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Cryogenic wind tunnels (CWT) perform high Reynolds number aerodynamic tests in a cryogenic environment and simulate hypersonic flights. However, condensation under cryogenic conditions presents a major challenge to test’s reliability. The detailed conditions remain unclear, and traditional strategies for reducing condensation in CWT need further exploration. In this paper, an Euler-Euler two-phase flow model is established based on the classical nucleation approach and the droplet growth theory, and also presents a method for the three-dimensional transonic flow modelling. This model integrates thermodynamic properties with aerodynamic behaviour to provide a reliable simulation approach. A series of operating conditions are simulated, involving three distinct Reynolds numbers. The simulated results revealed that condensation on the wing wall is caused by two main factors: the low-pressure region at the leading edge due to the angle of attack and the pressure drop due to an airflow expansion beyond the wing’s thickest section. Subsequently, the key operating conditions that influence the condensation phenomena are identified using Correlation Matrix Analysis and Principal Component Analysis. Furthermore, this study identifies the critical regions where condensation occurs under high Reynolds number flight conditions. The findings contribute to refining operational strategies in CWT, offering improved guidelines that minimise condensation risks under varied conditions. Highlights The effects of various operating conditions on condensation were simulated. The three-dimensional characteristics were described during condensation process. Operating pressure is identified as a key factor in CWT condensation phenomena. A critical area was proposed to support optimal conditions and avoid condensation. [ABSTRACT FROM AUTHOR]

Cryogenic wind tunnels (CWT) perform high Reynolds number aerodynamic tests in a cryogenic environment and simulate hypersonic flights. However, condensation under cryogenic conditions presents a major challenge to test’s reliability. The detailed conditions remain unclear, and traditional strategies for reducing condensation in CWT need further exploration. In this paper, an Euler-Euler two-phase flow model is established based on the classical nucleation approach and the droplet growth theory, and also presents a method for the three-dimensional transonic flow modelling. This model integrates thermodynamic properties with aerodynamic behaviour to provide a reliable simulation approach. A series of operating conditions are simulated, involving three distinct Reynolds numbers. The simulated results revealed that condensation on the wing wall is caused by two main factors: the low-pressure region at the leading edge due to the angle of attack and the pressure drop due to an airflow expansion beyond the wing’s thickest section. Subsequently, the key operating conditions that influence the condensation phenomena are identified using Correlation Matrix Analysis and Principal Component Analysis. Furthermore, this study identifies the critical regions where condensation occurs under high Reynolds number flight conditions. The findings contribute to refining operational strategies in CWT, offering improved guidelines that minimise condensation risks under varied conditions.Highlights The effects of various operating conditions on condensation were simulated.The three-dimensional characteristics were described during condensation process.Operating pressure is identified as a key factor in CWT condensation phenomena.A critical area was proposed to support optimal conditions and avoid condensation.

The objectives of this study are to determine the optimum design parameters of a IBCE for hypersonic flight and to investigate the relationship between engine performance and design parameters. The investigation of these objectives is made first in the literature in this study. The optimization of inverted Brayton cycle engine, IBCE, is performed using the particle swarm optimization method in this study. The optimum specific thrust, sT, value is reached by staying within the optimization constraints. When the total temperature of the cooling section is examined, a temperature above the freezing temperature of the air is obtained. A very high sT value, 451 N.s/kg is obtained at the hypersonic flight Mach Number (5 Mach) as a result of optimization. By the investigation, it is concluded that specific fuel consumption, SFC, reduces % 5.3 and sT increases % 5.6 dependent on preburner exit total temperature, PETT, change from 2100 K to 1400 K. Based on total temperature decrease at cooling section, T_cooling, change from 100 K to 500 K, it is seen that by the investigation, SFC increases %23.7 and sT increases % 13.1. It is seen that SFC reduces by % 6.3 and sT increases by % 35.9 depending on afterburner exit total temperature, AETT, change from 2000 K to 2300 K. It is observed that SFC reduces % 10.5 and sT increases % 11.7 dependent on total pressure ratio of turbine, p_t, change from 0.9 to 0.1. [ABSTRACT FROM AUTHOR]

Aero-propulsion, in particular applications involving hypersonic flight, benefits from ultra-high temperature ceramics (UHTC) efficiency. Atomistic simulations were carried out to beter understand the thermomechanical characteristics of these materials at the elevated temperatures and pressures of hypersonic flight. To determine the first oxidation of α-SiC (silicon carbide), both the classical molecular dynamics simulation code LAMMPS (Large-scale Atomic/Molecular Massively Parallel Simulator) and the reactive ReaxFF potential were used. The ablation of SiO3, CO, CO2, C2O2, and C2O3 molecules in these simulations provide evidence of the first ablative characteristics of α-SiC as well as the distinctive parabolic reaction kinetics found in very harsh settings. The temperature range selected for these simulations was 26 °C to 2,000 °C. Due to the unique oxidation kinetics present in these extreme environments, further research into the oxidation of UHTC would require the use of newly parameterized reactive potentials. This work assesses the effect of extreme environments and the volatilization of molecules from polycrystalline surfaces as a result of the unique oxidation kinetics that are present in these extreme environments. [ABSTRACT FROM AUTHOR]

A comprehensive state-to-state (StS) model for air was formed, including vibration-vibration-translation (VVT) reactions. The generated VVT reaction rates were compared with available first-principles calculations, and reasonable agreement was obtained. The influence of multiquantum transitions revealed that the possible reduction on the number of VVT transitions depends on the application. The influence of VV transitions revealed that the vibrational excitation becomes too fast if VVT transitions are reduced to VT transitions, invalidating this approximation. Comparisons were made with existing NO emission measurements in the ultraviolet and mid-infrared spectrum. At velocities of 3-4 km/s, some of the current O2 dissociation rates may be inaccurate. At a higher velocity of 6.81 km/s, the NO mole fraction predicted by the StS model is around an order of magnitude greater than that predicted by the two-temperature model. Finally, recommendations were given on the further development of the StS model. [ABSTRACT FROM AUTHOR]

Real gas effects in high entry temperatures are taken into consideration combining flight conditions during hypersonic flight into Earth's atmosphere. Convective stagnation heating loads are estimated from the original full Fay-Riddell formula and compared with other engineering analytical methods. A new engineering correlation formula is proposed for wall stagnation temperature distribution during reentry motion. The efficiency of the applied analysis gives satisfactory results compared with published data of verified experiments and computational codes for the lifting Space Shuttle STS-40 and ballistic Apollo4 capsule missions. [ABSTRACT FROM AUTHOR]

The interaction of an off-axis, spherical energy discharge with the Kh-47M2 Kinzhal missile at Mach 6 sea-level conditions was simulated. A parametric study of the energy discharge parameter ϵ (proportional to the ratio of the energy added to the internal energy within the energy deposition volume) was performed to examine the effect of energy added to the heated region in front of the missile and the changes it causes on the drag force, side (vertical) force, and pitching moment. The peak value for the pitching moment coefficient was shown to be 0.0045 for ϵ=400, which was shown to be 15.5% of the value due to a 10° front finset deflection. Both inviscid and Reynolds-averaged Navier-Stokes simulations were performed. The additional influence of the turbulent boundary layer on the peak value of the produced pitching moment was negligible. The time interval for the variation in the pitching moment coefficient due to energy deposition is 61 times faster than the standard actuation time (0.1 s) for conventional control surfaces, indicating the potential for multiple pulsed energy depositions as a means for hypersonic flight control. [ABSTRACT FROM AUTHOR]

The accuracy and sensitivity of plasma generation predicted by several different chemical kinetics models is investigated in the context of weakly ionized hypersonic flowfields around the Radio Attenuation Measurement (RAM-C) vehicle. A computational fluid dynamics analysis is used to examine 13 independent trajectory points along the RAM-C II flight, and an assessment of the chemistry models is made by comparing results to available flight measurements. The limitations of making such comparisons with the raw flight data are established in detail, including the inherent shortcomings associated with interpolating the flight data to assess a single trajectory point. Two separate geometries are evaluated in this study, as the initial RAM-C geometry was altered during flight after its nose cap was pyrotechnically ejected. The blunter post-ejection geometry generates more electrons in the stagnation region. In general, good agreement is found between each chemistry model and flight data from both the electrostatic probe and reflectometer stations above 56 km. An expected sizable gap exists between the simulations and reflectometer data at lower altitudes. The impact of forward reaction rates, equilibrium constants, and number of species varies considerably based on altitude, velocity, and position along the body. [ABSTRACT FROM AUTHOR]

To date, plasma sheath effects have not been incorporated into most CFD simulations of magnetohydrodynamics (MHD) or aerothermodynamics due to the high computational costs involved. The accurate modeling of both collisional and collisionless plasma sheaths for hypersonic applications is here presented. This is followed by the outline of recent numerical methods that enable the sheaths to be integrated with large time steps and to be incorporated within CFD codes. For the first time, the electron energy transport in non-equilibrium is here integrated in coupled form with the plasma sheath within a MHD process. Some example test problems are then presented including MHD in a scramjet combustor and electron transpiration cooling in a hypersonic flow. Such show that the plasma sheath effects can be added within CFD simulations with little loss in computational efficiency. The inclusion of plasma sheath models is seen to be critical in predicting accurately such problems. [ABSTRACT FROM AUTHOR]

The article describes a robust technology for numerical simulation of multicomponent flows in the presence of strong shock waves, for example, in the flight of a high-speed vehicle with a ramjet/scramjet engine. This takes into account the phenomena of turbulence, multicomponent diffusion and heat transfer. High-temperature phenomena that occur when moving at a high supersonic speed and are associated with dissociation and ionization of the gas environment are not considered, although they fit well into the framework of the general model of a multicomponent reacting gas. The advantage of the algorithm proposed in this article is its adaptability to both the simulation of strong shock waves and to the simulations on small scales, as well as a high potential for parallelization. Illustrative examples of simulations based on the developed method are given. [ABSTRACT FROM AUTHOR]

Scramjet engines are propulsion systems with high specific impulse performances, which are widely implemented for the realization of hypersonic flight. In this study, the influence of global hydrogen equivalence ratios on supersonic combustion under hypersonic flow was evaluated by conducting a large eddy simulation. First, numerical simulation results revealed a good agreement between the experimentally determined wall pressure and shockwave signal, which verified the accuracy of the numerical method. Moreover, the influence of the equivalence ratio on the characteristics of supersonic combustion was investigated. The results revealed that the complexity of the waveform increased with an increase in the equivalence ratio, owing to the higher combustor pressure and significant interactions between the flames and boundary layer. In particular, the combustion expanded to the center of the flow field, and the average Mach number of the flow field decreased. In addition, the size of the auto-ignition region decreased; however, it was under supersonic flow. The influence of the equivalence ratio on the heat release rate (HRR) under different combustion modes was investigated. The results revealed that the total HRR increased with an increase in the equivalence ratio. Moreover, the majority HRR was in the auto-ignition region, and supersonic combustion was pre-dominant. This study was based on the scramjet model engine without flameholders under hypersonic conditions; therefore, the findings cannot be applied to scramjet combustors under other conditions. • Different global hydrogen equivalence ratios in a scramjet are simulated by LES. • The combustion characteristics in a scramjet are studied. • The heat release rate of different global hydrogen equivalence ratios is assessed. [ABSTRACT FROM AUTHOR]

Recent developments in the world of hypersonic flight have brought increased attention to the thermal response of materials exposed to high-enthalpy gases. One promising concept is electron transpiration cooling (ETC) that provides the prospect of a passive heat removal mechanism, rivaling and possibly outperforming that of radiative cooling. In this work, non-equilibrium CFD simulations are performed to evaluate the possible roles of this cooling mode under high-enthalpy conditions obtainable in plasma torch ground-test facilities capable of long flow times. The work focuses on the test case of argon gas being heated to achieve enthalpies equivalent to post-shock conditions experienced by a vehicle flying through the atmosphere at hypersonic speed. Simulations are performed at a range of conditions and are used to calibrate direct comparisons between torch operating conditions and resulting flow properties. These comparisons highlight important modeling considerations for simulating long-duration, hot chamber tests. Simulation results correspond well with the experimental measurements of gas temperature, material surface temperature as well as measured current generated in the test article. Theoretical methods taking into consideration space charge limitations are presented and applied to provide design suggestions to boost the ETC effect in future experiments.

SU2-NEMO, a recent extension of the open-source SU2 multiphysics suite’s set of physical models and code architecture, is presented with the aim of introducing its enhanced capabilities in addressing high-enthalpy and high-Mach number flows. This paper discusses the thermal nonequilibrium and finite-rate chemistry models adopted, including a link to the Mutation++ physio-chemical library. Further, the paper discusses how the software architecture has been designed to ensure modularity, incorporating the ability to introduce additional models in an efficient manner. A review of the numerical formulation and the discretization schemes utilized for the convective fluxes is also presented. Several test cases in two- and three-dimensions are examined for validation purposes and to illustrate the performance of the solver in addressing complex nonequilibrium flows.

A tunnel/predictor display which presents guidance information in a 3-dimensional format is considered for improving trajectory control in hypersonic flight. The displayed 3-dimensional information comprises a tunnel image and a predictor for indicating the aircraft position at a specified time ahead. The 3-dimensional guidance information is introduced to support the pilot in controlling the flight path. It is considered that piloting problems can be avoided which exist with conventional trajectory control techniques due to path-attitude decoupling. A predictor control law is constructed which yields controlled element properties (predictor-aircraft system) requiring minimum pilot compensation. This predictor control law forms the basis of the trajectory control improvement goal. Results from hypersonic flight simulation tests at the NASA Dryden Flight Research Center are presented for experimental verification. This paper is an outcome of a joint research effort of the NASA Dryden Flight Research Center, and the Institute of Flight Mechanics and Flight Control of the Technische Universität München and the Department of Aeronautics, Naval Architecture and Railway Vehicles (former Department of Aircraft and Ships at the Budapest University of Technology and Economics. [ABSTRACT FROM AUTHOR]

The hypersonic waverider using the morphing configuration can improve the overall performance and broaden the flight range. On the other hand, the hypersonic waverider belongs to the lightweight construction, leading to the presence of the intrinsic elasticity. As a result, the integrated consideration of the active deformation and elastic action is important for the morphing waverider to realize global optimization. In this paper, the contro-integrated design methods are proposed for the morphing waverider to implement the trade-off analysis between the active deformation and intrinsic elasticity. First, the parametric model is built to describe the hypersonic flight characteristics. Then, the elastic dynamics is introduced to quantify the coupling responses associated with the angle of attack and elevon deflection. Furthermore, the integrated design is considered for the active deformation and inherent elastic action, and also the control law is designed to suppress the active and passive morphing disturbances. Finally, the simulation is conducted to test the effectiveness of the presented methods for the morphing waverider. [ABSTRACT FROM AUTHOR]

This paper discusses the influences of plasma sheath on antennas during hypersonic flights. The practical plasma sheath around an RAM-C vehicle when flying at 25.91 Ma at the altitude of 71 km is evaluated with the fluid analysis method, both dipole antenna and horn antenna are located at the stagnation region and aft body region, respectively, and the antenna performances are compared with those when the plasma sheath does not exist. Numerical results show that when the antennas are located at the stagnation region covered by plasma sheath, for dipole antenna, the gain decreases 7.3 dB and the frequency shifts to higher band by 9.88%, while for the horn antenna, the gain can decrease 4.01 dB and the return loss gets worse. However, when the antennas are located at the aft body region, the antenna performances show hardly any changes no matter the plasma sheath exists or not. The conclusion obtained may provide useful insights for designing antennas on hypersonic vehicles. [ABSTRACT FROM PUBLISHER]

In order to have an efficient control of its flight, hypersonic flight vehicle needs to have an accurate control of its airflow attitude angle. On the basis of realizing the accurate tracing of airflow attitude angle command, the projected flight path can be realized to accomplish flight missions. Considering the input time-delay caused by the system signal transmission of the aircraft and the actuator dynamic, compensate for the adverse effects caused by it, and design a time-delay compensation controller, which realizes the tracking control of airflow attitude angle for hypersonic flight vehicles with the input time-delay. Simulation result shows that this method has better properties of tracking control and rapid convergence. [ABSTRACT FROM AUTHOR]

Electromagnetically launched projectiles fly with hypersonic speeds in different regions of the Earth’s atmosphere. Because of their hypersonic speeds, these projectiles are designed to withstand extreme thermal loads. Drag forces should also be considered to maximize the operational range. To reduce the thermal loads and drag forces, the geometric shape of these projectiles should be carefully designed. We can utilize either experimental or numerical methods to calculate these heating effects and drag forces. Numerical methods are more economic in terms of monetary cost and time. However, we cannot use the same numerical method in different regions of the Earth’s atmosphere. In this paper, we used direct simulation Monte Carlo method to calculate thermal loads and drag forces of four different projectile geometries in the rarefied part of the atmosphere. Simulation results show that thermal loads and drag forces vary considerably depending on the projectile geometry. [ABSTRACT FROM PUBLISHER]

The waverider configuration is a type of aerodynamic shape providing a high lift-to-drag ratio and reasonable flight capacity for hypersonic vehicles. However, hypersonic waveriders are marred by serious design problems wherein the flight performance continually deteriorates if there are deviations in the expected flight conditions. As a result, a combination of hypersonic vehicle design with intelligent morphing technologies may be advantageous in addressing flight stability issues experienced during large flight envelopes. This paper discusses the concepts of nonlinear dynamics and integrated design methods with regard to morphing waveriders. First, a parameterized longitudinal model of a hypersonic vehicle is established in terms of the resulting lift, drag, moment, and thrust obtained from oblique shock theory, the Prandtl-Meyer equation, and the quasi-one-dimensional Rayleigh flow principle. Then the influence of active deformation on waverider control quality is considered by adjusting the configuration parameters in relation to the control action. Finally, the control system of a morphing waverider is designed using the model reference adaptive control methods, and a simulation example for morphing waveriders is demonstrated with regard to integrated action among aerodynamics, propulsion, and control. [ABSTRACT FROM AUTHOR]

The most recent idea of using the exothermic coating to reduce the drag coefficient on blunt noses in hypersonic flows is numerically simulated to study parametrically the effect of material properties. The conjugate heat transfer is considered by simultaneous solution of flow and solid phase governing equations in numerical procedure. The results show that changing some effective parameters in material properties may increase the coating effectiveness significantly. Using such analysis leads to choose the proper materials for specific flight missions. [ABSTRACT FROM PUBLISHER]

Two-stage winged space access vehicles consisting of a carrier stage with airbreathing turbo/ram jet engines and a rocket propelled orbital stage which may significantly reduce space transport costs and have additional advantages offer a great potential for mission safety improvements. Formulating the nominal mission and abort scenarios caused by engine malfunctions as an optimal control problem allows full exploitation of safety capabilities. The shaping of the nominal mission has a significant impact on the prospective safety. For this purpose, most relevant mission aborts are considered together with the nominal mission, treating them as an optimization problem of branched trajectories where the branching point is not fixed. The applied procedure yields a safety improved nominal trajectory, showing the feasibility of the included mission aborts with minimum payload penalty. The other mission aborts can be separately treated, with the initial condition given by the state of the nominal trajectory at the time when a failure occurs. A mission abort plan is set up, covering all emergency scenarios. [ABSTRACT FROM AUTHOR]

Recent developments in the world of hypersonic flight have brought increased attention to the thermal response of materials exposed to high-enthalpy gases. One promising concept is electron transpiration cooling (ETC) that provides the prospect of a passive heat removal mechanism, rivaling and possibly outperforming that of radiative cooling. In this work, non-equilibrium CFD simulations are performed to evaluate the possible roles of this cooling mode under high-enthalpy conditions obtainable in plasma torch ground-test facilities capable of long flow times. The work focuses on the test case of argon gas being heated to achieve enthalpies equivalent to post-shock conditions experienced by a vehicle flying through the atmosphere at hypersonic speed. Simulations are performed at a range of conditions and are used to calibrate direct comparisons between torch operating conditions and resulting flow properties. These comparisons highlight important modeling considerations for simulating long-duration, hot chamber tests. Simulation results correspond well with the experimental measurements of gas temperature, material surface temperature as well as measured current generated in the test article. Theoretical methods taking into consideration space charge limitations are presented and applied to provide design suggestions to boost the ETC effect in future experiments. [ABSTRACT FROM AUTHOR]

SU2-NEMO, a recent extension of the open-source SU2 multiphysics suite's set of physical models and code architecture, is presented with the aim of introducing its enhanced capabilities in addressing high-enthalpy and high-Mach number flows. This paper discusses the thermal nonequilibrium and finite-rate chemistry models adopted, including a link to the Mutation++ physio-chemical library. Further, the paper discusses how the software architecture has been designed to ensure modularity, incorporating the ability to introduce additional models in an efficient manner. A review of the numerical formulation and the discretization schemes utilized for the convective fluxes is also presented. Several test cases in two- and three-dimensions are examined for validation purposes and to illustrate the performance of the solver in addressing complex nonequilibrium flows. [ABSTRACT FROM AUTHOR]

Hypersonic transport (HST) aircraft are the subject of much current research as the need to reduce block times on long haul flights becomes more pressing. The dynamics of such HST aircraft are strongly influenced by the close interdependence of the aerodynamic and propulsion systems. The key dynamic features of such aircraft are their dynamically unstable longitudinal motion, and extremely sluggish response in pitch attitude, which make aircraft of this type difficult to fly. This paper presents the design of an automatic flight control system which can provide the required degree of stability and the required flying qualities. The optimal feedback control law was derived from the use of linear quadratic regulator (LQR) theory, in which the state weighting matrix was calculated from known closed-loop eigenvalues which were specified by the designer to obtain the required stability and flying qualities. The effectiveness of the multivariable control law is illustrated by means of dynamic responses obtained from a simulation based upon an HST aircraft flying at Mach 8.0 at a height of 85 000 feet. [ABSTRACT FROM PUBLISHER]

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. [ABSTRACT FROM AUTHOR]

• An aerodynamic heating model is derived from the molecular collision analysis. • The theoretical model is verified by molecular dynamics simulations. • The theoretical model is more accurate than the traditional NS equations. • The aerodynamic heating is attributed to the phonons excited by the colliding atoms. An aerodynamic heating model is derived from molecular collision analysis, in which the rising temperature of a hypersonic flying object as a function of the flying speed in a classical dense monoatomic gas environment is set up. The model predicts that the rising temperature of the hypersonic flying object is independent of the gas density but depends linearly on the gas atomic mass. A nonequilibrium molecular dynamics simulation is carried out to verify the theoretical model. Also, through analyzing the vibrational density of states in the flying objects, it is found that the excited phonon frequency is near the collision frequency, uncovering that the phonons are mainly excited by the face colliding gas atoms. Our study provides a new insight into understanding the intrinsic mechanism of aerodynamic heating and helps to develop the temperature-controllable hypersonic flying vehicle. [ABSTRACT FROM AUTHOR]

Numerical simulations of the 3D reacting flow in a model supersonic combustor with kerosene as the fuel have been carried out. The combustor entry Mach number is nominally 2.5. Tetrahedral grids with wall y+ less than 75 have been used. Details of the flow field inside the combustor such as contours of static pressure, static temperature, species mass fraction, reaction rate and Mach number are discussed. Also, the variation of metrics such as combustion efficiency and total pressure loss along the length of the combustor are presented. Predicted values of wall static pressure along the top wall of the combustor are compared with previously reported experimental data. The calculations over predict the pressure rise due to combustion in the cavity region, but the overall pressure rise is predicted well. [ABSTRACT FROM AUTHOR]

In this paper, numerical simulations of a full scale concept supersonic combustor using kerosene fuel are presented. Results are reported first for a baseline model. These results are then used for identifying the shortcomings in this model, which allows improvements to be made. In particular, the spacing between the ramps is widened so that the heat addition is not confined to a short region and is distributed over a larger volume. Also the design of fuel supply lines has been reworked so that they do not obstruct the supersonic flow. Details of the flow field inside the combustor such as contours of static pressure, static temperature, species mass fraction, reaction rate and Mach number are discussed. Wall static pressure distribution along the top wall of the combustor is presented. Intricate details of the flow field such as separation zones, shocks and expansion fans and their interaction are also brought out. Finally, thrust developed by the combustor is calculated using two different methods. [ABSTRACT FROM AUTHOR]