Rocket Engine Performance Research Articles

Liquid Rocket Engine Performance Characterization Using Computational Modeling: Preliminary Analysis and Validation

The need to refuel future missions to Mars and the Moon via in situ resource utilization (ISRU) requires the development of LOX/LCH4 engines, which are complex and expensive to develop and improve. This paper discusses how the use of digital engineering—specifically physics-based modeling (PBM)—can aid in developing, testing, and validating a LOX/LCH4 engine. The model, which focuses on propulsion performance and heat transfer through the engine walls, was created using Siemens’ STAR-CCM+ CFD tool. Key features of the model include Eulerian multiphase physics (EMP), complex chemistry (CC) using the eddy dissipation concept (EDC), and segregated solid energy (SSE) for heat transfer. A comparison between the complete GRI 3.0 and Lu’s reduced combustion mechanisms was performed, with Lu’s mechanism being chosen for its cost-effectiveness and similar output to the GRI mechanism. The model’s geometry represents 1/8th of the engine’s volume, with a symmetric rotational boundary. The performance of this engine was investigated using NASA’s chemical equilibrium analysis (CEA) and STAR-CCM+ simulations, focusing on thrust levels of 125 lbf and 500 lbf. Discrepancies between theoretical predictions and simulations ranged from 1.4% to 28.5%, largely due to differences in modeling assumptions. While NASA CEA has a zero-dimensional, steady-state approach based on idealized conditions, STAR-CCM+ accounts for real-world factors such as multiphase flow, turbulence, and heat loss. For the 125 lbf case, a 9.2% deviation in combustion chamber temperature and a 15.0% difference in thrust were noted, with simulations yielding 113.48 lbf compared to the CEA’s 133.52 lbf. In the 500 lbf case, thrust reached 488 lbf, showing a 2.4% deviation from the design target and an 8.6% increase over CEA predictions. Temperature and pressure deviations were also observed, with the highest engine wall temperature at the nozzle throat. Monte Carlo simulations revealed that substituting LNG for LCH4 affects combustion dynamics. The findings emphasize the need for advanced modeling approaches to enhance the prediction accuracy of rocket engine performance, aiding in the development of digital twins for the CROME.

Numerical investigation of liquid oxygen filling process in liquid rocket engine by a multi-dimensional co-simulation method

The liquid oxygen filling process in the thermal component downstream of the valve directly affects the starting performance and working reliability of the liquid rocket engine. The filling process is reproduced with a multi-dimensional co-simulation method including one-dimensional system simulation and three-dimensional component simulation, the former provides dynamic flow data for the latter to calculate the filling process. The characteristics of the flow and evaporation and the phenomenon of gaseous oxygen residue are investigated. The results indicate that the simulated pressure value and tendency in the thermal component agree well with the experiment. The flow and evaporation reach a temporary equilibrium due to the flash boiling of the superheated liquid oxygen in the thermal component. The pressure in the combustion chamber downstream of the thermal component plays an important role in breaking the equilibrium and promoting the transition from superheat to supercool state of liquid oxygen. Additionally, the volume proportion of residue gaseous oxygen is 35 % at ignition time, which is far from the ideal state and affects the temperature measurement in the experiment. 26 % of the gaseous oxygen has not been discharged in time, leading to the injection instability in the injection region away from the inlet.

Mass minimization approach for the optimal preliminary design of CMC inner liners in rocket thrust chambers

In the past decade, the world has witnessed a new space race, driven by a growing commitment to reducing the environmental impact of space missions. This has led to the widespread adoption of liquid-propellant rocket engines, which offer several advantages over their solid-propellant counterparts. One key advantage is their reusability, which not only helps to reduce the generation of space debris but also makes space exploration cheaper. To further enhance the performance of liquid rocket engines, researchers have been exploring innovative cooling techniques and advanced materials. Among these materials, Ceramic Matrix Composites (CMCs) have shown great potential in reducing the overall engine weight when used instead of high-tech metal alloys, resulting in lower fuel consumption and emissions during launches. This paper focuses on the mass minimization of inner liners made of CMCs in rocket thrust chambers. At this aim, a computationally efficient preliminary design approach, based on an analytical one-dimensional thermo-mechanical model, is proposed. A case study of mass minimization of an inner liner of rocket thrust chamber is also presented and discussed, by considering five different CMC materials.

A thrust vector control test system of large range solid rocket engine utilizing a four-point support piezoelectric dynamometer and a deflected flange structure

The accurate measurement of thrust vector is very important to improve the life and safety of aircraft engine. In order to meet the requirements of thrust vector test of a solid rocket engine and overcome the contradiction between large thrust (>50 kN) and high-frequency response (≫30 Hz)/high precision measurement (<3%), a thrust vector control test system utilizing a four-point support piezoelectric dynamometer (FSPD) and a deflected flange structure was proposed. The test system mainly consists of a four-point support piezoelectric dynamometer and some vector force simulation loading devices. The theoretical mechanical model and the thrust vector control test system was set up to study the performance of solid rocket engine. Firstly, the piezoelectric elements and four-point support piezoelectric dynamometer were calibrated by the three-direction force calibration loading device. Secondly, the rationality and feasibility of the thrust vector control test system were proved by the simulated loading experiment. The research results show that the static and dynamic performance of the test system meet the test requirements, and it has good linearity and repeatability, errors are less than 1.00 %, and interphase interference is less than 2.00 %. The measuring system proposed in this paper provides a new idea for the accurate measurement of vector force with large load and high-frequency response.

Hybrid Rocket Engine Burnback Simulations Using Implicit Geometry Descriptions

The performance of hybrid rocket engines is significantly influenced by the fuel geometry. Burnback simulations, to determine the fuel surface and fluid volume, are therefore an important tool for preliminary design. This work presents a method for the simulation of spatially constant burn-ups on arbitrary geometries. An implicit surface definition by means of a signed distance function is used to represent the fluid volume and the fuel block on tetrahedral meshes. Two methods each are used to determine the fluid volume and the burning surface. The first method is based on a direct integration of the signed distance function with the Heaviside function or the Dirac delta distribution, respectively. The second method linearly interpolates the position of an isosurface and thus reconstructs the fuel surface. Both methods are compared and validated with analytical results of four example geometries. Both calculations of the fluid volume and the calculation of the surface content with the interpolation method are characterized as first-order methods. With practicable mesh resolutions of one million computational cells, errors below two percent can be achieved. With the interpolation method, numerical meshes can also be exported for any time points of the burn. Finally, the application of the program to the fuel geometry of the Viserion hybrid rocket engine is demonstrated.

Development and Numerical validation of an Aerospike nozzle Contour Design

Aerospike nozzle simplified design and analysis, as well as test results, are presented. New techniques in nozzle design have been used to improve the performance of current rocket engines. The use of an aerospike nozzle is one such technique. It outperforms the existing proven bell nozzle in terms of overall performance. This study analyses the development and flow analysis of an aerospike nozzle at sea level, validating their results by taking into account currently used bell nozzles. The flow over the aerospike pump is investigated in higher depth using computational fluid dynamics under an autonomous fire testing condition. This test showed that the approximate design method used to determine an aerospike nozzle shape can lead to extremely effective nozzles. A model of the current 7.45 kN bell nozzle was selected as a specimen to develop an aerospike Nozzle. Gambit and Fluent software were used to model and examine the flow characteristics of building the aerospike nozzle. The developed aerospike nozzle shows a nearly 2% increase in performance over the existing nozzle.

Experimental and numerical study on liquid film cooling performance under heated wall condition

Liquid film cooling is broadly applied in the thermal protection of liquid rocket engine. An experimental system of liquid film cooling for a heated curved stainless-steel wall is built in this study to investigate the effects of jet mass flow rate, jet angle, and nozzle diameter on the cooling performance. The temperature distribution of the heated wall is demonstrated with infrared thermography technology and the flow spreading shape of the liquid film is captured with high-speed camera. A numerical liquid film cooling model based on the volume of fluid (VOF) coupled Level-set method is established to calculate the evaporative heat absorption and spreading shape of liquid film. The experimental study indicates that the peak film thickness in the hydraulic jump region increases with the jet mass flow rate, such as the peak film thickness increases from 265.15 to 632.25 μm at Q = 100–600 mL·min−1. Besides, a jet angle of 35 degree can maintain the maximum wall temperature drop of 29.4 °C. In addition, the simulated peak film thickness presents a minimum deviation of 5.2 % from the experimental data. The numerical model provides important reference for the study of liquid film cooling performance of the liquid rocket engine.

Design and analysis of a supersonic diffuser for altitude testing of upper-stage liquid rocket engines

This study conducted a numerical simulation of high-altitude environments for upper-stage liquid rocket engines that use second-throat supersonic diffusers. To obtain the minimum second-throat area, a typical normal shock model was modified and validated using two experimental procedures in an existing high-altitude facility: a cold flow test with a conical nozzle, and a hot test of a 7-ton class liquid engine. The modified normal shock model, known as the passive diffuser model, was employed to design a new second-throat supersonic diffuser and to numerically test it under various conditions. In the numerical analysis, a velocity parameter, such as the Mach number, should be maintained constant with minimal v-velocity at the entrance of the diffuser contraction to realize a good agreement with the quasi-one-dimensional estimation. In the full-scale facility, geometrical parameters of the second-throat supersonic diffuser normally have a marginal effect on the performance of the liquid rocket engine under test, unless they affect the starting condition of the second-throat supersonic diffuser. It is possible to reduce the throttling of the new upper-stage liquid rocket engine to 63 % by utilizing the self-evacuation feature of the newly designed second-throat supersonic diffuser. We find that, as throttling decreases, the hot region shifts to the second throat, necessitating the use of a cooling system to protect the entire second-throat supersonic diffuser from thermal damage.

Effect of water injection on the cooling performance of flame deflector in rocket engine testing platform

Ground test is an essential step to ensure the reliability of the design and assess the quantity and performance of rocket engine. However, supersonic and high-temperature rocket exhaust gas directly impinging on the flame deflector of test platform cause severe heat damage which brings about a low repeated utilization and construction cost of the deflector. Generally, the water is employed as the coolant for the spray cooling system to protect the flame defector of test platform. In this paper, Euler-Euler approach is adopted to investigate the multi-phase flow in a flame deflection of rocket engine test platform. An evaporation model based on concentration difference is chosen to simulate the mass and heat transfer between water injection and exhaust plume. Model validation shows that the predicted results are in good agreement with the experimental data. The effect of water injection, including spray location and angle, and mass flow rate, on the afterburning and cooling performance of flame deflector is numerically investigated. The results show that the afterburning of exhaust plume could be inhibited by the water injection. In terms of maximum and area-averaged temperature on the bottom and inclined surfaces of flame deflector, the configuration with spray nozzle angle of 60° has the lowest area-averaged temperature. For the simulated cases with different spray nozzle diameters, the difference of maximum and area-averaged temperatures is below 11.0 K and 6.1 K, respectively. In addition, the maximum and area-averaged temperatures increase first and then decrease with the increase of spray nozzle diameter.

Performance Evaluation of Hybrid Propellant Rocket Engine

In this article, the potential of the hybrid propellant rocket engine technique is discussed. it is a great dream to flight engineers to design and fabricate a rocket engine with more reliability and nil rate of combustion instability probably for long range applications. Solid propellant rocket engine and liquid propellant rocket engine has a considerable rate of advantages but combustion instability is a great threaten in both the engines. Hybrid propellant rocket engine has a very less rate of combustion instabilities when compared to both solid and liquid propellant engines. The ultimate aim of every engine design is to prevent occurrence of this combustion instability and maintain reliable operation during entire span of the combustion process. This article is mainly focused on occurrence of various type of combustion instabilities in hybrid propellant rocket engine and its preventing methods. The performance of the hybrid propellant rocket engine was analysed under various aspects and its unique characteristics were presented in detail.

Effect of dielectric barrier discharge methane reforming products on the combustion performance of rocket engine

Although the low-thrust liquid oxygen/methane rocket engine has broad application prospects, the low flame propagation speed and low combustion rate of methane fuel make the liquid oxygen/methane engine still face key technical challenges. Methane fuel is partially converted into hydrogen and ethane with higher combustion rate before being injected into the combustion chamber, which is positive for the use of dielectric barrier to improve the combustion performance of the engine. Therefore, this paper studies the effect of the main four products of dielectric barrier discharge reforming with methane conversion rate of 10%, on the flow field of the combustion chamber. The results show that the addition of reforming products can effectively improve the combustion efficiency of the engine. H2 in the reforming product can also improve the specific impulse performance of the engine by increasing the total pressure of the engine chamber. C2H4 will not affect the maximum temperature of the engine, However, it can expand the medium-high temperature range of engine temperature to different degrees. The addition of H2 accelerates the oxygen consumption rate, which provides a feasible way to reduce the design size of the engine and improve the combustion efficiency of the low-thrust engine.

Analysis on Propulsive Performance of Hollow Rotating Detonation Engine with Laval Nozzle

In the present study, an experimental performance analysis of hollow rotating detonation engines (RDEs) with Laval nozzles is carried out for the first time. Experiments of a hollow rotating detonation engine with a Laval nozzle were performed with a modular RDE at a backpressure condition of 1 atm. Two configurations with area ratios of the outlet throat to the inlet of and 2.7 have been tested with gaseous methane/oxygen as propellants. Three normalized metrics, usually used for evaluating the performance of conventional rocket engines, are introduced to analyze the performance deficit between the measured value of an RDE and the ideal value of an isobaric-combustion-based engine. These metrics allow for assessing the entire engine and each component separately. The metric analysis suggests a small outlet-to-inlet area ratio () is detrimental to the propulsive performance. To explain the mechanism, a gas-stratification flowfield model is further proposed. It is found that the unchoked region in the combustible gas layer, which is caused by unchoked injection on the injecting plate, is responsible for the performance deficit of the combustion chamber. This model is then validated by one-dimensional numerical simulations and experimental data. In addition, we also focus on the global performance, including the gross thrust, the specific impulse, and the utilization of the supplied stagnation pressure. The result implies a tradeoff space when choosing an appropriate .

Performance of Additively Manufactured Fuels for Hybrid Rockets

Hybrid rocket engine (HRE) performance is dependent on fuel/oxidizer selection and fuel grain geometry. A literature review was performed to identify key trends and findings related to the application of the additive manufacturing (AM) of fuel systems for HREs. The effects of complex combustion port geometries, embedded structures, and end-burning systems, along with the use of metallic additives, turbulators, diaphragms, gel-like fuels, powdered fuels, liquid fuels, and liquifying fuels and their impact on regression rates, combustion efficiencies, and/or mechanical strength are thoroughly documented here. In general, the application of AM to HRE fuels can be implemented to increase regression rates and combustion efficiency, and tailor HRE designs. Chemical equilibrium analysis computations were completed to characterize the theoretical performance of HTPB and common AM fuels (ABS, PLA, PC, PMMA, Nylon 6, and a UV-based fuel) with common oxidizers (LOX and N2O). AM fuels exhibit a similar theoretical performance as the commonly used HTPB fuel, and proper selection of the fuel can yield improved performance and design metrics. Development of AM approaches for HRE fuel design have significantly expanded their design trade space and should enable the competitive application of HREs for future propulsion missions.

Studies of Jet A1 Fuel Atomization Through Non Circular Orifices

&#x0D; &#x0D; &#x0D; The performance of the liquid rocket engine depends on the atomization behavior of the fluid being injected into the combustion chamber. Generally, a plain injector with a circular orifice has been used in the injector, but it has the disadvantage of having a low spray cone angle. The breakup length, mean droplet diameter, and Sauter mean diameter is also higher. Thus, to overcome these drawbacks, non-circular orifices have been utilized in the present study. The shapes used for non-circular orifices are semi-circular and plus. The results obtained with the non-circular orifice is compared with the circular orifices of the same area ratio. The working fluid used for the studies is Jet A1 fuel. Studies were also conducted with different L/D ratios by choosing an effective orifice length to reduce the upstream losses. The axis-switching phenomena were observed with the semi-circular as well as with the plus jets. The mean droplet size of the circular jets was more prominent compared to non-circular jets, and the Sauter mean diameter of non-circular jets droplets was smaller than that of the circular jet droplet. The spray cone angle has increased by 290% for plus jets and 30% for semi-circular jets compared to circular jets.&#x0D; &#x0D; &#x0D;

Charactering Catalytic Ignition and Combustion for the Ninety-Percent Hydrogen-Peroxide/Paraffin-Based Hybrid Rocket Engine

The present work performed a combined experimental and numerical investigation of the catalytic ignition and combustion characteristics for a hybrid rocket engine (HRE) using 90% hydrogen peroxide () with a paraffin-based fuel as the propellant system. The variation of the decomposition temperature used as the analytical premise was monitored by a thermocouple and laser absorption spectroscopy based on water absorption. The effects of preheating, the number of silver mesh layers, and the usage count for the catalytic bed on the decomposition characteristics were analyzed. A preheated 60-layer silver catalyst bed provided the best decomposition performance, including the shortest catalytic decomposition delay time and the highest decomposition efficiency. A two-dimensional numerical model was built to determine the effect of decomposition efficiency decay on the HRE combustion characteristics. Variations in the combustion temperature, combustion chamber pressure, and product distribution were compared for decomposition efficiencies ranging from 70 to 100%. As the decomposition efficiency of decreased to 80%, part of the oxidizer was blown out the chamber before participating in the combustion, yielding a significant decline in the combustion performance of the hybrid rocket engine. Static firing tests were also used to determine the catalytic ignition and reignition performance of the HRE. The results showed that the performance of the catalytic bed was stable while the reignition process was rapid without a pressure peak.

THERMAL PROTECTION SOLUTION FOR SOLID FUEL ROCKET ENGINES

This paper presents a new approach to the thermal protection of solid-fuel rocket engines. The method of thermal protection using coatings not only reduces the risk of damage to the engine casing but also increases the engine’s efficiency in working. Parameters such as thrust and the mass per square meter of the engine casing are used to compare the differences between engines with and without thermal protection coating. Numerical calculation applied to the cruise engine of the rocket 9M39. Calculation results show that the thickness of the thermal protective coating and the thickness of the combustion chamber have been optimized regarding rocket weight. The survey results showed that the mass per square meter of the engine casing decreased by 2.5 times, and the engine’s impulse total increased by 1.13% when using the thermal protective coating. These results are used as data to develop methods to improve the performance of solid propellant rocket engines.

Analysis of design parameters of fuel injector used in liquid propellant rocket engine

Performance of the liquid rocket engine is considered to be significant not only in terms of type of the fuel, fuel tank and structure used but also flow rate of the fuel takes prime importance in enhancing the engine performance. Variation in flow rate normally altered to greater extent by varying the design parameters of the fuel injectors. Keeping this in mind, the present study aims to alter the design parameters of pintle injector by changing the dimension such as diameter, impinging angle and L-open (fuel exit) by introducing orifice plate outlet of fuel injector to decrease the area and increase the velocity of fuel through designing CATIA model for fuel injector and to perform simulation using ANSYS workbench. The prime objective of the present study is to increase fuel flow rate and make it better comparable to the existing one. Implementing new design of pintle fuel injector by referring the standard value and study the fuel behaviour in the pintle injector. The present work also aims to improve the empirical design tools to get high performance of an engine and to achieve the better stability of the fuel injector through parametric design and analysis through CFD simulation.

Effect of Secondary Combustion on Thrust Regulation of Gas Generator Cycle Rocket Engine

Thrust regulation is applied to maintain the performance of the liquid propellant rocket engine. The thrust level of a rocket engine can be readily controlled by adjusting the number of propellants introduced into the combustion chamber. In this study, a gas generator design is proposed in which thrust regulation is maintained by performing secondary combustion in the divergent section of the nozzle of a gas generator. Tangential and normal injection techniques have also been studied for better combustion analyses. A normal injection technique is used for the experiment and CFD results are validated with the experimental data. Chemical equilibrium analyses are also performed by minimizing Gibbs free energy with the steepest descent method augmented by the Nelder–Mead algorithm. These equilibrium calculations give the combustion species as obtained through the CFD results. Performance evaluation of the rocket engine, with and without secondary combustion in the gas generator, led to an increase of 42% thrust and 46.15% of specific impulse with secondary combustion in the gas generator.

Comparison of Optimization Techniques and Objective Functions Using Gas Generator and Staged Combustion LPRE Cycles

This paper compares various optimization techniques and objective functions to obtain optimum rocket engine performances. This research proposes a modular optimization framework that provides an optimum design for Gas Generator (GG) and Staged Combustion (SC) Liquid Propellant Rocket Engines. This process calculates the ideal rocket engine performance by applying seven different optimization techniques: Simulated Annealing (SA), Nelder Mead (NM), Cuckoo Search Algorithm (CSA), Particle Swarm Optimization (PSO), Pigeon-Inspired Optimization (PIO), Genetic Algorithm (GA) and a novel hybrid GA-PSO technique named GA-Swarm. This new technique combines the superior search capability of GA with the efficient constraint matching capability of PSO. This research also compares objective functions to determine the most suitable function for GG and SC cycle rocket engines. Three single objective functions are used to minimize the Gross Lift-Off Weight and to maximize Specific Impulse and the Thrust-to-Weight ratio. A fourth multiobjective function is used to simultaneously maximize both Specific Impulse and Thrust-to-Weight ratio. This framework is validated against a pump-fed rocket, and results are within 1% of the actual rocket engine mass. The results of this research indicate that PSO and GA-Swarm produce optimum results for all objective functions. Finally, the most suitable objective function to use while comparing these two cycles is the Gross Lift-Off Weight.

Air core vortexing in liquid draining tanks: Influence of surface roughness

This paper study the influence of roughness of the tank base in the formation of air core vortex when liquid inside the cylindrical tank is rotated and subsequently drained. The vortex air core formation in drain tanks can lead to the blockage of drain port and can result in reduced discharge of draining liquid as an immediate consequence. The formation of such a vortex in propellant tanks of spacecraft and rockets will lead to underutilization of propellants and can adversely affect the performance of rocket engines. Ingestion of air core vortex in metal casting process can affect the mechanical properties of casted metal. Hence, suppression of air core vortex is inevitable in the fields of aerospace, metal casting and other hydraulic engineering systems. Current study makes use of roughness of the tank base to suppress vortexing which is the first of its kind and therefore, novel. Experimental investigations have been carried out at several values of roughness height and initial fluid rotation to study the characteristics of air core vortex. It is found that at the highest value of initial rotation provided to the liquid, air core vortexing can be subdued up to 25% by making use of base roughness. Current study also recommends drain tank manufacturers not to smoothen the tank base surface spending time at extra cost. In addition, applying roughness on the tank base will further weaken the formation of air core vortex, in the light of this study.