Supersonic Exhaust Research Articles

Optimization of the MSMGWB models used to predict remote infrared signals of jet engine in various spectral intervals

Further improvements are reported for an optimization method that was used in previous work to determine the best combination of Gauss quadrature schemes and wavenumber subinterval division results for the multi-scale multi-group wide-band k-distribution (MSMGWB) model, which is used in calculation of remote infrared signals of hot combustion gases from hydrocarbon fuels and gray walls in the 3–5 µm spectral interval. In addition, the application range of the MSMGWB model was extended to the 8–14 µm, 7.7–9.7 µm, 3.7–4.8 µm, and 2–2.5 µm spectral intervals. Compared with previous works, the new optimization procedure considered a broader range of thermodynamic state parameter combinations of the combustion gas and the atmosphere. The optimization effect was evaluated further via prediction of remote sensing thermal images in various spectral intervals for solid structures and hot jet of an aero-engine’s supersonic exhaust system. Furthermore, an optimized MSMGWB model that considers the spectral response characteristics of infrared detectors has been established and evaluated.

Spectral feature extraction of rocket exhaust plume using spectral proper orthogonal decomposition

The spectral characterization of flow-field parameters provides a new perspective for understanding the spatiotemporal evolution of unsteady supersonic exhaust plumes and for extracting typical structures. In this study, a large-eddy simulation is performed to calculate the three-dimensional unsteady supersonic plume flow field of rocket engines, and a spectral proper orthogonal decomposition (SPOD) method with a spatiotemporal separation is established. This approach is used to extract the coherent structural features of the unsteady exhaust plume flow field and analyze the mode space structure at different frequencies. The three-dimensional reconstruction and denoising of the exhaust plume flow-field parameters can be achieved via the frequency- and time-domain reconstructions of the SPOD algorithm and oblique projection method, respectively. The ground rocket exhaust plume of ballistic evaluation motor-II is analyzed. The results indicate that the SPOD method can effectively extract the single-frequency mode structure of the reactive supersonic flow field, and that low-order behavior appears in the m = 0 and m = 1 azimuth modes. The potential core exhibits a high-frequency wave-packet structure that is affected by shock waves and shear layers. Time-domain reconstruction based on the oblique projection method facilitates the capture of the dynamic characteristics of the flow field. For the first-order SPOD mode, the frequency- and time-domain reconstruction errors are 3.3% and 1.5%, respectively. The frequency-domain reconstruction method exhibits a 4% improvement in denoising ability compared to low-pass filters. This study provides a novel method for the spectral characterization and spatiotemporal feature extraction of supersonic exhaust plume flow fields.

Gas dynamics at starting and terminating phase of a supersonic exhaust diffuser with a conical nozzle

This research investigates the process of starting and breakdown of second throat diffuser during high-altitude test of conical nozzle with a high expansion ratio. The subscale experimental setup includes a conical nozzle with an expansion ratio of 53, plus a second throat diffuser with a contraction ratio of 1.85, using compressed air as the working fluid. Numerical simulation has been employed to identify the flow physics during the unsteady process of diffuser startup and breakdown. According to the results, the starting process is divided into three distinct stages. In the first stage, the exhausted flow from the nozzle enters the vacuum chamber, leading to an increase in vacuum chamber pressure. The second stage, corresponding to the period before full flow establishment in the conical nozzle, exhibits a relatively constant slope. In the final stage, associated with the transition from over-expanded to under-expanded states, the slope of the rate of evacuation development descends. Next, the vacuum degradation in termination process has been analyzed, and it has been found to include three stages: high slope, middle slope, and low slope. The flow physics during the start process, similar to the results observed in other conical nozzles, exhibits only a Mach reflection structure. However, during the termination process, the flow physics involve a combination of both structures, including Mach reflection and Cap Shock. The results indicate that during the start, an internal shock only interacts with the separation shock, and no special change occurs in the Mach reflection structure. In contrast, during the termination process, unlike what has been reported in previous studies on conical nozzles, the structure of cap shock waves and restricted shock separation patterns are also observed. Another distinction between the starting and termination processes is related to the pressure distribution in the diffuser wall. The wall pressure at the diffuser inlet during the termination process has been reported to be 90% higher than during the starting process.

Designing jet deflector configuration for a semi-cryogenic rocket engine

This study deals with the design of a suitable configuration of jet deflector for supersonic exhaust gases from a semi-cryogenic launch vehicle engine using computational fluid dynamics. Computational model of combustion in semi-cryogenic engine, with kerosene as fuel and liquid oxygen as oxidizer, was developed in which exhaust gases from engine were impinged on the deflector surface. Different design configurations of deflector structure, obtained by the combination of various impingement angles of 15°, 20°, 25°, 30°, and 45°, and different exit radii of 5000, 10 000, and 20 000 mm, were used in the developed computational models to analyze impingement, deflection, and associated flow properties of exhaust gases, such as acoustics, force on deflector imparted by jet, temperature, and Mach number of flow field. A new ablation model was developed based on Vieille's law to determine the amount of refractory material ablated from deflector structure for different configurations. The developed models of ablation, combustion, and flow impingement were also validated with existing literature. It was found that the configuration with 20° impingement angle and 10 000 mm exit radius had ablation, acoustics, force, and flow properties in desired limit. Furthermore, to find the optimum uplift angle for designing jet deflector, configurations with uplift angles of 0°, 5°, 15°, and 25° were studied using the computational models developed in the study. It was observed that the configuration with 20° impingement angle, 10 000 mm exit radius, and 15° uplift angle was best suited for impingement and deflection of exhaust jet from the specified semi-cryogenic engine.

Numerical Simulations of Supersonic Exhaust Diffuser At Various Operating Conditions

A numerical study was carried out to determine starting characteristics of a straight cylindrical supersonic exhaust diffuser of a high altitude simulation of high expansion rocket nozzle. The starting conditions of the supersonic exhaust diffuser are numerically simulated for contour nozzle of exit-to-throat area ratio of 21.77, diffuser diameter-to-nozzle throat area ratio of 23.36, and diffuser length-to-diameter ratio of 15. The evacuation performances of supersonic exhaust diffuser at various operating conditions are investigated with help of the Mach contour plots. The Mach variations along the centre line and wall of the diffuser indicate the location of the shock and flow characteristics. Validation of numerical simulation is carried out with experimental data and shows a good agreement between them. Flowfield inside the straight cylinder supersonic exhaust diffuser is presented out for stagnation to ambient pressure ratio of 5.5 - 28. The nozzle is started at the stagnation to ambient pressure ratio of about 17. The straight cylindrical supersonic exhaust diffuser is started at the stagnation pressure to ambient pressure ratio of about 21. The numerical simulations of the flowfield are able to capture the flows filled in the diffuser and the shock system is fully established in the duct.

Material selection based on joule heating simulation for resistojet thruster

Resistojets generate heat by passing an electric current through a resistive element, thereby heating a propellant at high pressure, which is subsequently expanded through a nozzle resulting in a supersonic exhaust jet. Specific impulse of a resistojet decides the performance of the thruster. Hence the goal here is to achieve highest specific impulse. The specific impulse of the resistojet is limited by the thermal design of the heater and the material selection. Since the design specifications on the resistojet model has been confirmed with calculated data points obtained from literature work, the material selection is done with the heat transfer simulation. Also, the materials selected as the simulation model is done on accordance with highest melting point and the factor of feasibility is taken into account. For achieving the heat transfer distributions among the flow channels through conduction allows the calculation of heat flux along the radial direction. The transient state condition of the turbulent fluid region is solved under the CFX solver. The simulation is done on different materials varying their thermophysical properties. The end result will be decided on the basis of the heat transfer rate and the residual obtained during higher iterations. Thus, this paper focuses on the choice of materials and the thermal analysis of the same to achieve the highest specific impulse and an efficient performance output.

Numerical Investigations At High Altitudes For Nozzles Performance Monitoring

Large area ratio nozzles that are employed for space propulsion applications are not meant to be tested at the ground-level conditions because of the flow conditions that occurs in their divergent section due to their low exhaust pressures. Therefore, it is desired to perform the experimental evaluation at high-altitude to evaluate the performance of the said nozzles. Generally, a high-altitude test facility, consisting of a supersonic exhaust diffuser, is generally employed. In thispaper, a second-throat exhaust diffuser has been numerically investigated to predict its minimum starting pressure, and to understand the flow physics in the diffuser using Computational Fluid Dynamics (CFD). Numerical computation of the flow field in the diffuser system is done for the two cases: 1) when the nozzle and the diffuser system are initially evacuated to a low pressure, and 2) when the nozzle and the diffuser system are initially at ambient pressure (1 bar). Simulations have been carried out for cold flow situation (Ȗ=1.4) over a range of nozzle inlet stagnation pressures. Numerical results compare favorably with the theoretical and experimental results and provide adequate insight to the flow physics and internal shock structures formed in thediffuser in addition to minimum starting pressure, diffuser wall pressure and vacuum thrust.

Afterburning Effects on Thermal Environment of Liquid Oxygen/Kerosene Rocket Under Low Altitudes

Heavy launch vehicles use liquid oxygen (LOX) and kerosene propellants for their high specific impulse and controlling precision. However, incomplete burnt exhaust gas reacts easily with air, which will influence the thermal environment. To investigate the afterburning effect on the plume flowfield at low altitudes, we established a supersonic exhaust plume model by using the three-dimensional Reynolds-averaged Navier–Stokes (RANS) methods and realizable turbulence model. Also, the nine-species and fourth-step chemical reactions are adopted to simulate the afterburning process. The validity of our numerical model is confirmed by the good agreement between calculation results and experiment data. On this basis, the thermal environment of the LOX/kerosene rocket between frozen and reaction flow at five typical altitudes are calculated and compared. The numerical results show that the secondary combustion reactions mainly occur in the mixing layer, with about an 11.9–18.6% increase in the peak temperature after considering the afterburning effects. The increase in range of the maximum temperature of the engine exhaust plume gradually reduces as the flight height increases. At the same altitude, the afterburning has more significant effect on flowfield for a larger distance from the nozzle exit. The results provide a vital foundation of analysis and calculation for the design of the LOX/kerosene rocket thermal protection.

Supersonic Exhaust from a Rotating Detonation Engine with Throatless Diverging Channel

Converging–diverging nozzles are common in rocket engine systems to increase the exhaust velocity and improve thrust performance. In this study, we focused on the acceleration of subsonic burned gas without a structural throat via detonation to realize a simple and compact engine. We developed and tested a rotating detonation engine (RDE) without a throat and with a diverging channel (constant diverging angle ). Gaseous and were used as the propellants, and the mass flow rate ranged from 62 to in the combustion tests under low back-pressure conditions. We measured pressure and thrust, as well as high-speed imaging of self-luminescence of the combustion and imaging of the exhaust plume. The pressure at the exit was less than one-fifth of the maximum pressure in the RDE, significantly below the value for a sonic flow. The results suggested that the exhaust flow was supersonic, with values up to Mach 1.7, without the need of a converging section within the engine. In addition to the estimated Mach number from the measured pressure, the exhaust plume images coherently indicated the existence of supersonic exhaust.

Numerical Study on the Flow and Heat Transfer Characteristics of a Second Throat Exhaust Diffuser According to Variations in Operating Pressure and Geometric Shape

A numerical study was conducted to investigate the flow and heat transfer characteristics of a supersonic second throat exhaust diffuser for high-altitude simulations. The numerical results were satisfactorily validated by the experimental results. A subscale diffuser using nitrogen was utilized to investigate starting pressure and pressure variation in the diffuser wall. Based on the validated numerical method, the flow and heat transfer characteristics of the diffuser using burnt gas were evaluated by changing operating pressure and geometric shape. During normal diffuser operation without cooling, high-temperature regions of over 3000 K appeared, particularly near the wall and in the diffuser diverging section. After cooling, the flow and pressure distribution characteristics did not differ significantly from those of the adiabatic condition, but the temperature in the subsonic flow section decreased by more than 1000 K. Furthermore, the tendency of the heat flux from the diffuser internal flow to the wall was similar to that of the pressure variations, and it increased with operating pressure. It was confirmed that the heat fluxes of the supersonic and subsonic flows in the diffuser were proportional to the operating pressure to the 0.8 and −1.7 power, respectively. In addition, in the second throat region after separation, the heat flux could be scaled to the Mach number ratio before and after the largest oblique shock wave because the largest shock train affected the heat flux of the diffuser wall.

Characterization of novel ceramic composites for rocket nozzles in high-temperature harsh environments

This paper presents the results of experimental tests for the characterization of Ultra-High-Temperature Ceramic Matrix Composite (UHTCMC) materials for near-zero erosion rocket nozzles. Two dedicated test set-ups were developed for preliminary screening of material candidates in a representative environment, characterized by relevant heat flux and temperature. The experimental set-up was based on a lab-scale 200N-class hybrid rocket engine, employing gaseous oxygen as the oxidizer and High-Density PolyEthylene as fuel; the configurations included free-jet test, in which small button-like samples were exposed to the supersonic exhaust jet of the rocket nozzle; and chamber inserts, in the shape and size of an annular element, placed inside the rocket combustion chamber. Computational Fluid Dynamic simulations, for modeling heat transfer and combustion chemical reactions, complemented the experimental observations and supported the characterization of test conditions. Samples with ZrB2-SiC matrix and continuous or chopped carbon fibers, sintered by either Hot Pressing or Spark Plasma Sintering were tested. Free-jet test samples demonstrated a substantially improved erosion resistance with respect to conventional graphite and in one case a negligible material recession. UHTCMC samples erosion was associated to the occurrence of a rapid rise in surface temperature, which achieved values over 2900 K. Chamber inserts, besides confirming the outstanding erosion resistance of UHTCMCs with respect to traditional materials (i.e. C/SiC), proved that long-fibers samples with sufficient porosity are more likely to withstand thermal shocks typical of the rocket combustion environment.

Numerical studies on four-engine rocket exhaust plume impinging on flame deflectors with afterburning

This paper studies the four-engine liquid rocket flow field during the launching phase. Using three-dimensional compressible Navier-Stokes equations and two-equation realizable k-epsilon turbulence model, an impact model is established and flow fields of plume impinging on the two different shapes of flame deflectors, including wedge-shaped flame deflector and cone-shaped flame deflector, are calculated. The finite-rate chemical kinetics is used to track chemical reactions. The simulation results show that afterburning mainly occurs in the mixed layer. And the region of peak pressure occurs directly under the rocket nozzle, which is the result of the direct impact of exhaust plume. Compared with the wedge-shaped flame deflector, the cone-shaped flame deflector has great performance on guiding exhaust gas. The wedge-shaped and cone-shaped flame deflectors guide the supersonic exhaust plume away from the impingement point with two directions and circumferential direction, respectively. The maximum pressure and temperature on the wedge-shaped flame deflector surface are 37.2% and 9.9% higher than those for the cone-shaped flame deflector. The results provide engineering guidance and theoretical significance for design in flame deflector of the launch platforms.

Numerical analysis on thermal environment of liquid rocket with afterburning under different altitudes

To achieve a detailed understanding of the characteristics of the liquid hydrogen-oxygen (LH2/LOX) rocket exhaust plume with afterburning, the three-dimensional compressible Reynolds-averaged Navier-Strokes (RANS) equation with the realizable k-Ɛ turbulence model is applied to calculate the flow field of the supersonic exhaust gas. An optimized 9-species and 14-steps chemical mechanism is used for the afterburning effects. Multi-grid method is adopted to establish structural grid of 8.87 million cells for the finite volume computation. The afterburning model is established under six inflow conditions and then validated by experimental data. The validated model is applied to compare the under-expanded supersonic plume with and without afterburning reaction at different flight altitudes. The simulation results indicate that afterburning mainly occurs in the mixed layer. Compared to the frozen flow, the peak temperature of the reaction flow has increased from 123 K to 318 K, the mole concentrations of hydrogen and water vapor change in the reacting plume. The afterburning has significant influence on the thermal environment of the exhaust plume at low altitude, but its effect becomes weak with the altitude increase. The results have provided a relevant evaluation of the flow field characteristics of supersonic plume and offer some usefulness to further numerical simulations of the liquid rocket exhaust jet with afterburning.

The generation mechanism of higher screech tone harmonics in supersonic jets

The generation mechanism of screech harmonics in supersonic exhausts is revealed using shadowgraph imaging and acoustic far-field measurements for a rectangular, high aspect-ratio nozzle. The coher ...

Numerical Studies on Flowfield Characteristics of Four-Engine Rockets Impinging Jet with Different Flame Deflectors

[Objective] To explore the further deep space, many countries are developing the new generation launch vehicle which has higher effective loads [1, 2]. To upgrade capacity, a scheme with multiple engines working simultaneously has often been adopted in the propulsion system of heavy launch vehicles [3, 4]. Therefore, the rocket jet impact problem has been the vital factor which seriously affected the launch safety. Due to the rocket engines are close to the launch platform during launching phase, the exhaust jets of rocket engine not only disturb the attitude of rocket in the take-off, but also have a thermal impact on the rocket base and ground control system [5]. Potential dangers would threaten secure and stable of launch missions if these problems could not be solved properly. Therefore, carrying out research on the flow field characteristics of multi-engine rockets exhaust plume has a great significance for the risk aversion and improvement of achievement ratio in launching [6-9]. This paper researches the four-engine rockets flow field during launching phase, compares and analyzes the plume impact characteristics with two shapes of flame deflector by numerical simulation methods. [Methods] The study object of this paper is a three-stage liquid rocket which utilizes kerosene/liquid oxygen (LOX) propellant as propulsion system for the first stage. Structured meshes with tensor product structure have better numerical ability than unstructured grids that using a multi-block, structured mesh to improve the fidelity of the flow solutions and assure a good resolution of shock waves in the flow field. The liquid rocket exhaust is considered as an ideal gas with the continuum assumption. The numerical solution of the compressible, multi-species, Reynolds-Averaged Navier-Stokes equations for multi-component flow has been obtained with finite volume spatial discretization on a structured three-dimensional grid implemented. The supersonic jet impingement on deflector was computed the realizable k-ε turbulence model. The model has shown superior improvements over the standard-model where the flow features include strong streamline curvature and vortices. The accuracy and effectiveness of the impact model was verified by comparison between the calculated results and the measured date. [Results] The temperature, pressure, and contours on the symmetry plane with wedge-shaped and cone-shaped deflectors are obtained, three shock cells are established before the plume impacts the deflector surface. A high temperature region occurs immediately through excessive high temperature gas accumulation on the deflector surface. The core and developed region and boundary of plume are clearly visible downstream of flow field. The supersonic exhaust plume was diverted into outlet with two directions and circumferential direction in the wedge-shaped and cone-shaped flame deflector, respectively. It can be shown from the temperature and pressure contours in the deflector surface that the region of peak temperature and pressure occur directly under the rocket nozzle, and the temperature drops gradually from the center to the outlet of the flame deflector. The simulation results show the planar static temperature contours for the rocket motor gas impingement on wedge-shaped and cone-shaped deflectors at Z/D = 3, 4.5, and 6 in axial plume direction, where Z/D = 0 corresponds to the nozzle exit. The exhaust gas inlet of the wedge-shaped deflector has the same temperature as atmosphere except for jet flow region, and not affected by the reverse flow. The temperature edge region of planar in the cone-shaped deflector, as well as the central region, is much higher than ambient temperature. The temperature in the edge at Z/D =3 where is closest to the exhaust gas inlet may reach up to 800 K. [Conclusions] By comparative analysis between the four-engine rockets impinging jet on different deflector, the main conclusions are as follow. The region of peak temperature and pressure occurs directly under the rocket nozzle. The wedge-shaped and cone-shaped deflectors have different diversion directions. The maximum pressure and temperature in the wedge-shaped deflector are, respectively, 62% and 18% higher than those in the cone-shaped flame deflector. Compared with the wedge-shaped deflector, the cone-shaped deflector has better performance for deflecting, and increased the distance and optimized the structure of the sidewalls can further improve the exhaust gas flow conductivity. The results provide engineering guidance and theoretical significance for design of the flame deflector of the launch platform.The full version of this extended abstract will appear in Defence Technology in 2020.

Numerical study for the flame deflector design of four-engine liquid rockets

The thermal environment of four-engine liquid rocket exhaust plume impinging on the flame deflector with different impingement and uplift angles is analysed. The supersonic exhaust gas impinging model was established by using the compressible, multi-component, Reynolds-Averaged Navier-Stokes (RANS) equations with the finite volume method. A comparison between the numerical results and experimental data in the literature is made, which verified the validity and accuracy of the numerical model. Additionally, the flow fields of the four-engine rocket impinging on the flame deflector under different impingement and uplift angles are simulated. The results show that high temperatures on the deflector surface mainly occur on the impingement point or the cambered surface. A large impingement angle causes the reverse flow intensity to increase whilst a small angle causes the exhaust gas to deflect a little, a suitable uplift angle can smoothly guide the exhaust gas away from the deflector that the best thermal environment of the deflector channel appears at an impingement angle of 25°and an uplift angle of 5°. This study demonstrates that our model can effectively simulate the impinging flow field, and can be of use for the design of the flame deflector under the multi-engine rocket exhaust gas impact.

A study on the performance characteristics of two-stream supersonic diffusers

This paper investigates the mixing of two supersonic streams in a supersonic exhaust diffuser using theoretical, numerical, and experimental approaches. It has focused attention on the startup and stationary operation of the two-stream supersonic diffusers. The diffuser geometry consists of a constant area throat which connects two ducts of convergent and divergent cross-sections. The effects of diffuser contraction ratio (CR) and convergence angle (θ) on the starting process, pressure recovery, and flow features. Experiments have been performed at five different secondary Mach number (Ms) values in the range of 1.8 <Ms < 2.6, at a fixed primary Mach number (Mp) value of 2. Synchronized pressure measurements are used to study the diffuser performance characteristics. Numerical simulations were also carried out to obtain a better insight into the flow field and computed results have been compared with the experimental results. It was found that the starting behavior has a hysteresis. As a result, the tunnel can be operated at lower pressure ratio than the starting pressure ratio. It is observed that the settling chamber pressure required to start a Constant-Area Diffuser (CAD) is decreased with increase in the diffuser length up to an L/D ratio of 9; subsequently, starting pressure remains almost constant with the further increase in the length due to frictional and pressure recovery losses. The present study provides valuable insights into the multi-stream supersonic mixing and diffusion problem.

On the subsonic near-wake of a space launcher configuration with exhaust jet

The Ariane 5 failure flight 157 made clear that the loads in the base region of space launcher configurations were underestimated and its near-wake dynamics required more attention. In the recent years, many studies have been published on buffet/buffeting in the critical high subsonic flow regime. Nevertheless, not much experimental data are available on the interaction of the ambient flow with an exhaust jet over a wide subsonic Mach number range. Further, a preceding study without exhaust jet revealed questions regarding a similar distribution of the velocity and Reynolds stress in the near-wake if scaled with the reattachment length. Consequently, a generic space launcher configuration featuring a cold, supersonic, over-expanded jet is investigated experimentally in the vertical test section Cologne (VMK) by means of particle image velocimetry (PIV) for five subsonic Mach numbers ranging from 0.5 to 0.9 with corresponding Reynolds numbers between Re_{text {D}}=0.8times 10^6 to 1.6times 10^6. The velocity and Reynolds stress distribution are provided for the near-wake flow and additionally for the incoming boundary layer. Just as in the preceding study, self-similar features are found in the flow field as long as the separated shear layer reattaches on the solid nozzle wall. Substantial changes are then measured for an alternating (hybrid) reattachment between the solid nozzle wall and supersonic exhaust jet as found for Mach 0.8, one of them being the increased axial turbulence in the recirculation bubble due to a ‘dancing’ large-scale, clockwise-rotating vortex.Graphic abstract

A wavenumber subinterval grouping strategy compatible with non-gray metal wall boundaries used in multi-scale multi-group full spectrum k-distribution gas radiation model

In the present work, strategies for the grouping of wavenumber subintervals were extended to deal with the compatibility of multi-scale multi-group full-spectrum k-distribution model with non-gray wall boundaries. The main idea is to enhance the correlated-k characteristics of absorption spectra of H2O-CO2 mixtures and emission spectra of non-gray wall within each divided group. The improvements in the calculation accuracies of radiative heat flux on the wall resulting from the new grouping strategy were evaluated using a series of 1D cases, in which combustion gases with strong inhomogeneities of temperature, pressure and molar ratio of water vapor to carbon dioxide were bounded by non-gray walls. The non-gray walls are assumed to be different metal materials with various surface states and temperatures. Finally, evaluation of the wavenumber subinterval grouping strategy was presented based on calculation of thermal imaging for directional radiation transfer from hot supersonic exhaust plume to a metal plate.

Design and analysis of a cooling system for a supersonic exhaust diffuser

High-altitude test facilities are usually used to evaluate the performance of space mission engines. The supersonic exhaust diffuser, a main part of high-altitude test facility, provides the required test cell vacuum conditions by self-pumping the nozzle exhaust gases to the atmosphere. However, the plume temperature is often much higher than the temperature the diffuser structure is able to withstand, usually above 2500 K. In this study, an efficient cooling system is designed and analyzed to resolve the thermal problem. A water spray cooling technique is preferred among various existing techniques. Here, a new algorithm is developed for a spray cooling system for a supersonic exhaust diffuser. This algorithm uses a series of experimental and geometrical relationships to resize the governing parameters and remove the required heat flux from the diffuser surface. The efficiency of the newly designed cooling system is evaluated via numerical simulations. The utilized numerical technique is based on the discrete-phase method. Various computational studies are accomplished to enhance the accuracy of numerical prediction and validation. The present numerical study is validated using experimental results. The results show that the realizable k-ɛ method is superior compared to other Reynolds-averaged Navier–Stokes models.