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This paper briefly reviews some national studies and new programs concerning hypersonic flight. The flight environment that will be encountered by this new class of hypersonic vehicles is described, and the fluid-dynamic and chemical phenomena that occur in hypersonic flight are examined. Ground-based facilities are briefly described, and their use in helping to validate the codes is examined.
Underexpanded axisymmetric jets are studied numerically using a full Navier-Stokes solver. Emphasis has been given to supersonic and hypersonic jets in supersonic and hypersonic ambient flows, a phehomenon previously being overlooked for the most part. The present work demonstrates that the shear layers and shock patterns in a jet plume can be captured without complicated viscous/inviscid and subsonic/supersonic coupling schemes. In addition, a supersonic pressure relief effect has been identified for underexpanded jets in supersonic ambient flows. While it is well known that an underexpanded jet in a quiescent ambience (or subsonic ambience) contains multiple shock cells, the present study shows that because of the supersonic pressure relief effect, an underexpanded jet in a supersonic or hypersonic ambience contains only one major shock cell.
The development of robust, accurate, and efficient implicit shock-capturing schemes for multidimensional compressible Navier-Stokes equations in the hypersonic and real gas flow regimes is presently undertaken by extending a class of implicit total variation-diminishing (TVD) schemes suitable for transonic and supersonic, multidimensional Euler and Navier-Stokes equations to hypersonic computations. Numerical aspects of TVD schemes are identified which affect the convergence rate for hypersonic Mach numbers and real gas flows, but which have a negligible effect on low Mach number or perfect gas flows.
A general discussion is given on the use of advanced computational fluid dynamics (CFD) in analyzing the hypersonic flow field around an airbreathing aircraft. Unique features of the hypersonic flow physics are presented and an assessment is given of the current algorithms in terms of their capability to model hypersonic flows. Several examples of advanced CFD applications are then presented.
Possible experimental facilities appropriate to a university environment that could make meaningful contributions to the solution of problems in hypersonic aerodynamics are investigated. Needs for the National Aerospace Plane and interplanetary flights with atmospheric aerobraking are used to scope the problem. Relevant events of the past two decades in universities and at the national laboratories are examined for their implications regarding both problems and prospects. Most striking is the emergence of computational fluid dynamics, which is viewed here as an equal partner with laboratory experimentation and flight test in relating theory with reality. Also significant are major advances in instrumentation and data processing methods, especially optical techniques. The direction of the study was guided by the concept of a companion program, i.e., the university effort should complement a major area of endeavor at NASA-Langley. Through this, both faculty and student participants gain a natural and effective working relationship. Existing and proposed major hypersonic aerodynamic facilities in industry and at the national laboratories are examined by type; hypersonic wind tunnels, arc-heated tunnels, shock tubes and tunnels, and ballistic ranges. Of these, the free piston tunnel and shock tube/tunnel are most appropriate for a university.
Linear stability theory for hypersonic boundary layers is presented. The theory is used to study the effects of real gas (under the assumption of local chemical equilibrium) and small nose bluntness on hypersonic boundary layer stability. It is found that chemical reactions have a stabilizing effect on the first mode instability and a destabilizing effect on the second mode instability in hypersonic boundary layers. There is also a tendency for the second-mode instability to shift to lower frequencies. The effect of small nose bluntness is found to be stabilizing.
Hypersonic computations are presently conducted with an extension of a class of high-resolution implicit TVD algorithms suited to transonic multidimensional Euler and Navier-Stokes equations. These conservative shock-capturing schemes, which are spatially second- and third-order, may be first- and second-order accurate in time and suitable for either steady or unsteady calculations. Attention is given to the enhancement of hypersonic flows' convergence rate and stability; accuracy and efficiency is achieved by these means for very complex two-dimensional hypersonic viscous and inviscid shock interactions.
The objective of the present investigation is to establish a benchmark experimental data base for a generic hypersonic vehicle shape for validation and/or calibration of advanced computational fluid dynamics computer codes. This paper includes results from the comprehensive test program conducted in the NASA/Ames 3.5-foot Hypersonic Wind Tunnel for a generic all-body hypersonic aircraft model. Experimental and computational results on flow visualization, surface pressures, surface convective heat transfer, and pitot-pressure flow-field surveys are presented. Comparisons of the experimental results with computational results from an upwind parabolized Navier-Stokes code developed at Ames demonstrate the capabilities of this code.
The Compressible Navier-Stokes (CNS) code was developed to compute external hypersonic flow fields. It has been applied to various hypersonic external flow applications. Here, the CNS code was modified to compute hypersonic internal flow fields. Calculations were performed on a Mach 18 sidewall compression inlet and on the Lewis Mach 5 inlet. The use of the ARC3D diagonal algorithm was evaluated for internal flows on the Mach 5 inlet flow. The initial modifications to the CNS code involved generalization of the boundary conditions and the addition of viscous terms in the second crossflow direction and modifications to the Baldwin-Lomax turbulence model for corner flows.
A new procedure is demonstrated for optimizing hypersonic wind-tunnel-nozzle contours. The procedure couples a CFD computer code to an optimization algorithm, and is applied to both conical and contoured hypersonic nozzles for the purpose of determining an optimal set of parameters to describe the surface geometry. A design-objective function is specified based on the deviation from the desired test-section flow-field conditions. The objective function is minimized by optimizing the parameters used to describe the nozzle contour based on the solution to a nonlinear least-squares problem. The effect of the changes in the nozzle wall parameters are evaluated by computing the nozzle flow using the parabolized Navier-Stokes equations. The advantage of the new procedure is that it directly takes into account the displacement effect of the boundary layer on the wall contour. The new procedure provides a method for optimizing hypersonic nozzles of high Mach numbers which have been designed by classical procedures, but are shown to produce poor flow quality due to the large boundary layers present in the test section. The procedure is demonstrated by finding the optimum design parameters for a Mach 10 conical nozzle and a Mach 6 and a Mach 15 contoured nozzle.
Computational results are presented for three issues pertinent to hypersonic, airbreathing vehicles employing scramjet exhaust flow simulation. The first issue consists of a comparison of schlieren photographs obtained on the aftbody of a cruise missile configuration under powered conditions with two-dimensional computational solutions. The second issue presents the powered aftbody effects of modeling the inlet with a fairing to divert the external flow as compared to an operating flow-through inlet on a generic hypersonic vehicle. Finally, a comparison of solutions examining the potential of testing powered configurations in a wind-off, instead of a wind-on, environment, indicate that, depending on the extent of the three-dimensional plume, it may be possible to test aftbody powered hypersonic, airbreathing configurations in a wind-off environment.
Turbulence modeling is generally recognized as the major problem obstructing further advances in computational fluid dynamics (CFD). A closed solution of the governing Navier-Stokes equations for turbulent flows of practical consequence is still far beyond grasp. At the same time, the simplified models of turbulence which are used to achieve closure of the Navier-Stokes equations are known to be rigorously incorrect. While these models serve a definite purpose, they are inadequate for the general prediction of hypersonic viscous/inviscid interactions, mixing problems, chemical nonequilibria, and a range of other phenomena which must be predicted in order to design a hypersonic vehicle computationally. Due to the complexity of turbulence, useful new turbulence models are synthesized only when great expertise is brought to bear and considerable intellectual energy is expended. Although this process is fundamentally theoretical, crucial guidance may be gained from carefully-executed basic experiments. Following the birth of a new model, its testing and validation once again demand comparisons with data of unimpeachable quality. This report concerns these issues which arise from the experimental aspects of hypersonic modeling and represents the results of the first phase of an effort to develop compressible turbulence models.
Work on hypersonic propulsion research using a ram accelerator is presented. Several different ram accelerator propulsive cycles have been experimentally demonstrated over the Mach number range of 3 to 8.5. The subsonic, thermally choked combustion mode has accelerated projectiles to near the Chapman-Jouguet (C-J) detonation velocity within many different propellant mixtures. In the transdetonative velocity regime (85 to 115 percent of C-J speed), projectiles have established a propulsive cycle which allows them to transition smoothly from subdetonative to superdetonative velocities. Luminosity data indicate that the combustion process moves forward onto the projectile body as it approaches the C-J speed. In the superdetonative velocity range, the projectiles accelerate while always traveling faster than the C-J velocity. Ram accelerator projectiles operating continuously through these velocity regimes generate distinctive hypersonic phenomena which can be studied very effectively in the laboratory. These results would be very useful for validating sophisticated CFD computer codes and in collecting engineering data for potential airbreathing hypersonic propulsive systems.
In this paper, preliminary qualitative and quantitative comparisons of the body weight of five hypersonic aircraft configurations are conducted. The five configurations are briefly described as follows: (1) a wing-and-body arrangement with a power-law, circular cross-section body and a delta wing; (2) an all-body vehicle with delta planform and elliptical cross-sections; (3) a wingless wave rider configuration; (4) a winged wave rider configuration; and (5) the spacewing concept, an oblique flying wing at low speed that yaws to 90 deg sweep and flies end-on at hypersonic speeds. The vehicles are defined by their external moldline geometries and by the interior arrangement of their fuel tanks and other components. Intersecting, circular-lobed tankage is used in vehicles with noncircular bodies. The nonusable volume of such concepts is calculated. The structural concept, structural materials, Thermal Protection System, and heat load are allowed to vary with vehicle longitudinal station. Relative strengths and weaknesses of the various hypersonic aircraft concepts in terms of body weight are summarized.
Developing methods of predicting flight control forces and moments for hypersonic vehicles, included a preliminary assessment of subsonic/supersonic panel methods and hypersonic local flow inclination methods for such predictions. While these findings clearly indicate the usefulness of such methods for conceptual design activities, deficiencies exist in some areas. Thus, a second phase of research was proposed in which a better understanding is sought for the reasons of the successes and failures of the methods considered, particularly for the cases at hypersonic Mach numbers. To obtain this additional understanding, a more careful study of the results obtained relative to the methods used was undertaken. In addition, where appropriate and necessary, a more complete modeling of the flow was performed using well proven methods of computational fluid dynamics. As a result, assessments will be made which are more quantitative than those of phase 1 regarding the uncertainty involved in the prediction of the aerodynamic derivatives. In addition, with improved understanding, it is anticipated that improvements resulting in better accuracy will be made to the simple force and moment prediction.
Piloting difficulties associated with conducting aircraft maneuvers in hypersonic flight are caused in part by the nonintuitive nature of the aircraft response and the stringent constraints anticipated on allowable angle of attack and dynamic pressure variations. An approach is documented that provides precise, coordinated maneuver control during excursions from a hypersonic cruise flight path and the necessary flight condition constraints. The approach is to achieve specified guidance commands by resolving altitude and cross range errors into a load factor and bank angle command by using a coordinate transformation that acts as an interface between outer and inner loop flight controls. This interface, referred to as a 'resolver', applies constraints on angle of attack and dynamic pressure perturbations while prioritizing altitude regulation over cross range. An unpiloted test simulation, in which the resolver was used to drive inner loop flight controls, produced time histories of responses to guidance commands and atmospheric disturbances at Mach numbers of 6, 10, 15, and 20. Angle of attack and throttle perturbation constraints, combined with high speed flight effects and the desire to maintain constant dynamic pressure, significantly impact the maneuver envelope for a hypersonic vehicle.
Research into hypersonic propulsion; i.e., supersonic combustion, was seriously initiated at the Langley Research Center in the 1960's with the Hypersonic Research Engine (HRE) project. This project was designed to demonstrate supersonic combustion within the context of an engine module consisting of an inlet, combustor, and nozzle. In addition, the HRE utilized both subsonic and supersonic combustion (dual-mode) to demonstrate smooth operation over a Mach 4 to 7 speed range. The propulsion program thus concentrated on fundamental supersonic combustion studies and free jet propulsion tests for the three dimensional fixed geometry engine design to demonstrate inlet and combustor integration and installed performance potential. The developmental history of the program is presented. Additionally, the HRE program's effect on the current state of hypersonic propulsion is discussed.
The effects of viscous interaction in the analysis and design of improved classes of viscous optimized hypersonic waveriders is examined. The Corda computer program is used to generate viscous optimized hypersonic waveriders from conical flow fields without viscous interaction. Each waverider is optimized for maximum L/D, and comparison studies are made between cases with and without viscous interaction. The results show that aerodynamic performance of the viscous interaction waveriders are reduced due mainly to a large increase in skin-friction drag associated with the viscous interaction phenomena that grows with increasing Mach number and altitude, but some of this performance loss can be recouped by including viscous interactions within the optimization procedure. When the waverider is optimized for viscous interaction, the shape can change dramatically. A significant result of the present work delineates on a velocity-altitude map the region where viscous interaction effects are significant for modern hypersonic waveriders by performing parametric runs to produce L/D, C sub L, and C sub D contour plots for Mach numbers from 6 to 30 at altitudes from 30 to 80 km.