching scheme to predict the quasistatic thermomechanical behavior of inelastic structures with rate and temperature-dependent material properties. In the coupled analysis, the displacements and temperature distribution at every time step are computed at once by solving the system of equations directly

Transition experiments were conducted on preablated graphite nosetips in the AEDC Track G facility. Nosetip radius and freestream static pressure were varied to insure transition front movement from sonic-line to stagnation-point proximity. Surface temperature contours were measured with image-converter cameras and results interpreted with interactive graphics to define instantaneous transition front contours. Statistical analyses showed that probability of transition occurrence on any given ray, between the minimum and maximum observed run lengths to transition, was Gaussian. Mean transition front location was observed to progress forward with a power-law dependence on range static pressure; coincidently, transition-front asymmetry, defined by a standard deviation about the mean, was noted to decrease linearly. All results were found to be independent of nosetip radius.

The problem of pitching oscillating slender delta wings with detached shock waves in hypersonic flow is studied using Messiter's thin shock-layer theory. The amplitude of oscillation is assumed small and a perturbation method is employed. Closed-form simple formulas are obtained for the unsteady pressure field and for the aerodynamic derivatives of the delta wings which are valid for general frequencies. It is found that within the thin shock-layer approximation, the slender delta wings with detached shock waves pitching in hypersonic stream are always stable dynamically. An accurate perturbation solution to Meissiter's functional-differential equation, which is required in calculating the steady and unsteady flowfields, is also obtained.

Spurk, J. H. and Gerber, N., "Dust Collection Efficiency for Power Law Bodies in Hypersonic Flight," AIAA Journal, Vol. 10, June 1972, pp. 755-761. Peddieson, J. and Lyu, C.-H., "Supersonic Two-Phase Flow Past a Cone," Proceedings of the 10th Southeastern Seminar on Thermal Sciences, edited by Watts and Sogin, Tulane Univ., New Orleans, La., 1974, pp. 467-482. Barren, R. M. and Wiley, J. T., "Remarks on Dusty Hypersonic Wedge Flow," AIAA Journal, Vol. 17, March 1979, pp. 317-318. Barren, R. M. and Wiley, J. T., "Newtonian Flow Theory for Slender Bodies in a Dusty Gas," Journal of Fluid Mechanics, Vol. 108, July 1981, pp. 147-157.

Thermal protection system based on radiation cooling for high altitude cruising hypersonic flight, achieving zero net mass transfer

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.

General considerations concerning the material and shape of a hypei sonic radome reveal a series of unifying physical restrictions on such structures The aerodynamics of such shapes at hypersonic speeds admits of unifying simplification as well, for the entropy layer and the highly ionized i egion of flow are nearly coincident A computational procedure is outlined which provides an approximate map of the entropy layer density, temperature, etc , for flow fields in which thermochemical equilibrium is maintained Using these data to generate a dielectric constant profile completes the specification of the electromagnetic problem This portion of the problem is attacked on an approximate basis, which is expected to be more accurate than ray-tracing, since the interaction zone should be of the order of a wavelength in dimension Interpretation of the results in terms of boresight-error-limited and powerlimited trajectories is discussed cursorily, and appendices illustrate implementation of the calculations by digital computer

The problem of optimizing the flight trajectory of a rocket vehicle moving in a resisting medium, in a general gravitational force field, is considered. Optimal laws for the modulation of the aerodynamic and propulsive forces are formulated in terms of the primer vector, the vector adjoint to the velocity. Relations for flight along an intermediate-thrust arc and integrals of the motion for several cases of practical interest are derived.

Heat transfer problems with a ballistic vehicle entering into the atmospheres of near planets at hypervelocity entry conditions

A theoretical analysis is made of the long-period (phugoid) oscillations of a lifting vehicle in hypersonic flight up to and including orbital speeds in any atmosphere. These oscillations correspond to a direct exchange between the kinetic energy and the potential energy along an oscillating flight path governed primarily by the trimmed steady-state zero pitching moment. The expressions derived for the period and damping of these oscillations show that for all speeds near or greater than sonic the atmospheric density gradient has an important effect in decreasing the period and increasing the damping of the phugoid oscillations. As orbital speeds are approached, it is found that the decrease in gravitational attraction with altitude overcomes the effect of the decrease in atmospheric density with altitude so that with increasing speed the phugoid period asymptotically approaches the corresponding satellite orbit period. In addition, simple explicit expressions are derived for the so-called "short period" or angle-of-attack oscillations. These and the phugoid relations are shown to be in excellent agreement with the numerical calculations previously presented by Etkin.