Scramjet Flight Research Articles

Estimation of Aerodynamic Heating on Scramjet Inlets and Validation With Measurements

Abstract Aerodynamic heating levels on a typical inlet configuration of a scramjet engine are estimated using both standard design correlations and numerical flow simulations. The stagnation point heat flux is estimated using the Fay and Riddell formula. Aerodynamic heating over the inclined ramps is estimated using Van Driest method. Numerical flow simulations are carried out using a Reynolds averaged Navier–Stokes (RANS) solver coupled with energy equations and the Shear Stress Transport (SST) k–ω turbulence model. The aerodynamic heat flux estimates are validated with in-house measurements in a shock tunnel and for a scramjet flight experiment in the Mach number range 1.59 to 7.92. The emergence of a good agreement between them confirms the appropriateness of design correlations for heat flux estimation in scramjet inlets. The choice of simplification and appropriateness of design correlations to complex geometries demand critical assessment. Numerical flow simulations capture flow features and enable the identification of potential augmented heating zones, which will be critical for long-duration scramjet missions.

Effect of the reaction mechanism on the numerical prediction of the performance of a scramjet combustor at cruise flight 8 Mach number

A numerical investigation of the effect of the hydrogen-air reaction mechanism on the prediction of supersonic combustion was conducted. Different reaction mechanisms with different number of species and reactions were investigated at a Scramjet hypersonic flight trajectory. A generic three-dimensional supersonic combustor was investigated at flight 8 Mach number condition and at an altitude of 30 km. Numerical simulations were performed to determine the influence of the size of the reaction mechanism on the prediction of supersonic lifted flames, in a Scramjet flight trajectory were the combustor inflow conditions are close to the hydrogen-air self-ignition limit. Results showed significant influence of the number of species and reactions involved in the mechanism on the prediction of the ignition length. Shorter ignition delays were obtained with a reduced mechanism. Similarly, combustion efficiency was affected by the reduced mechanism dropping nearly 15%, compared to the reference scheme. On the other hand, no significant influence of the reaction mechanism was observed on the combustor pressure drop. Pressure losses nearly converge to the same values by the exit of the combustor. The present data suggest that the main observed differences are on the different prediction of the location where self-ignition occurred. We show how the ignition delay is affected by the omission of secondary species and steps reactions leading to shorter ignition delays or even attached flames which are not expected to occur at the investigated hypersonic flight condition.

Numerical Study on Combustor Flow-Path Design for a Scramjet Flight Experiment

Numerical Study on Combustor Flow-Path Design for a Scramjet Flight Experiment

Combustion efficiency and pressure loss balance for the supersonic combustor

The purpose of this paper is to investigate the effects of intake’s compression process of the scramjet on its flight performance. The hydrogen injection to the supersonic cross-flow is considered as the problem configuration. The finite volume solver is developed to simulate the compressible reacting turbulent flow using the proper reaction mechanism as the finite rate chemistry. The combustion efficiency and the drag force are the most important parameters on the scramjet flight performance, and finding the design point to balance the higher combustion efficiency and the lower minimum drag, which depends on the total pressure loss, can be used to optimize the supersonic combustors. The performance of the supersonic intake is considered here using some oblique shock waves with equal flow-deflection angles to compute the combustor’s inlet condition. The variation of combustion efficiency and total pressure loss is presented for different combustor’s inlet conditions. The results are presented for the constant jet to inlet pressure ratios and also for the constant equivalence ratios, in which the last one is much appropriate and utilized to find the optimum design point of the intake and the combustor, for assumed flight condition.

CFD Analysis of Early Diaphragm Removal in Expansion Tubes

Impulse facilities generate transient high-velocity gas flows for ground testing in aerodynamics. The expansion tube is the facility type with the highest performance capability in terms of total flow enthalpy and total pressure, and is particularly useful for studies of atmospheric re-entry, chemical kinetics, scramjet flight and supersonic combustion. In operation, a thin film diaphragm that initially partitions two tube sections is required to rupture under the force of a shock-wave. Fragmented pieces are accelerated with the flow and can damage test models and instrumentation, and the rupture process itself affects flow properties. It has been proposed to replace the diaphragm with a fast open valve that clears the tube prior to shock arrival. This paper investigates the effect of early valve opening on the test flow.An inviscid, axisymmetric model was created using Eilmer3, a compressible CFD solver developed at the University of Queensland. Early valve opening was simulated at varying times for instantaneous diaphragm removal. The result was the formation of a secondary shock and expansion wave. The primary shockwave reaches the test section at a higher velocity due to passing through the expansion wave, creating a faster, higher-pressure test flow, but also higher temperature, leading to substantially reduced Mach number. The interactions with the secondary waves were found to cause unsteadiness in test flow properties.

Characteristics of Cascaded Fuel Injectors Within an Accelerating Scramjet Combustor

Performances of cascaded hydrogen injectors within an accelerating scramjet combustor were characterized via a three-dimensional numerical study. Two streamwise-aligned jets are employed, with the upstream injector half the diameter of the rear jet. Each are inclined at 45 deg to the freestream and achieve a jet-to-freestream momentum ratio of unity. Performance was evaluated over a range of freestream Mach numbers, modeling combustor entrance conditions on an accelerating access-to-space scramjet trajectory. Distance between injectors was varied to estimate the optimum jet-to-jet spacing to achieve robust performance across the Mach number range. It is shown that the downstream injector benefits jointly from shielding effects induced by the smaller upstream injector, as well as its increased diameter. Injector spacings greater than two total jet diameters displayed improved absolute jet penetration, spanwise spread, and entrainment rates over an equivalent single injector across all Mach numbers. Jet bow shock pressure recovery was improved at spacings above . Unique optimal injector spacings existed for each performance metric, at each freestream Mach number examined. Although no universal optimum spacing was determined, spacings between 4 and provided substantial performance enhancements over single injectors, throughout the accelerating scramjet flight conditions.

Skin Friction Sensor Development, Validation, and Application for High-Speed, High-Enthalpy Flow Conditions

This investigation concerns design and testing of skin friction sensors for high-speed, high-enthalpy conditions, such as re-entry vehicles, scramjets, jet engines, material testing, etc. Understanding these flows requires numerical and analytical modeling and reliable instrumentation. The sensor design is a direct measuring, non-nulling, cantilever beam arrangement. A multistep program tested the sensor through various facilities, culminating in simulated scramjet flight conditions. The first phase concerned calibrations for static, dynamic, pressure, thermal, and vibratory responses and electromagnetic interference. This phase concluded under Mach 3 conditions: stagnation temperatures and pressures of 300–655 K and 750–1000 kPa. Wall shear was measured as 132–248 Pa for skin friction coefficients 0.0010–0.0014. The measurements demonstrated agreement with flat-plate analytical estimations, numerical predictions, and historical test results. Total uncertainty was at 95% confidence. The second phase involved testing at Air Force scramjet facilities. Mach 2.0–2.2 flows were studied at stagnation pressures and temperatures (172–995 kPa and 294–1000 K). The sensors determined wall shear measurements (94–600 Pa) for skin friction coefficients (0.0016–0.0028). The sensors clearly indicated transient flow of a scramjet operability cycle with shock train movement and separation zones. The sensor demonstrated full functionality under sustained high-enthalpy flow conditions upward of .

The large-amplitude combustion oscillation in a single-side expansion scramjet combustor

The combustion oscillation in scramjet combustor is believed not existing and ignored for a long time. Compared with the flame pulsation, the large-amplitude combustion oscillation in scramjet combustor is indeed unfamiliar and difficult to be observed. In this study, the specifically designed experiments are carried out to investigate this unusual phenomenon in a single-side expansion scramjet combustor. The entrance parameter of combustor corresponds to scramjet flight Mach number 4.0 with a total temperature of 947K. The obtained results show that the large-amplitude combustion oscillation can exist in scramjet combustor, which is not occasional and can be reproduced. Under the given conditions of this study, moreover, the large-amplitude combustion oscillation is regular and periodic, whose principal frequency is about 126Hz. The proceeding of the combustion oscillation is accompanied by the transformation of the flame-holding pattern and combustion mode transition between scramjet mode combustion and ramjet mode combustion.

Aerothermal–Structural Analysis of a Rocket-Launched Mach 8 Scramjet Experiment: Ascent

This paper reports on the methodology and results of a weak-coupled aerothermal–structural analysis on the ascent phase of the SCRAMSPACE Mach 8 scramjet flight experiment. This vehicle was essentially unshrouded during the flight trajectory, relying on the thin, 5 mm thick aluminum external shell of the payload to maintain structural integrity and protect the flight experiment. As such, understanding the thermal–structural response of the vehicle was imperative to mission success. Using two- and three-dimensional models, an iterative procedure was employed to compute the flowfield, convective heating, wall temperatures and structural coupling at flight times covering both peak heating and peak surface temperature. Accounting for such coupling resulted in a 150 K reduction in wall temperature compared to the more conservative cold wall assumption. Despite this, peak temperatures remained of the order of 550 K. Further, thermally induced stresses within these regions were in excess of four times the material failure limits. Irreversible material failure during ascent was therefore concluded likely to occur on the external shell. Two alternate materials, steel 1006 and copper, were therefore assessed with the results indicating that steel sections on the external shell resulted in the best thermal–structural response of the payload.

Mach Hydrocarbon-Fueled Scramjet Flight Experiment: The HIFiRE Flight 2 Project

The Hypersonic International Flight Research Experimentation (HIFiRE) program is a collaborative international effort designed to study basic hypersonic phenomena through flight experimentation. On 1 May 2012, the HIFiRE Flight 2 project successfully flew a hydrocarbon-fueled, Mach 8 scramjet experiment and demonstrated the ability to fly an accelerating, constant dynamic pressure trajectory consistent with a scramjet-powered flight vehicle using unguided sounding rocket techniques. The project goals of capturing high-quality flight data from a research scramjet operating through dual-to-scram mode transition up to and beyond Mach 8 were achieved. HIFiRE Flight 2 is unique in its contribution to scramjet research, providing a reference dataset to the hydrocarbon scramjet community while maturing a novel and reduced-cost strategy for performing similar future experiments. This paper describes the programmatic approach, the overall scramjet flight test experiment mission objectives, and the flight test strategy. It also includes an overview of launch system and payload hardware, and brief discussions of flight activities, flight data, and experimental results. Preliminary results from flight indicate a fully successful mission and experiment. Based on evaluations of data collected during the project, including ground test, flight test, and multiple analysis activities, all project-level scientific objectives have been successfully achieved.

Investigations on the influence of the in-stream pylon and strut on the performance of a scramjet combustor.

The influence of the in-stream pylon and strut on the performance of scramjet combustor was experimentally and numerically investigated. The experiments were conducted with a direct-connect supersonic model combustor equipped with multiple cavities. The entrance parameter of combustor corresponds to scramjet flight Mach number 4.0 with a total temperature of 947 K. The research results show that, compared with the scramjet combustor without pylon and strut, the wall pressure and the thrust of the scramjet increase due to the improvement of mixing and combustion effect due to the pylon and strut. The total pressure loss caused by the strut is considerable whereas pylon influence is slight.

Multistability and Loops-Coupled Hysteresis: Flight-Test Analysis on Error Detection of Inlet Start/Unstart

DOI: 10.2514/1.B34349Anerrordetectionofinletstart/unstarthasbeenreportedasasignificantflightaccidentintheCentralInstituteofAviation Motors/NASA Mach 6.5 scramjet flight test. It is believed that nonlinear effects that give rise to thecatastropheandhysteresisbehaviorsinthetransitionsofinletstart/unstartmaybetheessentialfactorsthatresultinthe flight accident. Starting from nonlinear interaction effects and dynamical system theories, evidence ofmultistabilityandloops-coupledhysteresisphenomenonexhibitedinthehypersonicinletisreported.Multistabilityof a dynamic system consists of the ability to express multiple stable states and to switch between these states inresponse to some external input. The loops-coupled hysteresis phenomenon, which appears as a generic feature ofmultistability systems, is characterized by the existence of at least one hysteresis loop that involves more than twostablestates.Thepresentstudyshowstheloops-coupledhysteresisphenomenonhasoccurredintheflighttest,andaneglect of the multistability and loops-coupled hysteresis phenomenon has given rise to the error detection of inletstart/unstart in the Central Institute of Aviation Motors/NASA Mach 6.5 scramjet flight test.

Dispersion Reduction for a Sounding Rocket Scramjet Flight Experiment

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

Dispersion Reduction for a Sounding Rocket Scramjet Flight Experiment

A low cost method for reducing the dispersion in the trajectory of an unguided, spinstabilized, sounding rocket is developed and presented. The method is particularly suited to scramjet flight experimentation because the approach increases the likelihood of meeting Mach number and dynamic pressure objectives. The paper discusses the design and model of the scramjet payload, and two-stage launch vehicle, and the nominal trajectory, as well as a Monte Carlo analysis to quantify the likelihood of a successful scramjet test. Using the results of this analysis, a method is presented for reducing the dispersion in freestream conditions during the scramjet test window. The dispersion reduction is accomplished by modifying the time delay between the burnout of the first stage booster and the ignition of the second stage based on the vehicle state measured during the interstage coast. This method increases the likelihood of a successful test from 71% to 99% without adversely affecting range safety. Since the design and implementation of a vehicle guided control system is not required, this method is relatively inexpensive, making its use highly desirable for low cost scramjet flight experimentation.

Methodology of a combined ground based testing and numerical modelling analysis of supersonic combustion flow paths

In the framework of the European Commission co-funded LAPCAT (Long-Term Advanced Propulsion Concepts and Technologies) project, the methodology of a combined ground-based testing and numerical modelling analysis of supersonic combustion flow paths was established. The approach is based on free jet testing of complete supersonic combustion ramjet (scramjet) configurations consisting of intake, combustor and nozzle in the High Enthalpy Shock Tunnel Gottingen (HEG) of the German Aerospace Center (DLR) and computational fluid dynamics studies utilising the DLR TAU code. The capability of the established methodology is demonstrated by applying it to the flow path of the generic HyShot II scramjet flight experiment configuration.

Flight Data Analysis of the HyShot 2 Scramjet Flight Experiment

The development of scramjet propulsion for alternative launch and payload delivery capabilities has been composed largely of ground experiments for the last 40 years. With the goal of validating the use of short duration ground test facilities, a ballistic reentry vehicle experiment called HyShot was devised to achieve supersonic combustion in flight above Mach 7.5. It consisted of a double wedge intake and two back-to-back constant area combustors; one supplied with hydrogen fuel at an equivalence ratio of 0.34 and the other unfueled. Of the two flights conducted, HyShot 1 failed to reach the desired altitude due to booster failure, whereas HyShot 2 successfully accomplished both the desired trajectory and satisfactory scramjet operation. Postflight data analysis of HyShot 2 confirmed the presence of supersonic combustion during the approximately 3 s test window at altitudes between 35 and 29 km. Reasonable correlation between flight and some preflight shock tunnel tests was observed.

Experimental Investigation of Hydrocarbon-fuel Ignition in Scramjet Combustor

The direct-connected supersonic combust or experiment is finished for kerosene fuel ignition in H2/O2 preheated impulse facility. The entrance parameter of combustor corresponds to scram jet flight Mach number 3.5. Kerosene ignition is realized by using hydrogen as pilot flame. Wall pressure distributions of combustion are measured and flame photographs of ultraviolet ray are got. Experiment indicates that it is very difficult for kerosene fuel to realize self-ignition at low entrance temperature (below 900K) in supersonic combustor. Hydrogen pilot flame is one of the efficient methods for realizing kerosene ignition.

Numerical analysis of supersonic combustion ramjet with upstream fuel injection

AbstractThis paper describes possible fuel injection scheme for airbreathing engines that use hydrocarbon fuels. The basic idea is to inject fuel at the spike tip of the supersonic inlet to achieve mixing and combustion efficiency with a limited length combustion chamber. A numerical code, able to solve the full Navier–Stokes equations in turbulent and reacting flows, is employed to obtain numerical simulations of the thermo‐fluidynamic fields at different scramjet flight conditions, at Mach numbers of M=6.5 and 8. The feasibility of the idea of the upstream injection is checked for a simple axisymmetric configuration and relatively small size. The results are discussed in connection with the potential benefits deriving from the use of new ultra high temperature ceramics (UHTC). Copyright © 2003 John Wiley & Sons, Ltd.

Aerodynamic Database Development for the Hyper-X Airframe-Integrated Scramjet Propulsion Experiments

An overview of the activities associated with the aerodynamic database that is being developed in support of NASA's Hyper-X scramjet flight experiments is provided. Three flight tests are planned as part of the Hyper-X program. Each will utilize a small, non-recoverable research vehicle with an airframe-integrated scramjet propulsion engine. The research vehicles will be individually rocket boosted to the scramjet engine test points at Mach 7 and 10. The research vehicles will then separate from the first stage booster vehicle and the scramjet engine test will be conducted before the terminal decent phase of the flight. An overview is provided of the activities associated with the development of the Hyper-X aerodynamic database, including wind-tunnel test activities and parallel computational fluid dynamics analysis efforts for all phases of the Hyper-X flight tests. A brief summary of the Hyper-X research vehicle aerodynamic characteristics is provided, including the direct and indirect effects of the airframe-integrated scramjet propulsion system operation on the basic airframe stability and control characteristics. Brief comments on the planned postflight data analysis efforts are also included.

Review of shock-induced supersonic combustion research and hypersonic applications

This article focuses on research in supersonic and kinetics in high-speed flow between 1959-1968, and the application of the experimental results to hypersonic propulsion. The analysis discusses both advantages and problems for premixing the fuel and employing as an ignition method for a scramjet flying at a high Mach number. The experimental tests are discussed, including implications to the chemical kinetics of the high-velocity process. The conditions were confined to relatively low pressure, less than 2 atm (200 kPa). The results were considered to be mainly applicable for high-altitude scramjet flight, at low static pressure, where chemical reaction distances will be long. At these lower pressures, shock-induced combustion may be the predominant effect in a scramjet application, and it has some advantages that are discussed. The relation between and is also discussed. In addition, an attempt is made to resolve the conflicting experimental data published in the 1960s relating to standing detonation waves and combustion.