Thrust Chamber Research Articles

Transcritical Behavior of Methane in the Cooling Jacket of a Liquid-Oxygen/Liquid-Methane Rocket-Engine Demonstrator

The successful design of a liquid rocket engine is strictly linked to the development of efficient cooling systems, able to dissipate huge thermal loads coming from the combustion in the thrust chamber. Generally, cooling architectures are based on regenerative strategies, adopting fuels as coolants; and on cooling jackets, including several narrow axial channels allocated around the thrust chambers. Moreover, since cryogenic fuels are used, as in the case of oxygen/methane-based liquid rocket engines, the refrigerant is injected in liquid phase at supercritical pressure conditions and heated by the thermal load coming from the combustion chamber, which tends to experience transcritical conditions until behaving as a supercritical vapor before exiting the cooling jacket. The comprehension of fluid behavior inside the cooling jackets of liquid-oxygen/methane rocket engines as a function of different operative conditions represents not only a current topic but a critical issue for the development of future propulsion systems. Hence, the current manuscript discusses the results concerning the cooling jacket equipping the liquid-oxygen/liquid-methane demonstrator, designed and manufactured within the scope of HYPROB-NEW Italian Project. In particular, numerical results considering the nominal operating conditions and the influence of variables, such as the inlet temperature and pressure values of refrigerant as well as mass-flow rate, are shown to discuss the fluid transcritical behavior inside the cooling channels and give indications on the numerical methodologies, supporting the design of liquid-oxygen/liquid-methane rocket-engine cooling systems. Validation has been accomplished by means of experimental results obtained through a specific test article, provided with a cooling channel, characterized by dimensions representative of HYPROB DEMO-0A regenerative combustion chamber.

Modeling liquid rocket engine coolant flow and heat transfer in high roughness channels

The design of thrust chambers of modern liquid rocket engines is moving towards the use of additive layer manufacturing techniques, which are replacing traditional techniques because of its inherent advantages of customization and flexibility of design and prototyping phase that in the end leads to advantages in terms of production costs and times. In this framework, the realization of suitable cooling channels, which are characterized by tiny cross section, can lead to high relative surface roughness, as a result of the manufacturing process. Despite the increased friction loss, high roughness channels are considered often with interest because of the corresponding increase of heat transfer and thus cooling efficiency. However, the correlation between the increase of friction and heat transfer changes at high roughness. As a consequence, the commonly used assumption for heat transfer prediction at low roughness may lead to order of magnitude errors if extended to high roughness channels. To respond to the request of a reliable heat transfer prediction model, the correction introduced by Aupoix to the Spalart-Allmaras turbulence model to take into account for high roughness in external flows is extended and validated here for the application to high roughness channels. The implementation of the proposed correction allows to predict realistic values of the convective heat transfer coefficient, in agreement with the correlations reported in the literature.

Unsteady Stream-Tube Model for Pulse Performance of Bipropellant Thrusters

We propose a theoretical model for predicting the impulse bit of bipropellant thrusters under pulse-firing operations. The present theoretical model, which considers the nonuniformity of the mixture ratios created inside the thrust chamber, extends the stream-tube approach, which is limited to the prediction of the steady performance. In pulse-firing operation, the fuel or the oxidizer alone can be injected in isolation due to the mismatched injection timing before the rated injection, which leads to the deterioration of performance as compared with the steady operation. The present approach (the unsteady stream-tube model) successfully implements a time-dependent stream-tube structure inside the thrust chamber, allowing for the prediction of the impulse bit as a straightforward function of the injection conditions. Three different pulse-firing tests using distinct hypergolic propellants demonstrate the validity of this model, typically reproducing the notable trend of deterioration in the impulse bit in the short-pulsed mode. We also examine the time-averaged specific impulse and mass flow rate to improve the impulse bit during short-pulsed operations.

Thermal Analysis of Cooling Channels on Electromechanical Actuator During Actual Flight

Liquid methane is regarded as a potential cryogenic propellant that can be used as coolant for regenerative cooling channels of high-power electrical devices, in addition to the oxygen/methane thrust chamber, owing to its moderate working temperature. This study examines the thermal performance of supercritical methane in cooling channels of the electromechanical actuator of a flight vehicle at varying angles of inclination and flight acceleration overloads. The main effects of overload and attitude at various angles of inclination on heat transfer are shown as variation in buoyancy force, which can be classified into gravitational and centrifugal buoyancy forces. Both forces can disturb the turbulent structure and complicate heat transfer under the influence of the drastically variable properties of the supercritical fluid. The authors also discuss conjugate impacts of the two kinds of buoyancy forces on the thermal performance of the cooling channel at various angles of inclination and flight acceleration overloads. The physical model was constructed as a helically coiled tube with angles of inclination varying from 0 to 90 deg, and flight overload was implemented as and (where is the gravitational acceleration). The analysis shows that both the flight attitude and overload played an essential part in the thermal performance of supercritical methane in the cooling channel. Some novel findings were obtained; an enhancement and a deterioration in local heat transfer could occur at large overloads with inclined orientations. The gravitational buoyancy force leading to secondary flow was an essential factor in the deterioration of local heat transfer but was inconspicuous under the gravity of Earth. Moreover, the trends of development of and were consistent, implying that the buoyancy force generated by gravity and centrifugal forces was not isolated.

Energy Estimation and Testing Verification on Ignition of a Torch Ignition System

The torch type electric ignition system is a desirable choice for reusable liquid vehicle engine. It is generally believed that the ignition and start-up process of engine are crucial and high-risk. However, due to the complexity of combustion experiments, the ignition and start-up process can be efficiently analysed from the perspective of energy exchange. By building the energy analysis equations of electric spark plug, ignition chamber and thrust chamber, the energy estimation method of torch type electric ignition system is established. Furthermore, the time scheduling and examining method for multi-injector ignition and start-up process of large flow liquid rocket engine with multi-injector are established. The hot-firing test of LOX/CH4 rocket engine is carried out according to the design parameters to validate the effectiveness of the energy estimation method.

Assessment of thrust chamber stability margins to high-frequency oscillations based on mathematical modeling of coupled ‘injector – rocket combustion chamber’ dynamic system

High-frequency instability of a liquid-propellant rocket engine (LRE) during static firing tests is often accompanied by a significant increase in dynamic loads on the combustion chamber structure, often leading to a chamber destruction. This dynamic phenomenon can also be extremely dangerous for the dynamic strength of a liquid-propellant rocket engine. The calculation of acoustic combustion product oscillation parameters is important in the design and static firing tests of such rocket engines. The determination of the oscillation parameters (natural frequencies and stability margins such as oscillation decrement) is one of the problems solved in the LRE design period as part of the development of measures to ensure the engine stability. The main aim of the paper is to develop a numerical approach to determining the parameters of acoustic oscillations of combustion products in liquid-propellant rocket engines combustion chambers taking into account the features of combustion space configuration and the variability of gaseous medium physical properties depending on the axial length of the chamber, acoustic impedance in critical throat and dissipation effects (damping experimental values) in the shell structure and the gas media in the chamber. The approach is based on mathematical modeling of the coupled ‘chamber shell structure – gas’ dynamic system by using the finite element method and the CAE (Computer Aided Engineering) system. The developed approach testing and further analysis of the results for the RD 253 engine using nitrogen tetroxide and unsymmetrical dimethylhydrazine as a propellant pair were carried out. The dynamic system shapes and frequencies of longitudinal, tangential and radial modes are determined. The results of mathematical modeling of the dynamic system indicate a satisfactory agreement of the calculated decrements of the first longitudinal oscillation mode and third tangential oscillation mode with the experimental decrements obtained by hot-fire tests data. From system harmonic analysis of the thrust chamber, it follows that the dynamic pressure gain factor of the gas media in the chamber at the first longitudinal mode frequency is 1.6 times greater than the system dynamic gain in the tangential mode. At the same time, the oscillation decrement for the system tangential mode is 2 times smaller than that of the first longitudinal mode. This means that the thrust chamber tangential mode is more dangerous and can lead to rocket engine combustion instability. The effect of the injector on the high-frequency stability of the combustion chamber and the possibility of partial suppression of combustion chamber thermoacoustic oscillations by adjusting the high-frequency dynamics of the injector are shown theoretically.

Thermal–Mechanical FEM Analyses of a Liquid Rocket Engines Thrust Chamber

The Italian Ministry of University and Research (MIUR) funded the HYPROB Program to develop regeneratively cooled liquid rocket engines. In this type of engine, liquid propellant oxygen–methane is used, allowing us to reach very good performances in terms of high vacuum specific impulse and high thrust-to-weight ratio. The present study focused on the HYPROB final ground demonstrator, which will be able to produce a 30 kN thrust in flight conditions. In order to achieve such a thrust level, very high chamber pressures (up to 50 bar) and consequently high thermal fluxes and gradients are expected inside the thrust chamber. Very complex and high-fidelity numerical FEM models were adopted here to accurately simulate the thermal–mechanical behavior of the thrust chamber cooling channels, accounting for plasticity, creep, and low-cycle fatigue (LCF) phenomena. The aim of the current work was to investigate the main failure phenomena that could occur during the thrust chamber’s service life. Results demonstrated that LCF is the main cause of failure. The corresponding number of loading cycles to failure were calculated accordingly.

Thrust Chamber Dynamic and Propulsive Performance of Biofuels Under Detonation Combustion

Detonation is a shock wave obtained from the energy that releases after the combustion. Chapman-Jouguet(CJ) theory can be used to identify the behaviour of the detonation in gasses. Pulse Detonation Engines (PDEs) is one of the engines that implement the detonation in its combustion system. The Humphrey cycle is the thermodynamic cycle which is similar to the Pulse Detonation Engines (PDEs). It is the modification of the Brayton cycle where the constant-pressure heat addition process of the Brayton cycle is replaced by a constant-volume heat addition process. The Humphrey cycle can provide the pressure rise combustion by utilizing the shock inside the combustion chamber. Compared to the Brayton cycle, the Humphrey cycle has higher thermodynamic efficiency. However, the detonation process has unsteady combustion which makes it more difficult to handle. The purpose of the study is to calculate the performance of the aircraft by using alternative fuel in the ZND model. An analytical model is started by having the molecular structure of each biofuel and it has been used to determine detonation velocity, Mach number at C-J point, temperature ratio, pressure ratio, density ratio, Brayton and Humphrey efficiency, specific impulse, and specific thrust. In addition, the physical properties of the flow are investigated by changing the initial temperature, initial pressure, and mass flux. The pressure ratio, temperature ratio, and density ratio will all decrease as the initial pressure varies. The variation of mass flux and initial temperature, on the other hand, generates the opposite outcome as the change of pressure. The feasibility of the fuels in the detonation combustion can be known as they show high propulsive performance after the initial condition is changing.

Impact of shear-coaxial injector hydrodynamics on high-frequency combustion instabilities in a representative cryogenic rocket engine

The excitation mechanism of a thermoacoustic instability in a 42-element research rocket thrust chamber with representative operating conditions with respect to European cryogenic rocket engines is investigated in detail. From previous research it was known that the chamber 1T mode can be excited by persistent heat release rate oscillations which are modulated by the resonant modes of the liquid oxygen injectors. The excitation source of the longitudinal injector eigenmodes is investigated in this study. Fibre-optical probes measuring the OH* dynamics from the recess volume of two injectors showed additional frequency content which could neither be explained by the chamber acoustics, nor the acoustics of the injection system. Instead, the temporal evolution of these frequencies correlate with the oxidizer flow velocity. In this work we show that the additional flame modulation originates from a hydrodynamic effect in the injection system. Even though the exact process cannot be precisely identified, an effect designated orifice whistling at the injector inlet orifice seems to be a likely candidate. Combining the new results with previous publications about this combustor, it is now possible to explain past and present observations in terms of the hydrodynamic and thermoacoustic conditions which are necessary for the combustion instability to appear. The conditions, which lead to an injection-driven excitation of the 1T mode are matching frequencies of the 2L mode of the injectors and the chamber 1T mode as well as a Strouhal number between 0.2 and 0.4 based on the length and flow velocity of the injector inlet orifice.

A Bi-stage Multi-objective Reliability-based Design Optimization Using Surrogate Model for Reusable Thrust Chambers

To reduce the computational burden of Multi-Objective Reliability-Based Design Optimization (MORBDO) and promote its application in complex practical systems, a bi-stage MORBDO procedure using surrogate models is proposed in this paper. The novel process consists of two stages: in the first stage, a narrowed solution space named as MORBDO design space is obtained by a pre-multi-objective deterministic design optimization in the whole original design space; In the second stage, the MORBDO design space is extended to form an augmented space by considering uncertainties, and then the MORBDO is executed with the double-loop strategy over the augmented space. Besides, the multiple response surface (MRS) is adopted in the bi-stage MORBDO procedure to replace the time-consuming finite element analysis. The proposed method is validated by conducting MORBDO on a lab-scale reusable LOX/H2 thrust chamber aiming at maximum residual strength and cyclic life. Compared with the traditional MORBDO implemented in the whole original design space, the bi-stage MORBDO can significantly reduce the computational burden of the surrogate model by 49.33% and only one-quarter of the optimization population is required in NSGA II to search for the optimal solution. Finally, a sensitivity analysis is performed to quantify the importance ranking of design variables of thrust chambers. The results indicate that the geometry of cooling channels near the throat plays the most important role in the thrust chamber's residual strength and cyclic life. The proposed method is easy to be conducted and holds great potentials in the MORBDO of other complex practical systems.

Performance Evaluation of LowThrust Rocket Nozzles of Various Geometries

Issues of viscous loss in rocket nozzle become increasingly more important as the thrust chamber is reduced in size. A properly designed nozzle is critical since small geometrical differences can result in dramatically modified performance. The present study analyzes the performance of conical converging-diverging nozzle for low-thrust rocket application. A total of five nozzles with different area ratio and divergence length were numerically tested; two optimum area ratio nozzles, one underexpanding nozzle and two overexpanding nozzles. The main aim is to analyze the flow phenomena of compressible gas flows within these nozzles and its relation to the rocket performance (i.e. thrust). The axisymmetric flow problems are numerically solved using a well-validated software with the Spalart-Allmaras turbulence model is employed to model the effect of turbulent on the flow. The results revealed technique of eliminating flow separation and obviously, knowledge of the point of separation is essential for performance enhancement. The elimination of flow separation inside the overexpanding nozzle by means of truncating it at a point of separation increases the thrust produced by 18% for a relatively low combustion pressure. It is anticipated that the thrust enhancement would be greater for a higher nozzle pressure ratio (NPR, i.e. the ratio of combustion chamber total pressure and atmospheric static pressure). With the demonstrated performance enhancement, this work serves as a driver for further nozzle enhancement using this technique.

Spray characteristics of the nanoparticle-containing gel propellants by using an improved single-phase nozzle

Due to the high viscosity and complex rheological properties, the atomization problem continues to be one of the key problems of gel propulsion technology. By analyzing the rheological properties of the gel propellants containing nanoparticles, an improved single-phase nozzle was proposed in this manuscript, which added a cone-like structure inside the nozzle. The equivalent cross-sectional area of propellant passage decreases gradually from inlet to outlet, which ensures that the apparent viscosity of gel propellant decreases continuously until it flows out of the nozzle. A series of experiments were conducted to investigate the atomization characteristics of the improved nozzle. A traditional direct current (DC) nozzle was also used for comparison. The results showed that the improved nozzle proposed in this manuscript can effectively improve the atomization fineness and uniformity of the gel propellant containing nanoparticles by efficiently decreasing Sauter mean diameter (SMD) and also reducing the volume fraction of large droplets in the spray field. Meanwhile, the spray cone angle of the improved nozzle was obviously larger than that of the DC nozzle, which was favorable for mixing and evaporation in the thrust chamber.

Recurrent Neural Network based Soft Sensor for flow estimation in Liquid Rocket Engine Injector calibration

Injector is the critical element in the Liquid Rocket Engine (LRE), to ensure proper mixing of propellants (fuel and oxidizer) in the thrust chamber for achieving the optimum thrust. LRE injector is calibrated in order to deliver required flow rates of propellants by sizing the orifices through simple injector water calibration (IWC) techniques. In LRE-IWC process, a huge 6” turbine flow meter (TFM) is employed for the flow-rate measurement. In order to achieve and maintain the required accuracy and precision in the LRE-IWC process, periodical calibration of TFM is mandatory. It involves tremendous time, cost and human effort. Soft sensors can provide an economical and effective alternative solution for TFM flow-rate measurement. The objective of the proposed work is to develop and implement a recurrent neural network based soft sensor (RNN-SS) for TFM flow-rate measurement. In the LRE-IWC process, experimental flow trials were carried out for different flow patterns, and the necessary measurement data were generated for the soft sensor design. The designed RNN-SS was trained by tuning various hyper parameters to replace the TFM, using three related measurement parameters acquired during the experimental trials. The precise TFM flow-rate estimation was achieved by the designed RNN-SS, with a worst-case mean absolute percentage error (MAPE) of 1.91% for the experimental flow patterns considered, with good repeat-ability. The proposed RNN-SS model for TFM flow-rate estimation gives a MAPE of 0.58%, for the required flow-pattern which is well suited for practical use.

Optimal Design of Regenerative Cooling Structure Based on Backpropagation Neural Network

Aiming at the problems of relying on traditional experience and high cost of schemes selection in the current research on regenerative cooling design, this paper establishes an optimal design model based on the backpropagation (BP) neural network method. In this paper, the cooling structure is parameterized to realize the segmented description. A BP neural network is constructed through the chamber performance simulation data sets to get the mathematical description between structural size parameters and optimization objectives. Taking the specified thrust chamber lifetime, safety margin of chamber wall, and upper limit of maximum wall temperature and pressure loss as constraints, the Globalsearch algorithm is applied to obtain the optimal design results of the regenerative cooling structural size parameters. The results show that the trained BP neural network model is able to predict the regenerative cooling performance parameters with the calculation accuracy similar to numerical simulation. The study of the space shuttle main engine–main combustion chamber indicates that under the same thrust chamber operating parameters and coolant inlet parameters, the optimized regenerative cooling channel reduces both the maximum wall temperature and the coolant pressure loss, and improves the overall cooling performance.

Thermo-structural behavior of thermal barrier coatings for thrust chamber applications

The investigation of the failure mechanism is essential for the application of thermal barrier coating in the thrust chamber of liquid rocket engines. In this work, a four-layer thermal barrier coating system in a typical regeneratively-cooled thrust chamber was simulated by finite element method, and the interface morphology between layers was represented by an ideal cosinusoidal curve. A thermo-structural analysis, the boundary conditions of which were calculated by a fluid-thermal coupling method, was carried out to study the influence of the interface roughness and thermally grown oxide on the temperature and stress distributions under thermal cycles. Elasto-plastic deformation and creep behavior were also taken into consideration. The computed results reveal that the roughness parameters including interface amplitude and wavelength as well as the thermally grown oxide thickness have limited effects on the temperature distribution in the thrust chamber wall, but a significant impact on the residual stress distribution in thermal barrier coatings. As these parameters grow, the magnitude of residual stress generally increases, meanwhile, the dangerous regions transform in a complicated way. Much attention should be paid to the interface of thermal barrier coatings in thrust chambers.

System scheme and thermal performance of a third fluid cooled rocket engine

In this study, a novel expander rocket engine cycle that uses a third fluid as the combustor coolant and the turbine driver was discussed. A detailed system scheme with two stage cryogenic heat exchangers of the third fluid cooled (TFC) liquid rocket engine was improved. The design highlighted the phase-change heat transfer model of the regenerative cooling channels (RCC) on the basis of previous experiments. Subsequently, the state parameters distribution for the system was first proposed. In addition, the exergy analysis of cooled Rankine cycle (CRC) loop was investigated. The effects of the third fluid methane mass flow rate (m˙tf) on the exergy efficiency of CRC loop were further analyzed. The results demonstrated that the m˙tf and the design geometric dimension of the RCC can well meet the thermal protection requirements of the thrust chamber. In the exergy analysis process, the exergy loss of the RCC was the largest. The exergy loss ratio of cryogenic heat exchanger Ⅰ and Ⅱ vary greatly due to the difference of the heat transfer performance. With the increase of the m˙tf, the exergy efficiency of CRC loop decreased slightly, while the maximum output power from the heat transferred through the thrust chamber wall (Wmax) increased. The results will be benefit for a better thermal performance of the TFC engine through the optimization of the m˙tf and subassemblies.

Experimental and Numerical Investigations on the Ignition of RBCC Gas-Oxygen/Kerosene Rocket Ejector

The rocket ejector refers to a core component of a rocket-based combined cycle (RBCC) engine. The ignition is of critical significance for rocket ejection. Reliable and stable ignition crucially determines the normal operation of the engine. In this paper, a thrust chamber with coaxial swirl injector for the RBCC rocket ejection was developed and tested. Gas oxygen (GOX) and kerosene acted as propellants. As revealed from the test results, the process of ignition pressurizing comprised four phases. The oxygen prefilling time before ignition slightly impacted the ignition time, whereas it affected the peak pressure of ignition. In a confined range, the peak pressure decreased as the prefilling time was extended. The ignition was simulated by building a numerical model, and the results well complied with the experimentally achieved results. The numerical model is capable of specifically indicating the position of the kernel of fire and the process of flame propagation. The simulation results reveal that the propellant could form a combustible condition within 4 ms. The kernel was 6 mm away from the injector, located at the oxygen and kerosene mixing interface and approaching the upper wall. The above results reflected the vital role of the central recirculation zone formed by the prefilled oxygen. The ignition energy was transported near the injector under the convection effect, which ignited the stoichiometric mixture, and the entire ignition could reach a stable state within 20 ms. The numerical model which was developed in this paper can help clarify the combustion mechanism.

Investigating the effective stress function algorithm for the thermostructural analysis and design of liquid rocket engines thrust chambers

The thermostructural behavior of a regeneratively cooled thrust chamber for aerospace applications is strongly non-linear as plastic and creep phenomena come into play. Indeed, the fuel combustion inside the thrust chamber generates high heat fluxes and, then, elevated temperatures are expected in the inner copper structure of the chamber which is made of cooling channels separating the hot gases from the coolant flow. Since the analysis of those non-linear phenomena requires the adoption of very complex and heavy numerical models, it is crucial to employ very efficient numerical algorithms such as the Effective Stress Function algorithm; such an algorithm allows to obtain accurate results in terms of stresses in finite element thermostructural analyses taking into account both plastic and creep phenomena. The aim of the current paper is, then, to demonstrate that the Effective Stress Function algorithm is well suited for the analysis of the thrust chamber providing accurate results in a reasonable amount of computation time. The numerical results obtained are very similar to those achieved with the classical approach adopted by finite element commercial codes with significant computation time savings.

Analysis of a transpiration cooled LOX/CH[formula omitted] rocket thrust chamber

Transpiration cooling has been theoretically identified as a more efficient thermal protection system, compared to regenerative and film cooling techniques. The availability of porous light-weight high-temperature materials, such as ceramic matrix composites (CMCs), is making transpiration cooling an attractive solution for use in future re-usable rocket thrust chambers for cheaper access to space. In this study, a numerical framework, utilizing the industry standard computational fluid dynamics (CFD) code ANSYS CFX, for the simulation of transpiration cooled rocket thrust chambers is presented. The framework is shown to be suitable for predicting the performance, both in-terms of thermal protection and thrust produced, of thrust chambers that utilise transpiration cooling systems. A liquid oxygen (LOX) and methane (CH4) oxidiser and fuel combination is used, although the method is not limited to these substances. Foreign gas transpiration is modelled by a multi-component model. Different porous liner geometries and reservoir conditions are studied to demonstrate their effect on the performance of the thrust chamber. The numerical framework was developed to design a sub-scale transpiration cooled rocket thrust chamber ground test.

Computational investigation of cooling effectiveness for film cooled dual-bell exhaust nozzle for LO2/LH2 liquid rocket engines

ABSTRACT A dual bell nozzle (DBN) consists of two bell nozzles of different geometric area ratios attached at the region called inflection. This distinct geometry modification allows the nozzle to adapt to altitudes through its different operating modes. A numerical investigation with hot flow is done on a 2-dimensional axisymmetric model of a DBN having coolant injection at its inflection. Analysis has considered the secondary injection of gaseous hydrogen film into the gas mixture resulting from the combustion of liquid hydrogen and liquid oxygen in the thrust chamber of a LO2/LH2 engine. Model gave a root mean square deviation of 0.0012 from the experimental result at 30 NPR. Flow phenomena inside the nozzle are studied for different altitudes with and without coolant injection. Shift of separation location inside the nozzle on changing the quantity of gaseous hydrogen injected for three nozzle pressure ratios in the range of operation of the nozzle is calculated. The nondimensional shift in separation location is estimated to be 1.34 at 30 NPR, 1.10 at 45 NPR and 0.70 at 60 NPR. Temperature distribution on nozzle wall and cooling effectiveness of coolant on nozzle wall is determined with varying coolant mass flow rate for LO2/LH2 and LO2/RP1 rocket engines and the results are compared. The effectiveness of coolant for LO2/LH2 engine reduces from 1 to 0.78 on moving downstream of the nozzle whereas it reduces from 1 to 0.27 for LO2/RP1 engine. The study also predicts reduction in specific impulse of 5% for LO2/LH2 engine and 3.3% for the LO2/RP1 engine at MR = 15 due to the film cooling.