Gas Generator Cycle Research Articles

Scheme design and assessment of hybrid pump feed system with energy management for throttleable liquid rocket engine

Currently, there is considerable emphasis on the electric pump-fed cycle for liquid engine, primarily due to its design simplicity. However, its development is hindered by the underdeveloped state of power battery technology. Drawing inspiration from hybrid power technology used in electric vehicles and turbochargers, a hybrid pump feed system for throttleable engines is originally proposed as a promising solution. This system integrates the electric motor into the gas generator cycle, with several topologies evaluated. The parallel configuration featuring a mid-motor is selected for its compact structure, efficient power-splitting and energy recovery. Additionally, customized energy management strategies and optimization models are developed to effectively allocate power throughout the operational processes of liquid engines. A comparative analysis of four engine cycles is conducted under the typically variable-thrust mission. The results indicate that attributed to the conservation of turbo-gas and battery energy, the optimized hybrid pump achieves a reduction of 2.39 % compared to the turbopump and 7.15% to the electric pump in total mass. Adaptability assessment further indicates that the mass advantage of the hybrid pump system is more significant during prolonged engine burning and deep throttling. Specific working conditions are found in which the system prefers electric-motor driving or regenerating turbine energy. Although energy-recovery results in the system efficiency decrease, it serves to lower energy demand of battery pack, thus easing the burden on cell thermal management and structural design. This study provides a practical design framework for hybrid pump-fed rocket engines in future variable-thrust missions.

ANALYSIS OF THE CORRECTNESS APPROACHES TO DETERMINING THE SPECIFIC IMPULSE OF THRUST GENERATED BY THE WORKING FLUID OF A GAS GENERATOR CYCLE ROCKET ENGINE TURBINE

For liquid rocket engines (LREs) utilizing the gas generator cycle, the resulting engine thrust is composed of two components: the thrust of the engine chamber and the thrust of the turbine exhaust nozzle. Over the years of LRE utilization, methodologies for calculating ideal thermo gas dynamic parameters of the chamber have been developed, large volumes of statistical data regarding specific impulse losses in engine chambers have been obtained, and approaches for predicting specific impulse and, consequently, thrust of the main chambers have been refined. A much more complex situation arises with determining the specific impulse of the thrust generated by the working fluid of an autonomous turbine: significant limitations in information regarding methods for calculating this parameter have been found during literature analysis, and the available statistics exhibit contradictory characteristics. Accurate prediction of the specific impulse of the thrust from turbine exhaust nozzles allows for a more rational selection of design parameters (e.g., combustion chamber pressure) during the development of a new engine and is particularly important in cases where the spent gas generator gas is used to create control efforts. As the rocket and space industry historically developed under conditions of increased secrecy, many approaches and design methods for LREs remain known only to design bureaus and research institutes that were founded and developed during the heyday of rocket and space technology. This complicates the situation for new, young aerospace companies, as they must either conduct physical tests, which require expensive specific test equipment and increase the time and cost of engine development, or use numerical modeling, which, although to some extent allows for the replacement of physical experimentation, requires model verification and confirmation of the assumptions made. Furthermore, considering the military focus of rocket-related topics in the past, it cannot be excluded that intentionally distorted information may be present in open sources. Thus, considering the global trend of increasing the number of startups in the rocket and space industry, the identification of reliable and open sources of information becomes an important task. In this work, data from open sources regarding the specific impulse of the thrust from turbine exhaust nozzles of existing LREs were analyzed, revealing significant inconsistencies in the declared energy parameters. Numerical modeling of generator gas leakage from the nozzle of the exhaust gas duct of the RD-111 engine turbine has been conducted, and the results were compared with the calculations carried out by other methods. As a result of the work, various approaches to evaluating the thrust of exhaust nozzles during the design of an engine utilizing the gas generator cycle were analyzed, and conclusions were drawn regarding the correctness of their application.

Thermal Performance Characteristics of an 80-Ton Variable-Thrust Liquid Engine for Reusable Launch Rockets

In this paper, an 80-ton thrust liquid rocket engine (hereinafter referred to as an LRE) with a gas generator cycle, a 5:1 thrust throttling ratio, and an integrated flow regulator/gas generator (hereinafter referred to as an IFRGG) is analyzed. This LRE can be used during the first stage of launching, second-stage and upper-stage space missions, and moon/mars low-orbit hovering and soft landing, and it can also be used with various near-space multipurpose flight vehicles. The thermal performance model of the variable-thrust LRE is established, the influence of the main LRE design parameters on the performance optimization is analyzed, and an optimal selection of the design parameters under certain constraints is completed. A performance evaluation was successfully conducted, and we determined the main structural parameters at the sea-level design point; additionally, we evaluated the LRE performance, matched the system parameters under a 100–20% variable-thrust operation, and conducted an analysis on the LRE operating characteristics under a wide range of variable-thrust operations. The LOX/kerosene propellants were selected, and the vacuum-specific impulse of the LRE decreased from 303.2 s to 289.2 s; this followed an approximate linear law, with a decrease of about 4.62% when the thrust was varied in the wide range of 100–20%. The variable-thrust LRE still had a better vacuum performance under a very deep throttling condition. The reason why the specific impulse was low under the deep throttling condition is that it was greatly affected by the different atmospheric pressure that was caused by the varying flight height and the insufficient atomization and combustion of the propellant; however, because of its wide range of variable-thrust working abilities, it is suitable for various special flight missions.

Overspeeding characteristics of turbomachinery for gas generator cycle air turbo ramjet engine

The main interest in the current study focuses on the possibility of overspeeding for the gas-generator cycle air turbo ramjet (GG-ATR) engine. The authors developed the air turbo ramjet engine and investigated its compressor performance. Based on those data, the authors developed the analytical code for the air turbo ramjet engine, which calculates the performances of turbomachinery, gas-generator, and ram combustor. The previous study described that the rotor overspeeding would not occur in the air turbo rocket engine. However, the current results show that degraded ram combustion can decrease the compressor pressure ratio and the compressor power. This reduced compressor power can cause overspeeding for the air turbo ramjet engine. The experimental results of compressor power and turbine inlet pressure support those analytical results.

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.

Low toxic nitromethane based monopropellant for gas generator cycle air turbo ramjet engine

The present study proposes ethanol-nitromethane mixture as a safe, storable, and low toxic monopropellant for a gas generator cycle air turbo ramjet engine and evaluates its propulsion performance. The present study proposes that nitromethane is mixed with ethanol to adjust gas generator combustion temperature. The author developed the computational code for the present analysis and verified its accuracy. The specific thrust of ethanol-nitromethane is nearly identical to that of ethanol-liquified oxygen and is higher than methanol-hydroxylammonium nitrate aqueous solution and hydrazine. Hydrazine has the highest Isp among the propellants in the present analysis. However, Isp of ethanol-nitromethane is nearly equal to that of ethanol-liquified oxygen and higher than those of methanol-hydroxylammonium nitrate aqueous solution, and hydroxyl terminated polybutadiene-ammonium perchlorate. With the propellant-to-air ratio range in the present study, ethanol-nitromethane has the stoichiometric condition at the ram combustor around the propellant-to-air ratio range from 0.2 to 0.25 and can obtain high ram combustion temperature. This result is favorable for ethanol-nitromethane because it improves the specific thrust and the specific impulse. Therefore, ethanol-nitromethane can be a promising low-toxic liquid monopropellant for the air turbo ramjet engine.

Performance Analysis and Mass Estimation of a Small-Sized Liquid Rocket Engine with Electric-Pump Cycle

The propellant supply system of a liquid rocket engine using an electric pump has high reliability because of the relatively small number of components. The system also has the merit of a quick response and ease of control owing to its simple configuration. Recently, the rocket lab developed the Rutherford engine, which has an electric pump cycle, because of the improved technology in the electric motor and battery. This paper examined the development of the electric-pump cycle and compared the performance with other cycles for a small-sized low-thrust rocket engine. Performance analysis and mass estimation were conducted using the developed analysis program, in which reliability in mass estimation was improved based on the designed configuration or real performance data from commercial products. In addition, the modeling method and analysis procedure were described in detail. The results showed that it is possible to develop a small-sized engine with an electric-pump cycle when the present technologies are applied. The electric-pump cycle had a smaller dry mass than the gas-generator cycle, even at a low thrust level of 500 N, and showed higher performance in specific impulse and speed increments.

Effects of the O/F Ratio on the Performance of a Low Thrust LOX/Methane Rocket Engine with an ElecPump-fed Cycle

Recently, electrically driven pump-fed cycle rocket engines (i.e., ElecPump engines) have received considerable attention. ElecPump engines are superior to pressure-fed engines and will likely replace gas generator cycle engines under a 100 kN thrust level. This study focuses on LOX/methane ElecPump engines with a low-thrust level and estimates the nominal operation point of the engine. The objective variable was set to the maximum velocity increment of the stage. The mixture ratio and combustion pressure were explored as the design variables. Thus, the optimal mixture ratio was proposed as the design variable, considering not only the engine weight but also the structural weight.

Index allocation for a reusable LOX/CH4 rocket engine

Reusable rocket engines are the core components of reusable launch vehicles, and have thus become a major focus of aerospace engineering research in recent years. In practice, subsystem design is based on the overall index allocation of an engine; therefore, a multidisciplinary optimization approach is necessary. In this study, design of a reusable methane/liquid oxygen (LOX/CH4) rocket engine with a gas generator cycle was investigated using multidisciplinary optimization. Two parameters were chosen as design variables: pressure and fuel mix ratio of the main combustion chamber. Optimization objectives were specific impulse, structural mass, and life cycle cost of the reusable rocket engine, and constraints were assigned to each discipline according to rocket design requirements. Then, an optimization model was developed, and optimal design parameters were acquired for the LOX/CH4 rocket engine. The proposed method is effective for designing the index allocation of reusable rocket engines and takes into account the multidisciplinary nature of complex systems.

Optimum characteristic length of gas generator for liquid propellant rocket engine

Gas generator is an essential part of gas generator cycle Liquid Propellant Rocket Engine (LPRE). While designing the Gas Generator (GG), the selection of an unoptimized characteristic length result in an increase in mass flow rate of propellants required to drive turbine. As the exhaust of the gas generator is of little importance in the thrust of the engine, therefore, the specific impulse of the engine decreases with increase in the propellant flow rate to the gas generator. In literature, it is witnessed that researchers have conducted investigation of gas generator performance in terms of variation in mixing ratio or some models have been used to predict the characteristic length; however, it lacks the experimental investigation of gas generator performance in terms of characteristic length. In this paper, an experimental investigation of the gas generator of LPRE is conducted to determine the optimum characteristic length of the gas generator with liquid oxygen and kerosene. Hot tests are conducted with two different types of injectors, namely pressure swirl injector and impinging injectors in a modular water cooled gas generator. The gas generator is tested using six different characteristic lengths (i.e., from 1.75 m up to 9.67 m) for pressure swirl injector and with three different characteristic lengths (i.e., from 1.75 m up to 7.15 m) for impinging injector. The experimental results of hot tests show that the highest characteristic velocity of gas generator is achieved at a characteristic length of 4.27 m. It is demonstrated that a higher characteristic velocity of gas generator is achieved using impinging injector as compared to liquid-liquid pressure swirl coaxial injector at mixing ratio of 0.28 at optimum characteristic length. The effects of increase in mixing ratio with decrease in total mass flow rate on the chamber pressure are also discussed.

Alcohol and alkane fuel performance for Gas Generator cycle Air Turbo Ramjet Engine

The present study investigates the feasibility of alcohol fuels for a Gas Generator cycle Air Turbo Ramjet (GG-ATR) Engine in comparison with alkane fuels. The alkane and alcohol fuels in the present study range from C1 species (CH4 and CH3OH) to C4 species (n-C4H10 and 1-C4H9OH).The present study analytically evaluates Specific thrust and Specific impulse (Isp) of GG-ATR engine as functions of gas generator combustion temperature. For both alcohols and alkanes, the peak Gas Generator (GG) combustion temperatures exist for specific thrust and ram combustor temperature. Comparing alcohol to alkane in the lower GG combustion temperature, specific thrusts and Isp of alcohol tend to be higher than those of C2H6, C3H8, and 1-C4H10. GG combustion gases for alcohol fuels contain higher mole fractions of H2O and CO2 than those for alkanes. Because of high mole fractions of H2O and CO2, the ratio of GG combustion gas for alcohols to air flow rate is fortunately close to the stoichiometric ratio, resulting in higher ram combustion temperature. The high ram combustion temperature can contribute to the higher specific thrust and Isp. Lower GG combustion temperature is preferable because of a thermal limit of a turbine blade. Thus, alcohols have superiorities to alkanes for the GG-ATR engine application.

An MPC Approach to Transient Control of Liquid-Propellant Rocket Engines

Abstract The current context of launchers reusability requires the improvement of control algorithms for their liquid-propellant rocket engines. Their transient phases are generally still performed in open loop. In this paper, it is aimed at enhancing the control performance and robustness during the fully continuous phase of the start-up transient of a generic gas-generator cycle. The main control goals concern end-state tracking in terms of combustion-chamber pressure and chambers mixture ratios, as well as the verification of a set of hard operational constraints. A controller based on a nonlinear preprocessor and on linear MPC (Model-Predictive Control) has been synthesised, making use of nonlinear state-space models of the engine. The former generates the full-state reference to be tracked while the latter achieves the aforementioned goals with sufficient accuracy and verifying constraints for the required pressure levels. Robustness considerations are included in the MPC algorithm via an epigraph formulation of the minimax robust optimisation problem, where a finite set of perturbation scenarios is considered.

Performance assessment of electrically driven pump-fed LOX/kerosene cycle rocket engine: Comparison with gas generator cycle

An electrically driven pump-fed cycle for rocket engine is proposed and a viability of the proposed cycle is assessed compared to a gas generator cycle. The maximum possible thrust level is determined considering the technological maturity of the electric motor. Four types of battery cells were assessed in a screening test for the proposed cycle and the necessity of regenerative cooling for the battery pack is shown. The mass expressions of the proposed cycle and gas generator cycle are derived in terms of pump power and burning time. The basic features are demonstrated with respect to combustion chamber pressure, burning time, and thrust level. The results show that it is favorable to maintain a lower combustion chamber pressure, a longer burning time, and a higher thrust level to remedy the payload penalty incurred when the gas generator cycle is not used. In addition to focusing on the battery pack, the regenerative cooling effect on the battery pack mass is discussed. Further, the impact of optimal battery cell discharge time on the payload is explained. To estimate the payload for the proposed cycle quantitatively, hypothetical low earth orbit (LEO) and KSLV-II sun synchronous orbit (SSO) mission cases are used. In the analysis of the hypothetical LEO mission, it is found that the proposed cycle payloads are only 2.1% to 3.5% lower than those of the gas generator cycle when the combustion chamber pressure is 3.0 MPaA. For the KSLV-II SSO mission, the cargo payload is increased by 3.7% compared to that of gas generator cycle if the proposed cycle is employed for the third-stage engine.

Energy Efficiency and Economic Aspects of Mining Wastes Utilization within the Closed Cycle of Underground Gas Generator

Energy efficiency of coal gasification with possible utilization of mining wastes within ecologically closed gas generator cycle has been considered. Technical and technological performance of such gas generator and mechanism of material and heat balance on the basis of the available analytical methods and practices as well as the developed author software have been proposed. Heat carrier formed in the process of coal gasification has been used for the utilization. Temperature of the utilization process within the industrially expedient limits being supported with the help of either activation or attenuation of the gasification process. After specific treatment, organogenic waste and domestic wastes are utilized by means of thermal decomposition within a gas generator. Economic evaluation of the proposed means confirms the expediency of their implementation in mines with industrial and balanced coal reserves as well as within the areas where this energetic source has already been already mined out. Results of this investigation were partially presented on international scientific and practical conference “Forum of Miners - 2017”. They contain the researches, which were conducted within the project GP – 489, financed by Ministry of Education and Science of Ukraine.

Improving the performance of LOX/kerosene upper stage rocket engines

Improved liquid rocket engine cycles were proposed and analyzed via comparison with existing staged combustion and gas-generator cycles. The key features of the proposed cycles are regenerative cooling of thrust chamber by oxygen and subsequent use of this oxygen for driving one or two oxygen pumps. The fuel pump(s) are driven in a conventional manner, for example, using a fuel-rich gas-generator cycle. Comparison with staged combustion cycle based on oxygen-rich pre-burner showed that one of the proposed semi-expander cycles has a specific impulse only on 0.4% lower while providing much lower oxygen temperature, more efficient tank pressurizing system and built-in roll control. This semi-expander cycle can be considered as a more reliable and cost-effective alternative of staged combustion cycle. Another semi-expander cycle can be considered as an improvement of gas-generator cycle. All proposed semi-expander cycles were developed as a derivative of thrust chamber regenerative cooling performed by oxygen. Analysis of existing oxygen/kerosene engines showed that replacing of kerosene regenerative cooling with oxygen allows a significant increase of achievable specific impulse, via optimization of mixture ratio. It is especially the case for upper stage engines. The increasing of propellants average density can be considered as an additional benefit of mixture ratio optimization. It was demonstrated that oxygen regenerative cooling of thrust chamber is a feasible and the most promising option for oxygen/kerosene engines. Combination of oxygen regenerative cooling and semi-expander cycles potentially allows creating the oxygen/kerosene propulsion systems with minimum specific impulse losses. It is important that such propulsion systems can be fully based on inherited and well-proven technical solutions. A hypothetic upper stage engine with thrust 19.6kN was chosen as a prospective candidate for theoretical analysis of the proposed semi-expander cycles. The newly-developed software RECS was used for the comparative analysis of engine cycles.

Rocket Engine Feasibility Study for the JAXA Future Transportation Reference System

Bioethanol is a suitable candidate as a reusable rocket engine fuel because it is carbon neutral and environmentally friendly. Based on the JAXA future transportation reference system, the feasibility of a gas generator cycle fueled by bioethanol was examined in this paper. The result is considered to be a general design guideline of an ethanol-fueled rocket engine. A combustion pressure of 7 MPa, an expansion ratio of 25, a turbine efficiency of 0.6 and a pump efficiency of 0.7 enabled the engine to satisfy the required specifications with an Isp efficiency of 0.96. In addition, a flow experiment of bioethanol under high temperature and high pressure was conducted. Resulting from the EPMA analysis, the presence of sulfur and coking was recognized on the copper-alloy tube.

Advantage of ethanol fuel for gas generator cycle air turbo ramjet engine

Performance analyses of Gas Generator cycle Air Turbo Ramjet (GG-ATR) engine are conducted to investigate the feasibility of ethanol fuel, as comparing Liquefied Hydrogen (LH2), Liquefied Natural Gas (LNG), and n-C12H26. Liquefied Oxygen (LOX) is considered as an oxidizer. Gas generator combustion temperature and compressor pressure ratio are selected as the analytical parameters, which are varied from 900 to 1700 K and from 2.0 to 4.0, respectively.In general, specific impulses of light molecule weight fuels, such as LH2 and LNG, are higher than those of large molecule weight fuel. For storable fuel, Isp of ethanol is larger than that of n-C12H26 when gas generator temperature is less than 1300 K, although specific heat release of ethanol is only two thirds of n-C12H26 heat release. The main chemical species of n-C12H26 gas generator (GG) combustion gas are CH4 and CO. On the other hand, ethanol GG combustion gas is mainly composed of H2 and CO, resulting in smaller molecular weight than n-C12H26. These characteristics contribute to larger Isp of ethanol than n-C12H26 and indicate the feasibility of ethanol for GG-ATR engine fuel.

System performance tests and start transient analysis of a liquid rocket engine turbopump

This article describes a series of development tests of a turbopump, which can be applied to a gas generator cycle rocket engine with liquid oxygen and kerosene propellants. A turbine drives both an oxidizer pump and a fuel pump in the turbopump assembly. In the tests, liquid oxygen and kerosene are supplied to the oxidizer pump and the fuel pump, respectively, while either cold hydrogen gas or hot gas from the gas generator is supplied to the turbine. The turbopump is operated reliably at both on-design and off-design conditions, meeting all the performance requirements. The test results are compared with those of the turbopump component tests, where model fluids are used, that is, water for the oxidizer pump and the fuel pump, and cold air for the turbine. The turbopump tests results agree well with the turbopump component test results. The speed buildup of the turbopump at start period is calculated when pressurized gas is used to initially spin the turbine. A differential equation which represents the torque balance between the turbine and the pumps is solved. The calculation shows a good agreement with the test result. When the mechanical loss of the turbopump is considered, a better estimation is obtained.

유전알고리즘을 이용한 액체로켓엔진 설계 최적화

유전알고리즘을 사용하여 액체로켓엔진의 연소실 압력과 노즐 확장비, O/F 비 등 주요 설계변수를 최적화하였다. 대상엔진은 LO2/RP-1을 추진제로 사용하는 개방형 가스발생기 사이클을 대상으로 하였다. 연소실의 물성치는 CEA2를 이용하였으며, 무게 산출은 참고문헌을 바탕으로 모델링 하였다. 최적 설계의 목적함수는 비추력과 추력중량비를 다중목표로 설정하여 가중치 방법을 사용하였다. 유전알고리즘을 최적화 과정을 거친 결과 비추력은 최대 4%, 추력중량비는 최대 23% 정도 증가하였다. 또한 다양한 추력에 대해서 Pareto frontier line을 얻었다. A genetic algorithm (GA) has been employed to optimize the major design variables of the liquid rocket engine. Pressure of the main combustion chamber, nozzle expansion ratio and O/F ratio have been selected as design variables. The target engine has the open gas generator cycle using the LO2/RP-1 propellant. The gas properties of the combustion chamber have been obtained from CEA2 and the mass has been estimated using reference data. The objective function has been set as multi-objective function with the specific impulse and thrust to weight ratio using the weight method. The result shows about 4% improvement of the specific impulse and 23% increase of the thrust to weight ratio. The Pareto frontier line has been also obtained for various thrust requirements.

추진제 공급압력이 액체로켓엔진의 성능에 미치는 영향

가스발생기 사이클 액체로켓엔진에서 추진제의 공급압력 변화에 대한 성능 즉, 연소압, 터빈 파워, 엔진 혼합비, 가스발생기 연소가스의 온도 변화를 제시하였다. 로켓엔진의 주요 13개 시스템 레벨변수를 이용하여 엔진 성능을 수치적으로 계산한다. 산화제 공급압이 증가하면 연소압과 터빈 파워는 증가하며 연료 공급압이 증가하면 연소압과 터빈 파워가 감소한다. 연료 유량 증가에 따라 감소된 가스 발생기의 혼합비는 연소가스 온도를 감소시키며 터빈 구동매질로서의 연소가스 물성을 저하시킨다. 연료 유량 증가에 따라 감소된 터빈 파워는 엔진 추력에 직접 영향을 미치는 주연소기의 연소압을 감소시킨다.