Scramjet Test Research Articles

Large eddy simulation of a dual-mode scramjet with direct injection: Flame dynamics under flow separation induced by heat-release

The combustion process in the scramjet engines generally takes place in unsteady conditions. The unsteadiness of the flame can be present locally or globally in supersonic combustion, for instance, reactant mixing, ignition, flame stabilization, and blow-off. Thus, unsteady supersonic combustion is an urgent problem that needs to be solved for real scramjet engine applications. This work helps to advance the knowledge of dual-mode scramjet combustion operating with different equivalence ratios by conducting quantitative analyses that characterize flame dynamics under flow separation. A hydrogen-fueled direct-injection scheme was computed using blastFoam code via large-eddy simulation of a scramjet test rig at a stagnation temperature of 950 K and a pressure of 0.82 MPa. The computational results were in good agreement with the experimental data, and two typical flame structures (cavity shear-layer and cavity assisted jet-wake stabilized flame) corresponding to the scram/ram operating modes were reproduced. The local flame structure and flow regime were investigated using a modified flame index and filter functions. A particular focus was set on the continuous flame flashback during mode transition, which could be divided into three stages involving jet/shear-layer interaction, unsteady heat release, and the establishment of a large-area separation zone, with the decay of the cavity recirculation zone. Additionally, the unsteady jet-wake ram combustion presented three oscillation modes, giving rise to intermittent flameout of varying levels. The coherent-structures motion caused by the combined effects of low-speed separated flow and the fuel-jet was the main cause for this kind of unsteady combustion.

Non-Reacting Examination on Transverse Injection Upstream of Aft Wall Angled Cavities in an Axisymmetric Supersonic Flow

Experiments on an air-fueled non-reacting scramjet test facility with four distinct aft wall angled cavity designs were performed in a Mach 1.8 flow field to explore the effects of transverse fuel injection system on scramjet combustors. Axisymmetric combustor cavity models were featured with two consecutive angles being inclined towards the downstream flow direction at the rear end. The wall mounted injector was located upstream from the cavity at a distance of 10 mm. Mixing performance within the various rear wall angled cavities were evaluated and compared with No-injection case for different injection pressures. Results revealed that the flow within the combustor was more uniform mixing compared to No-cavity configuration and was greatly influenced by injection pressures. Increase in injection pressures were characterized by enhanced mixing, greater stagnation pressure loss values and unstable flow.

Expansion tube nozzle design using a parallel simplex algorithm

A new hypersonic Mach 12 nozzle has been designed and manufactured for a large-scale expansion tube. The nozzle design goals were to produce a Mach 12 flow, with a core flow diameter of at least 300 mm and a maximum exit flow angle non-uniformity of less than $${2}^{\circ }$$. The nozzle has been designed by coupling a RANS CFD solver with a parallel simplex algorithm to solve the computationally expensive optimisation problem. Novel aspects of this analysis are that it addresses nozzle optimisation specifically for scramjet test flows, characterised by high pressures and thick boundary layers. A new MPI implementation of a block-marching technique is used to solve the flowfield, which is optimised with a parallel Nelder–Mead algorithm. The validity of the objective functions is discussed through a robust bootstrapping analysis, and off-design performance of the nozzle is also characterised. The analysis demonstrates that the optimised contour achieves the design objectives and has excellent off-design performance. Initial commissioning experiments confirmed the results of the numerical analysis. Despite the final operating conditions being slightly off-design, the nozzle was able to produce an experimental core flow exceeding the numerical predictions. Indeed, the manufactured nozzle, 2.8 m long, with an exit diameter of 573 mm, has been shown experimentally to produce a core flow size of 360 mm, enabling full-scale Mach 12 scramjet experiments.

Characterisation of the eddy dissipation model for the analysis of hydrogen-fuelled scramjets

ABSTRACTThe eddy dissipation model (EDM) is analysed with respect to the ability to address the turbulence–combustion interaction process inside hydrogen-fuelled scramjet engines designed to operate at high Mach numbers (≈7–12). The aim is to identify the most appropriate strategy for the use of the model and the calibration of the modelling constants for future design purposes. To this end, three hydrogen-fuelled experimental scramjet configurations with different fuel injection approaches are studied numerically. The first case consists of parallel fuel injection and it is shown that relying on estimates of ignition delay from a 1D kinetics program can greatly improve the effectiveness of the EDM. This was achieved through a proposed zonal approach. The second case considers fuel injection behind a strut. Here the EDM predicts two reacting layers along the domain which is in agreement with experimental temperature profiles close to the point of injection but not the case any more at the downstream end of the test section. The first two scramjet test cases demonstrated that the kinetic limit, which can be applied to the EDM, does not improve the predictions in comparison to experimental data. The last case considered a transverse injection of hydrogen and the EDM approach provided overall good agreement with experimental pressure traces except in the vicinity of the injection location. The EDM appears to be a suitable tool for scramjet combustor analysis incorporating different fuel injection mechanisms with hydrogen. More specifically, the considered test cases demonstrate that the model provides reasonable predictions of pressure, velocity, temperature and composition.

Development of high temperature material for cellular ceramic regenerative storage heater

Ceramic Regenerative Storage Heater (RSH) is the preferable solution to provide clean hot air in Hypersonic Wind Tunnel (HWT) and Scramjet Test facilities to prevent the liquefaction of air during expansion in the high Mach number nozzles. Selection of the heater storage material for such operational requirements is a challenging task. As per the theoretical background; alumina, zirconia or mullite compositions have the capability to withstand at such high operating conditions. In the current study, a series of experiments were carried out to estimate the thermal characteristics of the material. This paper highlights the simplified approach by combining theoretical and experimental methods for economic (reduced time and cost) selection process of high temperature material. The desired material must possess good thermal shock resistance and low dust generation characteristics. Al2O3 and SiO2 based matrix composition fused in a transcrystalline phase exhibits significant improvement on the overall material integrity. Several Al2O3 and SiO2 matrix compositions were examined and the final matrix composite ratio of 90% of Al2O3 and 10% of SiO2 is considered as the desired candidate material. The ceramic matrix was sintered at a temperature of 2123 K, under the specific pressure of 1400 kg/cm2 for a period of 2 h, at a heating rate of 3.5 °C/min. The results of the thermal shock tests on various samples are presented. To improve the anti-dust characteristics, the final samples were coated with fine particles (5 µm) of Al2O3 and SiO2 matrix. Based on the experimental evaluation, Al2O3 and SiO2 fused matrix composition found to be meeting all the requirements of high temperature storage material.

Experimental and numerical investigations of freejet and direct-connect dual-mode scramjet

To understand scramjet operations, comparisons between direct-connect and freejet dual-mode scramjet tests were investigated in this paper via experiment and numerical simulation. Being in accord with the condition of flight Mach number 6.0, the inlet flow conditions of the isolator were achieved with Mach number 3.0, total temperature 1500 K and total pressure 2.1 MPa. Based on these parameters, two tests with equivalence ratio 0.6–600 and 1.0–604 were carried out. In the experiment, pressure at the centerline of the combustor down wall was measured. Simulation including the Mach number distribution at the combustion chamber, the shock wave structures in the isolator, and heat release of the combustor were proceeded and compared. Experimental results show that the measured pressures match the numerically calculated outcomes. Simulation results show that when the combustion case is 1.0–604, the combustor operates in the ram-mode for the freejet test while it is in the scram-mode for the direct-connect test. When the case is 0.6–600, the combustor in both freejet and direct-connect tests operates in the scram-mode. In addition, the specific thrust in the freeject test presents a prominent difference to that in the direct-connect test when the engine operates in the same combustion mode. For case 1.0–604, the thrust difference is 4.38% and for case 0.6–600, the discrepancy is 8.19%. Besides, influence of the heat release on the thrust performance was analyzed for the two tests.

Selection of High Temperature Heat Storage Materials for Cored Brick Heaters of Hypersonic Wind Tunnel and Scramjet Facilities

Hypersonic Wind Tunnel (HWT) and Scramjet test facilities use cored brick heaters to generate high temperature pure air. The selection of right material for storage medium of cored brick heaters is a challenging task. Many tests are required to be carried out to finalise the suitable material for cored brick heaters. Some of these qualification tests require numerous number of samples of different sizes. This evaluation cycle demand years of time and huge financial resources. It is very important to know which test to be carried out first so that the right material is chosen at the initial stages. The proposed paper is aimed at development of a sequential flow which can potentially reduce the time and financial requirements, to evaluate the candidate material for cored brick heaters. This paper is based on the experience of material selection followed for cored brick heater of HWT project of DRDL. In the present study, the initial challenges involved in identifying the material for cored brick heater are discussed along with factors to be considered for efficient storage and release of energy. Experimental results of characterization of the storage materials by considering critical parameters such as compressive strength, thermal shock resistance and abradability index are presented. Based on the experimentations and observations, sequence of testing required to be followed to finalize the candidate material, is evolved. A detailed methodology for the selection of high temperature heat storage material is presented.

Scramjet test flow reconstruction for a large-scale expansion tube, Part 1: quasi-one-dimensional modelling

Large-scale free-piston driven expansion tubes have uniquely high total pressure capabilities which make them an important resource for development of access-to-space scramjet engine technology. However, many aspects of their operation are complex, and their test flows are fundamentally unsteady and difficult to measure. While computational fluid dynamics methods provide an important tool for quantifying these flows, these calculations become very expensive with increasing facility size and therefore have to be carefully constructed to ensure sufficient accuracy is achieved within feasible computational times. This study examines modelling strategies for a Mach 10 scramjet test condition developed for The University of Queensland’s X3 facility. The present paper outlines the challenges associated with test flow reconstruction, describes the experimental set-up for the X3 experiments, and then details the development of an experimentally tuned quasi-one-dimensional CFD model of the full facility. The 1-D model, which accurately captures longitudinal wave processes, is used to calculate the transient flow history in the shock tube. This becomes the inflow to a higher-fidelity 2-D axisymmetric simulation of the downstream facility, detailed in the Part 2 companion paper, leading to a validated, fully defined nozzle exit test flow.

Scramjet test flow reconstruction for a large-scale expansion tube, Part 2: axisymmetric CFD analysis

This paper presents the second part of a study aiming to accurately characterise a Mach 10 scramjet test flow generated using a large free-piston-driven expansion tube. Part 1 described the experimental set-up, the quasi-one-dimensional simulation of the full facility, and the hybrid analysis technique used to compute the nozzle exit test flow properties. The second stage of the hybrid analysis applies the computed 1-D shock tube flow history as an inflow to a high-fidelity two-dimensional-axisymmetric analysis of the acceleration tube. The acceleration tube exit flow history is then applied as an inflow to a further refined axisymmetric nozzle model, providing the final nozzle exit test flow properties and thereby completing the analysis. This paper presents the results of the axisymmetric analyses. These simulations are shown to closely reproduce experimentally measured shock speeds and acceleration tube static pressure histories, as well as nozzle centreline static and impact pressure histories. The hybrid scheme less successfully predicts the diameter of the core test flow; however, this property is readily measured through experimental pitot surveys. In combination, the full test flow history can be accurately determined.

X3 expansion tube driver gas spectroscopy and temperature measurements

The University of Queensland’s X3 facility is a large, free-piston driven expansion tube used for super-orbital and high Mach number scramjet aerothermodynamic studies. During recent development of new scramjet test flow conditions, experimentally measured shock speeds were found to be significantly lower than that predicted by initial driver performance calculations. These calculations were based on ideal, isentropic compression of the driver gas and indicated that loss mechanisms, not accounted for in the preliminary analysis, were significant. The critical determinant of shock speed is peak driver gas sound speed, which for a given gas composition depends on the peak driver gas temperature. This temperature may be inaccurately estimated if an incorrect fill temperature is assumed, or if heat losses during driver gas compression are significant but not accounted for. For this study, the ideal predicted peak temperature was 3750 K, without accounting for losses. However, a much lower driver temperature of 2400 K is suggested based on measured experimental shock speeds. This study aimed to measure initial and peak driver gas temperatures for a representative X3 operating condition. Examination of the transient temperatures of the driver gas and compression tube steel wall during the initial fill process showed that once the filling process was complete, the steady-state driver gas temperature closely matched the tube wall temperature. Therefore, while assuming the gas is initially at the ambient laboratory temperature is not a significant source of error, it can be entirely mitigated by simply monitoring tube wall temperature. Optical emission spectroscopy was used to determine the driver gas spectra after diaphragm rupture; the driver gas emission spectrum exhibited a significant continuum radiation component, with prominent spectral lines attributed to contamination of the gas. A graybody approximation of the continuum suggested a peak driver gas temperature of 3200 K; uncertainty associated with the blackbody curve fit is ±100 K. However, work is required to quantify additional sources of uncertainty due to the graybody assumption and the presence of contaminant particles in the driver gas; these are potentially significant. The estimate of the driver gas temperature suggests that driver heat losses are not the dominant contributor to the lower-than-expected shock speeds for X3. Since both the driver temperature and pressure have been measured, investigation of total pressure losses during driver gas expansion across the diaphragm and driver-to-driven tube area change (currently not accounted for) is recommended for future studies as the likely mechanism for the observed performance gap.

Thermo-mechanical characterization of laser textured LaMgAl11O19/YSZ functionally graded thermal barrier coating

The ablation resistance and adhesion strength of laser textured and as-sprayed functionally graded LaMgAl11O19/YSZ based thermal barrier coating (FG-TBC) was studied and compared with a duplex coating of LaMgAl11O19 and YSZ (DC-TBC). The thermal barrier coatings were sprayed over Hastelloy C263 using atmospheric plasma spray process (APS). The laser texturing was performed using a picosecond pulsed laser. The width-to-depth ratio and groove spacing were varied to analyse and correlate the significance of groove parameters in providing strain tolerance under extreme thermal loading conditions. The ablation behaviour of the coatings was investigated using a scramjet test rig. The shear force induced through the ablation load spalled the as-sprayed DC-TBC. The thermal stress mismatch instigated through the difference in coefficient of thermal expansion (CTE) between discrete YSZ and LaMgAl11O19 layers led to the failure of the as-sprayed DC-TBC. The as-sprayed FG-TBC exhibited an improved ablation resistance. The varying thermo-mechanical properties along the ceramic layer thickness in as-sprayed FG-TBC minimise the concentration of thermal stress across the coating and imparts better thermal stability. The improved mechanical interlocking of molten splats during re-solidification of five discrete layers of ceramic composites inhibits the concentration of stress around the weak links. Compared to the textured DC-TBCs, the textured FG-TBCs exhibited minimal level of distortion and better resistance to ablation. The textured surfaces restrict the propagation of crack across the top surface and also provide improved strain tolerance. The graded structure minimises the stress built up and improves the coating resistance compared to textured DC-TBCs. The adhesion strength of all test surfaces was analysed following the standard test procedure. The textured surfaces having higher surface roughness (in terms of surface area) exhibited improved adhesion strength. The groove parameters were observed to have significant influence in improving the ablation resistance and adhesion strength of the coatings.

Numerical analysis on the heat transfer of three types of nozzles for the hypersonic long-run wind tunnel

The hypersonic long-run scramjet test tunnel is one of the key ground facilities for the studies of ramjet/scramjet and hypersonic thermal management. Due to the significantly large heat loading, the nozzle of the tunnel facility demands effective cooling protection. In this work, the two-dimensional, three-dimensional and axisymmetric Mach 6.5 nozzles at an inlet total temperature of 1840 K and a total pressure of 6.4 MPa were studied with main focuses on the properties of aerodynamic heating of nozzles. The present work aims to provide insights into the design of an effective cooling system for the nozzle and other components of the hypersonic long-run wind tunnel.

충격파 터널을 이용한 고속추진기관 시험 연구

충격파 터널은 초음속 및 극초음속 비행조건의 유동장을 지상에서 가장 유사하게 모사할 수 있는 시험설비로서, 거의 반세기 동안 전 세계적으로 유용하게 운용되고 있다. 국내의 극초음속 비행체에 대한 많은 관심의 결과로 충격파 시험장치에 대해 주목하고 있으며, 최근에 KAIST에서는 중간 크기의 충격파 터널을 갖추게 되었다. 본 논문은 개발된 충격파 터널의 성능과 최근 수행된 모델 스크램제트 시험자료를 제시하고 있다. 본 연구에서 모사된 자유흐름 유동조건은 마하수 5.7 비행속도에 해당하며, 이러한 비행조건에서 스크램제트의 초음속 연소기내에 유동 환경을 모사하였다. Shock tunnels are known to be capable of simulating flow-field environments of supersonic and hypersonic flights. They have been operated successfully world-wide for almost half a century. As a consequence of the strong interest in hypersonic vehicles in Korea, attention has been given on this type of facility and so an intermediate-sized shock tunnel has lately been built at KAIST. In the light of this, this paper presents our tunnel performance and some of the model scramjet test data. The freestream flow used in this work replicates a supersonic combustor environment for a Mach 5.7 flight speed.

Hypersonic Scramjet Testing via Diode Laser Absorption in a Reflected Shock Tunnel

A wavelength-multiplexed two-color tunable diode laser absorption spectroscopy sensor probing transitions near was developed to measure temperature and column density simultaneously across three lines of sight in a ground-based model scramjet combustor. High-enthalpy scramjet conditions equivalent to Mach 10 flight were generated with a reflected shock tunnel. The sensor hardware development and optical engineering to overcome noisy combustor conditions, including short test time, beam steering, and mechanical vibration, are discussed. A new scanned-wavelength-modulation spectroscopy technique was used to acquire the complete spectral lineshape while maintaining the high signal-to-noise ratio characteristic of wavelength-modulation spectroscopy measurements. Two combusting flow experiments were conducted, and the results compared with steady-state computational fluid dynamics calculations, with both the simulations and the measurements finding formation of from combustion during the test time. This work represents the first use of wavelength-modulation spectroscopy in a reflected shock tunnel and the first use of scanned-wavelength-modulation spectroscopy in a scramjet combustor.

Thermal Radiative Simulations and Measurements of a Scramjet Test Rig

Thermal radiation is a poorly understood process in scramjet engines, but it may play a significant role in the flow and wall heating of the combustion chamber. However, current simulation methods for predicting the thermal radiation in a flight scramjet combustion chamber have yet to be validated with experiments. An experimental measurement apparatus is placed at the rear exit of the HIFiRE 2 direct-connect rig at NASA’s Langley Research Center. An array of photodetectors gather emission in the infrared along several lines of sight across the combustor exit. The fields of view are simulated using a ray-tracing-method program that postprocesses a computational fluid dynamics flowfield simulation of the test rig. The ray-tracing program employs a simplified two-point correlated- spectral model with spectral model error bars. The predictions show an overlap in sensor and experimental predictions for 13 of 16 photodetectors.

Dual-Pump Coherent Anti-Stokes Raman Spectroscopy Measurements in a Dual-Mode Scramjet

In this paper, the authors describe dual-pump coherent anti-Stokes Raman spectroscopy (CARS) measurements of mixing and combustion in a direct-connect scramjet combustor operating at equivalent flight Mach typical of the ramjet–scramjet transition, in the scram mode. Measurements were performed in the University of Virginia’s scramjet test facility in which the air is heated by electrical resistance heaters. The CARS technique is used to acquire temporally and spatially resolved measurements of temperature and species mole fraction. Measurements were at four planes: one upstream of an fuel injector and three downstream. Contour plots of mean flow and standard deviation statistics are presented for cases with and without reaction of the fuel. The vibrational temperature at the exit of the facility nozzle, and in the freestream of the scramjet combustor, is elevated compared with the rotational temperature; the air– vibrational temperature is the same as the facility stagnation temperature. There are spatial nonuniformities of temperature exiting the heater. The mixing of the fuel jet from the single ramp wall injector and the growth of the plume downstream is shown. The flame is attached at the injector and propagates from the wall adjacent to the fuel plume, around the periphery of the plume, before engulfing the whole plume. Computational fluid dynamics modeling shows that the flow can be predicted well using hybrid large-eddy simulation/Reynolds-averaged Navier–Stokes methods with relatively simple subgrid models, provided facility effects are properly accounted for.

Production of High-Mach-Number Scramjet Flow Conditions in an Expansion Tube

The expansion tube is the only type of ground test facility currently able to simulate high-Mach-number scramjet test flows. These access-to-space flow conditions are characterized by total pressures of the order of gigapascals. The University of Queensland’s X2 expansion tube facility has recently been used to generate scramjet flow conditions between Mach 10 and 14, relevant to a 96 kPa dynamic pressure ascent trajectory, with total pressures up to 10 GPa. These experiments have matched flight enthalpy and Mach number while achieving test flow static pressures considerably higher than the true flight values. This demonstrates that Reynolds number can also be matched, even for subscale models, thereby addressing one of the major challenges for ground testing of high-Mach-number scramjet design concepts. The routine operation of expansion tubes at high flight enthalpies has already been established; but for operation at lower enthalpies (relevant to scramjet testing), the wave processes within these machines become more tightly coupled and difficult to tune. This paper details the combined analytical and numerical processes used to develop new flow conditions in the X2 facility. Experimental results are presented, and axisymmetric computational fluid dynamics analysis is used to fully characterize the test flow properties.

Comparison of hierarchical and non-hierarchical error indicators for adaptive mesh refinement for the Euler equations

This article deals with two automatic mesh refinement methods for the Euler equations discretized by the discontinuous Galerkin finite element method of lowest order. The main purpose is the comparison of an hierarchical error indicator with a non-hierarchical one. The underlying mesh refinements are based on quadrilateral meshes with hanging nodes. The hierarchical indicator is based on the comparison of the solutions of two consecutive globally refined meshes. It therefore includes information of the nonlinear problem on the fine mesh. In contrast to this, the non-hierarchical indicator employs only information on the given mesh in form of jumps of the solution over internal edges. As often done in practice, in both cases, we chose a certain physical quantity such as the density or the Mach number as a sensor variable to which the indicator is applied. Three different standard test problems are considered: •A scramjet test case representing an internal supersonic flow, •A cylinder test case representing an external supersonic flow, •A Naca wing case representing an external transonic flow.

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.

Measurement of Water Vapor Levels for Investigating Vitiation Effects on Scramjet Performance

C OMBUSTION heated wind-tunnel facilities are required for many different aspects of ground-based testing of hypersonic airbreathing propulsion systems. This is particularly true for largescale scramjet testing in facilities with run times on the order of tens of seconds to minutes at equivalent flight Mach numbers above 4. Unfortunately, hydrogen and hydrocarbon combustion-heated facilities both result in water vapor levels in the test flow that are significantly greater than for atmospheric air. Hydrocarbon heated facilities additionally introduce carbon dioxide into the test flow. The effect of test gas vitiation due to combustion heating on scramjet combustor performance is an area of active research. Several recent studies have shown that the presence of elevated levels ofwater vapor may significantly affect combustor performance in scramjet engine test articles [1–4]. The experimental results reported in [1,2] were obtained in an electrical resistance heated wind tunnel where steam was artificially added to the flow to simulate a combustion-heated facility. Although the amount of steam addition can be calculated, an independent measurement is useful to confirm the presence of water vapor in the scramjet combustor at the reported levels. To that end, this note reports on results of the application of an instream optical measurement technique whereby water vapor levels in the test flow through the scramjet were independently measured and verified using a tunable diode laser absorption diagnostic. This paper first presents and describes the method of steam addition, and the associated makeup oxygen and optional carbon dioxide vitiation systems. Next, the tunable diode laser absorption technique that was applied to the facility is described. Finally, the added and measured water vapor levels are compared to validate the water vapor vitiation technique and to provide confidence in the similarmakeup oxygen and carbon dioxidevitiation techniques. This study represents the first time steam addition levels have been independently verified for the facility of [1,2] and it provides confidence in the claims made by the investigators in [1,2] with respect to the effects of test gas vitiation on scramjet performance and operation.