On supersonic combustion and hypersonic propulsion

Criteria for hypersonic airbreathing propulsion and its experimental verification

Standing oblique detonation for hypersonic propulsion: A review

Hypersonic airbreathing propulsion, specifically scramjet technology, represents a transformative advancement in high-speed flight. This literature review examines critical studies and technological developments that have contributed to our understanding and implementation of scramjet systems. The primary reference is the comprehensive resource provided by Utah State University, which offers detailed insights into various propulsion systems. Acknowledgements are due to Dora E. Musielak, Ph.D., for her contributions through the AIAA Training Course on Hypersonic Air Breathing Propulsion in 2018. Foundational efforts by Heiser and Pratt (1995) laid the groundwork for theoretical and practical aspects of hypersonic propulsion. Significant advancements include the development and testing of the X-43A vehicle, achieving Mach 10, and subsequent studies addressing materials, thermal protection, and aerodynamic heating challenges. Further research has optimized scramjet performance through advanced design techniques and computational simulations. This review highlights contributions from notable works on scramjet inlets, combustion systems, propulsion system airframe integration, and numerical simulations of thermodynamic nonequilibrium. The innovative design of the Rectangular-to-Elliptical Shape Transition (REST) scramjet inlet/engine is also discussed, showcasing the progress in hypersonic propulsion technology. In addition to reviewing the literature, this study presents a Python code developed for modeling scramjets or ramjets with options for normal shocks or combined shocks. The code allows users to study one of three models and analyze efficiency against inlet Mach number or turning angles. It also demonstrates the difference in performance under dual nature ramjet/scramjet operations, providing a practical tool for evaluating and comparing propulsion system efficiencies.

Hypersonic air-breathing propulsion has been a focus technology in hypersonic aviation in the past decades [1]. Three-dimensional cavity may act as the flame holder of a Scramjet engine in air-breathing hypersonic propulsion. An interesting three-dimensional cavity is “swallowtail” cavity which has a special inner shape like a swallowtail. With three-dimensional cavity in supersonic chamber, threedimensional vortexes may be organized optimally, and the exchange of mass, momentum and energy between cavity flow and supersonic flow may be enhanced to provide better performance of mixing and combustion[2]. Also, three-dimensional cavity may avoid the sharp heat release in local region of chamber and suppress the subsonic combustion oscillation induced by the cavity in a supersonic combustor. It is necessary to study the heat release distribution of a supersonic combustor with three-dimensional cavity.

During the 10 last years, a large Research and Technology effort has been led by MBDA and ONERA to develop knowledge on high-speed airbreathing propulsion and master associated technologies. Development of operational, civilian or military, application of the hypersonic airbreathing propulsion depends of two key points : development of needed technologies for the fuel-cooled structure of the propulsion system, capability to predict with a reasonnable accuracy and to optimise the aeropropulsive balance (or generalized thrust-minus -drag balance). The most part of the technology development effort can be led with available ground test facilities and classical numerical simulation (thermics, mechanics ...). On the contrary, before any operational application, it is mandatory to demonstrate our ability to predict the aeropropulsive balance (generalized thrust-minusdrag balance) of a hypersonic vehicle, providing sufficient margins to start a costly technological program. Considering this mandatory step, MBDA and ONERA are leading a specific scientific program, called LEA, organized as follows: • Define a methodology for the development of a hypersonic vehicle using ground tests and numerical simulation • Develop the required tools (experimental or numerical) for this purpose. • Apply this methodology to the development of a simplified, scientific experimental vehicle • Validate this methodology through a series of flight tests Started in January 2003, this program is planned to end in 2012 after 6 autonomous flight tests of the experimental vehicle in the Mach number range from 4 to 8. Copyright © 2003 by MBDA France and ONERA. Published by American Institute for Aeronautics and Aerospace,Inc., with permission. Introduction During the two past decades, a lot of system studies, generally based on large technology development efforts, have been performed around the world to assess the interest of combined propulsion (airbreathing + rocket) for space launcher application. In France, thanks to studies led by the French Space Agency (CNES) during eighties, then in the framework of PREPHA program (Research and Technology Program for Advanced Hypersonic Propulsion), it was clearly established that combined propulsion could have an interest for space launcher application only if airbreathing mode can provide a subsequent part of the total speed increment (e.g. airbreathing mode must ensure propulsion from at least Mach 1.5/2 to Mach 10/12) (Ref [1] to [7]). By another way, hypersonic airbreathing propulsion could have very interesting application for military purpose. As a matter of fact, flying at very high Mach number and altitude can strongly improve the penetration capability while having a high average speed allow to deal with critical time targets (Ref [8] to [10]). Nevertheless, the development of operational, civilian or military, application of the hypersonic airbreathing propulsion depends of two key points : • development of needed technologies for the propulsion system as a low weight, high robustness fuel-cooled structure for the combustor, • capability to predict with a reasonnable accuracy and to optimise the aero-propulsive balance (or generalized thrust-minus -drag). Technology development effort Even if technologies will finally need to be flight proven, a large part of the technology development effort can be led with available ground test facilities (Ref[11]) and classical numerical simulation (thermics, mechanics...). 12th AIAA International Space Planes and Hypersonic Systems and Technologies 15 19 December 2003, Norfolk, Virginia AIAA 2003-7031 Copyright © 2003 by MBDA France and ONERA. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.

For super- and hypersonic air-breathing propulsion, the air intake can be regarded as the key component which is responsible for the performance of the propulsion sub-system (PSS). Here, performance means the achievable thrust (via captured air mass flow and total pressure recovery), distance from buzz and/or backflow (i.e.: stability of the system represented by its ability to allow “a certain amount” of so called sub-critical spillage), and its drag, namely in the below shock-on-lip Mach number regime and, especially, in the transition regime where missile switches from the booster driven acceleration phase to the air-breathing propulsion system powered sustain phase. For this purpose, Bayern-Chemie has developed a wide range of air-intakes which cover all applications of supersonic air-breathing missile propulsion, ranging from air-to-ground, anti-radar missile, and ship-to-ship applications up to advanced, highly maneuverable beyond visual range air-to-air missiles. The types and configurations of air-intakes cover a broad design Mach number range (shock-on-lip Mach number) as well as various types of rotational-symmetric and two-dimensional air intakes, mounted to the fuselage of the missile in various positions (inverted, radial, perpendicular), and so-called chin-inlets, shrouding the missile fore-body (conical section of ogive-shaped nose) to a certain amount. On the base of the experiences gained, this article will present the results of various airintake development programs and the resulting findings with respect to performance (isolated and integrated), stability, and spillage and bleed drag.

Preface. Mixing and Combustion Issues in Hypersonic Air-Breathing Propulsion D.M. Bushnell. Experiments. Hypersonic Combustion -- Status and Directions G.Y. Anderson. Recent Experiments on Hypersonic Combustion in an Expansion Tube Test Facility J.I. Erdos. Supersonic Combustion Experiments in Free Piston Reflected Shock Tunnels R.G. Morgan. Fuel Disperson in Supersonic Airstreams B.C.R. Ewan. Reacting Free Shear Layers. Reacting Compressible Mixing Layers: Structure and Stability C.E. Grosch. Suppression and Enhancement of Mixing in High-Speed Reacting Flow Fields J.P. Drummond, P. Givi. Modeling Turbulent Scalar Mixing with Mapping Closure Methods S.S. Girimaji. Finite-Rate Chemistry Effects in Subsonic and Supersonic Combustion R.W. Pitz, T.M. Brown, T.S. Cheng, S. Nandula, J.A. Wehrmeyer, O. Jarrett, Jr., G.B. Northam, J.-Y. Chen. Detonations. Laser-Initiated Conical Detonation Wave for Supersonic Combustion -- a Review G.F. Carrier, F.E. Fendell, Mau-Song Chou. Recent Advances in Ram Accelerator Technology J.M. Powers. Detonation Waves and Propulsion J. Shepherd. Studies on Detonation Driven Hollow Projectiles P.A. Thibault, J.D. Penrose, A. Sulmistras, S.B. Murray, J.L.D.S. Labbe. Ignition and Structure. The Role of Mathematical Modeling in Combustion J. Buckmaster. Ignition and Flame Spread in Laminar Mixing Layers A. Linan. Numerical and Asymptotic Analysis of Ignition Processes C. Trevino, A. Linan. Unsteady Behavior. Steady and Unsteady Aspects of Detonation Initiation J.W. Dold, J.F. Clarke, M. Short. Weakly Nonlinear Dynamics ofNear-CJ Detonation Waves J.B. Bdzil, R. Klein. Some Fundamental Problems of Detonation Instabilities and Its Relation to Engine Operation J.H. Lee, Fan Zhang, R.S. Che. Godunov-Type Schemes Applied to Detonation Flows J.J. Quirk.

Detonation combustion is characterized by the high thermodynamic efficiency and fast heat release. Benefitting from these potential advantages, an oblique detonation wave (ODW) is introduced into the combustion chamber and oblique detonation engine (ODE) plays an important role in hypersonic air-breathing propulsion systems. Previous studies mainly focused on the initiation structures, standing features and wave systems of oblique detonation, but the global analysis of ODE propulsive performance is still absent at the macro-level. In this paper, the flow and combustion processes of an ODE are decomposed into four basic modules, named as inlet model, mixing model, combustion mode and nozzle model, respectively. We solve these four basic flow processes using theoretical methods and propose a systematically theoretical approach that can be used to predict the ODE propulsion performance. On the basis of previous ODW initiation structures and waves systems, four different combustion modes, i.e., over-driven ODW, Chapman-Jouguet ODW, over-driven normal detonation wave and oblique shock-induced constant-volume combustion, are chosen to describe the heat release processes of combustible mixture in the ODE combustor. The effects of different combustion modes on fuel specific impulse of the ODE are also analyzed. In addition, the influence mechanisms of inflow parameters, combustor parameters and intake-exhaust parameters on the thrust performance of ODE are also obtained, and the results show that the major factor of fuel specific impulse of an ODE consists mainly of the inflow Mach number and the expansion ratio of engine nozzle. Finally, combined with precious detonation research results, such as the standing features and initiation structures of oblique detonation in a confined space, the preliminary design direction of oblique detonation engine are proposed, which mainly involve some constrained conditions, such as geometrical constraints, inflow velocity limitations and stability ranges of a detonation wave in ODE combustor.

In the framework of the joint DLR-ONERA project JAPHAR on hypersonic air-breathing propulsion, supersonic combustion experiments were conducted in the configuration of an axial injection of hydrogen at Mach 2 in an air flow also at Mach 2. The structure of the reactive mixing layer between the fuel and the air has been investigated by means of Particle Image Velocimetry (PIV). PIV measurements are based on two successive images of tracer particles seeding the flow. The correlation between the two images yields the quasiinstantaneous two-dimensional velocity field. The effect of seeding particles on the velocity measurements has been studied and the results are compared to previous data obtained with single point Laser Doppler Velocimetry. The combustion regime in the different zones explored by the PIV technique was previously identified by wall pressure measurements and direct visualizations of the flame. The explored zones display the boundary layer, potential core of the hydrogen jet, reflection of compression and expansion waves. The decrease of the axial velocity of the hydrogen jet gives information on the rate of mixing.

Review of combustion stabilization for hypersonic airbreathing propulsion

An overview of some of the activities in hypersonic airbreathing aerodynamics and propulsion airframe integration is presented for the Space Vehicle Technology Institute. The Institute is a multi-university joint NASA-DoD program that was created as a center for research and education in future launch vehicle technologies. Perhaps more than any other type of flight vehicle, next generation space launchers will have to be analyzed as completely integrated aerodynamic-propulsion systems. Optimal aerodynamics will be vital to the development of efficient, engine-integrated launch vehicle forms, especially using airbreathing propulsion. To this end, inverse design approaches, design tradeoffs, and an understanding of relevant basic flow physics are all part of the Space Vehicle Technology Institute program. The relevance of these efforts to NASA activities is also described.

During last ten years, a large effort has been undertaken in Europe, and particularly in France, to improve knowledge on hypersonic airbreathing propulsion, acquire a first know-how for components design and develop needed technologies. MBDA France and ONERA brought major contributions to this effort by participating in different programs (PREPHA, WRR, JAPHAR, PROMETHEE).

In 1997, French ONERA and German DLR started a common research project to pursue, beyond the national programs PREPHA and Sager, the study of high speed airbreathing propulsion. The project focuses on the dual mode ramjet, the goal being, on one hand, to design and test at ground such an engine, and, on the other hand, to study its propulsive balance on a vehicle. Activities are organised around the blue-print of a possible experimental vehicle, that provides a solid context for the studies and a basis to synthesise the results through the thrust minus drag budget.

Hypersonic airbreathing propulsion - Flight test needs

This paper is aimed at development of an integrated approach based on analytical and computational aerothermodynamics for the special case of design of a 75% (low process-efficiency), hydrogen-fuelled, constant area combustor of a hypersonic airbreathing propulsion (HAP) system thereafter undertaking study of two types of HAP systems. The results of configurational aerothermodynamics implied that the most appropriate constant area configuration had a 30 degrees downstream wall-mounted fuel injector with a single acoustically stable cavity placed downstream of the fuel injection point. Moreover for identical flow inlet parameters and system configurations at lower levels of thermodynamic process efficiencies, the constant combustor-area (i.e. Scramjet 1) engine is superior in its performance to the constant combustor-pressure (i.e. Scramjet 2) engine for all values of fuel-air ratios.