Fuel-rich Gas Generator Research Articles

Thermodynamic analysis of the hydrocarbon-fuelled air-turborocket engine with complete-combustion gas generator

The Air-Turborocket (ATR) engine can work at Mach 0∼4 or even higher speed, which is considered one of the best low Mach number propulsion systems for reusable hypersonic vehicles. However, because the hydrocarbon-fuelled ATR engine uses a fuel-rich gas generator, the combustion product contains a large amount of C(gr) that can cause coking in the turbine in a few minutes. To solve this problem, an ATR engine cycle with a complete-combustion gas generator (ATR-CCGG) was proposed. The performance of this cycle has been analysed through the thermodynamic model, and the influence factors and the sensitivity study of the cycle performance have been investigated. The results show that when the equivalence ratio is 1, the cycle can get more than 700 s of specific impulse and 1000 m/s of specific thrust at supersonic speed. Although the performance at subsonic speed is lower than that of the LOX/Kerosene ATR engine, the gas generator without C(gr) can ensure the engine to work for hours without coking in the turbine at different Mach numbers, which can be used in reusable hypersonic vehicles or single-stage-to-orbit missions.

Influence of Doped H2O or H2 on Soot Production and Power Capability in the Fuel-rich Gas Generator

ABSTRACT Influence of doped H2O or H2 on both soot production and power capability in the fuel-rich gas generator has been studied together by using the program of chemical equilibrium with applications (CEA). The oxidant is LOX, and the fuel is composed of Jet-A and the additive. The parameters of the gas generator are as follows: the range of combustion temperature is 800–1700 K, combustion pressure is 0.1–5.0 MPa, oxidant/fuel ratio is 0.1–1.2, and the mass percent of the additive in fuel is 0–60%. The results indicate that the addition of either H2O or H2 can obviously reduce the mass percent of soot in combustion products, and the reasons are discussed on the base of the products distributions. Moreover, the minimum amounts of addition to surrender mass percent of soot less than 0.1% are present. The effects of combustion pressure on soot mass percent in combustion products appear turning appoints around 1100 K, no matter the additive is H2O or H2. The addition of H2 can obviously improve the power capacity of combustion products in the whole temperature range. The addition of H2O can also improve slightly the power capacity of combustion products, when the combustion temperature is less than 1400 K. Effective molar weight of combustion products is the main factor affecting power capacity.

Concept and performance study of turbocharged solid propellant ramjet

This study proposes a turbocharged solid propellant ramjet (TSPR) propulsion system that integrates a turbocharged system consisting of a solid propellant (SP) air turbo rocket (ATR) and the fuel-rich gas generator of a solid propellant ramjet (SPR). First, a suitable propellant scheme was determined for the TSPR. A solid hydrocarbon propellant is used to generate gas for driving the turbine, and a boron-based fuel-rich propellant is used to provide fuel-rich gas to the afterburner. An appropriate TSPR structure was also determined. The TSPR's thermodynamic cycle was analysed to prove its theoretical feasibility. The results showed that the TSPR's specific cycle power was larger than those of SP-ATR and SPR and thermal efficiency was slightly less than that of SP-ATR. Overall, TSPR showed optimal performance in a wide flight envelope. The specific impulses and specific thrusts of TSPR, SP-ATR, and SPR in the flight envelope were calculated and compared. TSPR's flight envelope roughly overlapped that of SP-ATR, its specific impulse was larger than that of SP-ATR, and its specific thrust was larger than those of SP-ATR and SPR. Attempts to improve the TSPR off-design performance prompted our proposal of a control plan for off-design codes in which both the turbocharger corrected speed and combustor excess gas coefficient are kept constant. An off-design performance model was established by analysing the TSPR working process. We concluded that TSPR with a constant corrected speed had wider flight envelope, higher thrust, and higher specific impulse than TSPR with a constant physical speed determined by calculating the performance of off-design TSPR codes under different control plans. The results of this study can provide a reference for further studies on TSPRs.

혼합식 가스발생기의 연료과농 연소특성

본 논문은 하이브리드 로켓 성능향상을 위하여 가스발생기형과 후방 연소형 개념을 결합한 혼합형 하이브리드 로켓을 제안하고 있다. 특히 고체 추진제를 사용하는 기존의 가스발생기와 달리, 고체연료와 액체/기체 산화제를 적용한 혼합식 가스발생기를 제안하였으며 혼합식 가스발생기의 연료과농 연소특성을 확인하기 위하여 연료 길이, 산화제 유량, 연료 내경 그리고 연료 종류를 변화하며 연소가스 온도 변화를 측정하였다. 그러나 이들 인자 변화에 의한 온도변화가 매우 제한적이므로 또 다른 인자로 <TEX>$O_2$</TEX>와 <TEX>$N_2$</TEX>를 혼합한 혼합산화제를 사용하였다. 이때 가스발생기의 연소가스 온도의 요구조건은 1600 K이하로 설정하였으며 연소 시험에서 혼합식 가스발생기는 온도조건을 만족하는 연료과농 연소가스가 생성되었음을 확인하였다. 그러나 온도에 따른 검댕의 발생특성과 다른 이전 연구들에서 제시하는 가스발생기 연소가스 온도 요구조건이 1200 K이하임을 고려할 때, 최종적으로 이 조건을 만족하는 연료과농 연소가스를 생성할 계획이다. In this study, a combined hybrid rocket system is newly introduced which has characteristics of both gas generators and afterburner type hybrid rockets. In particular, a combined gas generator utilizing solid fuel and liquid/gas oxidizer was designed as a primary combustor of the system. Combustion tests were carried out with various equivalence ratio affected by parameters such as fuel length, oxidizer flow rate, fuel port diameter and fuel type. In general, fuel-rich gas generator produces low combustion gas temperature to meet the temperature requirement and the target temperature was transiently set less than 1600 K. Since it was found that controlling parameters showed limited effects on the change of equivalence ratio, mixture of <TEX>$O_2$</TEX> and <TEX>$N_2$</TEX> as an oxidizer was additionally introduced. As a result, a combined gas generator successfully produced combustion gas temperature of less than 1600 K Future studies will carry out more combustion tests to attain fuel-rich combustion gas temperature less than 1200 K, which was a temperature requirement of a gas generator system in the previous studies.

Genetic algorithm to optimize the design of main combustor and gas generator in liquid rocket engines

A genetic algorithm was used to develop optimal design methods for the regenerative cooled combustor and fuel-rich gas generator of a liquid rocket engine. For the combustor design, a chemical equilibrium analysis was applied, and the profile was calculated using Rao’s method. One-dimensional heat transfer was assumed along the profile, and cooling channels were designed. For the gas-generator design, non-equilibrium properties were derived from a counterflow analysis, and a vaporization model for the fuel droplet was adopted to calculate residence time. Finally, a genetic algorithm was adopted to optimize the designs. The combustor and gas generator were optimally designed for 30-tonf, 75-tonf, and 150-tonf engines. The optimized combustors demonstrated superior design characteristics when compared with previous non-optimized results. Wall temperatures at the nozzle throat were optimized to satisfy the requirement of 800 K, and specific impulses were maximized. In addition, the target turbine power and a burned-gas temperature of 1000 K were obtained from the optimized gas-generator design.

Combustion Dynamics of Swirl Coaxial Injectors in Fuel-Rich Combustion

An experimental study was carried out to investigate the combustion characteristics of liquid–liquid swirl coaxial injectors in a fuel-rich gas generator. Propellants of liquid oxygen and kerosene (jet A-1) were burned in a range of mixture ratios (0:29 0:38) and chamber pressures (53:6 57:9 bar) in several combustion chambers with differentdiametersandlengths.Aninjectorheadhad37identicalswirlcoaxialinjectorswithlargerecesslengthsand internal-mixingcharacteristics.Theinjectorheadwasconnectedtocombustionchambersandachokednozzleeither with or without an extension pipe. It was found that the combustion characteristic velocity, combustion gas temperature, and combustion dynamics were seldom influenced by the increase of injector recess length in the present fuel-rich conditions. The results of dynamic pressure data and swirl-injector dynamics suggest that the longitudinal-modecombustioninstabilityinthefuel-richgasgeneratorequippedwithbiswirlcoaxialinjectorscould besignificantlyaffected bytherelationship betweenthe resonant frequencyin thecombustion chamberandthe flow dynamics of the swirl injector.

Conceptual design for a kerosene fuel-rich gas-generator of a turbopump-fed liquid rocket engine

A design method for a kerosene fuel-rich gas-generator of a liquid rocket engine using turbopumps to supply propellant was performed at a conceptual level. The gas-generator creates hot gases, enabling the turbine to operate the turbopumps. A chemical non-equilibrium analysis and a droplet vaporization model were used for the estimation of the burnt gas properties and characteristic chamber length. A premixed counter-flow flame analysis was performed for the prediction of the burnt gas properties, namely the temperature, the specific heat ratio and heat capacity, and the chemical reaction time. To predict the vaporization time, the Spalding model, using a single droplet in convective condition, was used. The minimum residence time in the chamber and the characteristic length were calculated by adding the reaction time and the vaporization time. Using the characteristic length, the design methods for the fuel-rich gas-generator were established. Finally, a parametric study was achieved for the effects of the O/F ratio, mass flow rate, chamber pressure, initial droplet temperature, initial droplet diameter and initial droplet velocity.

Fuel-Rich Combustion Characteristics of Biswirl Coaxial Injectors

An experimental study was performed to investigate the combustion characteristics of liquid–liquid swirl coaxial injectors in fuel-rich conditions. Liquid oxygen and kerosene (Jet A-1) were burned in a range of mixture ratios (0.29–0.41) and chamber pressures (46–65 bar) in a gas generator for a liquid rocket engine. An injector head was connected to a water-cooled chamber and a short nozzle with or without an extension pipe between the chamber and the nozzle. The extension pipe acoustically simulated a turbine inlet manifold. The injector head had 37 identical swirl coaxial injectors. It is found that the characteristic velocity and combustion gas temperature are seldom influencedby the extension pipe, but are only functions of themixture ratio. The dynamic pressure data show that the combustion instability in the fuel-rich gas generator equipped with biswirl coaxial injectors can be significantly affected by themixture ratio and also by the extension pipe, which influences the resonant frequency in the chamber.

Combustion Dynamics and Stability of a Fuel-Rich Gas Generator

The dynamic characteristics of fuel-rich combustion have been studied using an experimental combustor simulating a gas generator for a liquid rocket engine. The combustor burns liquid oxygen and fuel (Jet A-1) at a mixture ratio of about 0.32 and a chamber pressure ranging from 4.10 to 7.24 MPa, which covers subcritical to supercritical pressures of oxygen. For the investigation of combustion dynamics, pressure fluctuation measurements using piezoelectric sensors have been a major probe throughout the present study. Two different types of injector heads equipped with biliquid swirl coaxial injectors and either a short nozzle or a turbine manifold nozzle have been used in the study. Fuel-rich combustion of both injector heads with the short nozzle revealed pressure pulsation at frequencies of about 128 Hz, which attenuates along with an increase of a chamber pressure. Combustion tests with the turbine manifold nozzle conducted at chamber pressures lower than the oxygen critical pressure showed combustion instabilities at a frequency of 330 Hz, which has been identified as a longitudinal resonant mode by a linear acoustic analysis. The combustion instabilities seem to be induced by inherent pressure fluctuations from the biliquid swirl coaxial injector when the chamber pressure is below the oxygen critical pressure.

추력 30톤급 액체산소/케로신 로켓엔진 연소장치 개발(II)-가스발생기

추력 30톤급 펌프공급방식 액체로켓엔진을 위한 가스발생기 개발 과정에 관하여 기술하였다. 액체산소와 케로신을 추진제로 연료 과농 조건에서 작동하는 가스발생기의 개발을 위해 분사기 개발에서부터 시작하여 축소형, 실물형 개발시제를 거쳐 가스발생기 단품 개발을 성공적으로 완료하였다. 가스발생기 설계 과정에서 다양한 해석적 방법을 적용하였으며 점화 시험, 연소성능 및 연소안정성 평가시험, 내구성 평가시험 등을 통해 가스발생기의 성능요구사항을 시험적으로 검증하였다. 개발된 가스발생기는 연소압 및 혼합비 운용 영역 내에서 안정적으로 작동하며 성능 및 수명 요구조건을 만족시킴을 확인하였다. The development process of a gas generator for a 30-tonf pump-fed space liquid rocket engine is described. Starting from the development of an injector, followed by subscale and full-scale test specimens, the development of LOx/kerosene fuel-rich gas generator has been concluded successfully. Various analytical methods have been utilized in the course of design and the performance requirements have been verified experimentally through ignition tests, combustion performance and stability assessment tests and duration tests. The gas generator has proven its workability and stability within a defined operation window of varying chamber pressure and mixture ratio and demonstrated compliance to the performance and life time requirements.

연료 과농 가스발생기의 연소 가스 물성치에 관한 연구

액체 로켓 엔진용 가스발생기 개발을 위해서는 추진제 O/F ratio에 따른 연소 가스의 열역학적 물성치 예측이 필수적이다. 본 연구에서는 액체산소/Jet A-1 조합의 연료 과농 가스발생기의 실 추진제 연소 시험을 통해 O/F ratio 변화에 따른 연소 생성 가스의 온도를 계측하였다. 또한 연소실 내 동압 섭동 측정 및 정압 측정 결과를 이용하여 비열비, 가스 상수, 정압 비열과 같은 연소 가스의 열역학적 물성치를 간접적으로 산출해내었다. 본 실험값은 화학평형코드 결과를 통해 구한 보간 계수를 이용한 예측 결과와 비교해보았을 때 동일한 경향 및 유사한 값을 가지는 것으로 밝혀졌으며 이는 보간 계수 예측 방법이 설계 도구로 충분히 적용 가능하다는 결과를 확인하였다. It is essential to predict thermodynamic properties of combustion gas with respect to a propellant mixture ratio for the development of a gas generator for a liquid rocket engine. The present study shows the temperature measurement of exit combustion gas as a function of a mixture ratio through the series of combustion tests of a fuel-rich gas generator with liquid oxygen and Jet A-1. The measurements of dynamic and static pressures, and combustion gas temperatures allowed the estimation of thermodynamic properties like a specific heat ratio, a gas constant, and a constant pressure specific heat of the combustion gas. The comparison of the experimental results with predictions made by interpolation parameters obtained from the modification of the chemical equilibrium code indicates that the interpolation method calibrated using the temperature measurements can be utilized as an effective tool for the initial design of a fuel-rich gas generator.

액체로켓엔진에서 충돌형 분사기 형태의 연료과잉 가스발생기 연소특성

본 논문에서는 충돌형 분사기를 장착한 연료 과잉 가스발생기에서 수행한 설계점 및 탈설계점 연소시험의 전반적인 결과에 대하여 논하였다. 가스발생기는 충돌형 분사기와 추진제 공급 메니폴드로 구성된 분사기 헤드, 물냉각 채널을 가진 연소실, 혼합을 증가시키는 turbulence ring, 온도 및 압력을 측정하는 링, 그리고 노즐로 구성되었다. 여러 운영조건에서 연소시험은 성공적이었으며 가스발생기 손상은 발생하지 않았다. 체류시간 4～6msec 정도에는 출구온도변화가 거의 없었지만 압력변동에 따라 출구온도는 변하였다. 측정되는 압력, 유량 그리고 노즐목 크기로 계산한 연소효율는 출구에서 측정한 온도의 제곱근에 비례하는 관계식을 저 혼합비 가스발생기에서도 갖고 있었다. 가스발생기의 O/F 비 변화에 가장 민감하게 출구온도가 변화하였으며 이에 대한 관계식을 도출하여 향후 설계 기초 자료로 활용되게 하였다. The overall results of hot firing tests of fuel-rich gas generator with impinging injector at design and off-design points are described. The gas generator consists of an injector head with impinging injector, a water cooled combustor wall, a turbulence ring to enhance mixing, an instrument ring measuring temperature and pressure and a nozzle. The combustion tests were successfully performed without damage of gas generator. Test results show that the outlet temperature is not dependent on residence time of hot gas within 4～6msec but dependent on chamber pressure. The relation between outlet temperature and combustion efficiency resulting from measured pressure, mass flow rate and area of nozzle throat is shown. The overall O/F ratio is the critical parameter to determine the outlet temperature and the linear correlation between two parameters is established.

액체로켓의 농후 가스발생기 최적설계

액체로켓 엔진에 사용되는 가스발생기를 최적설계 하였다. 추진제는 RP-1/LOx 이고, open cycle터보펌프 시스템을 사용하였으며, 가스발생기는 농후 (fuel-rich) 연소를 적용하였다. 최적설계의 목적함수는 주연소설의 비추력의 최대화이고 설계 제한조건은 가스발생기 연소 실온도와 터빈-펌프의 출력일치이다. 가스 발생기의 설계에 사용된 설계변수는 가스발생기 유량, O/F비, 터빈 노즐 입구 각, 부분분사비, 그리고 터빈 원주속도이며 이들을 이용하여 가스발생기의 열역학적 성능을 계산하였다. 그리고 설계 제한조건을 만족하면서 목적함수를 최대화 할 수 있는 가스발생기의 크기와 성능조건을 확인하였다. 설계된 가스발생기 기본형상은 연소시험에 적용된 후 최종적으로 결정된다. An optimal design of the gas generator for Liquid Rocket Engine (LRE) was conducted. A fuel-rich gas generator in open cycle turbopump system was designed for 10ton in thrust with RP-1/LOx propellant. The optimal design was done for maximizing specific impulse of thrust chamber with constraints of combustion temperature and for matching the power requirement of turbopump system. Design variables are total mass flow rate to gas generator, O/F ratio in gas generator, turbine injection angle, partial admission ratio, and turbine rotational speed. Results of optimal design provide length, diameter, and contraction ratio of gas generator. And the operational condition predicted by design code with resulting configuration was found to maximize the objective function and to meet the design constraints. The results of optimal design will be tested and verified with combustion experiments.

Nonequilibrium combustion model for fuel-rich gas generators

The efflux composition characteristics of an experimental fuel-rich gas generator have been modeled with a detailed hydrocarbon kinetics package that includes the effects of liquid droplet vaporization. The chemical reaction mechanism, consisting of 107 chemical species and 642 reversible reactions, is capable of describing a complex kerosene-type fuel mixture for the case where the constituents (aliphatic, aromatic content) can be specified. A liquid fuel droplet vaporization model for octane and toluene mixtures has been included. A flame zone model is incorporated to address the fast reaction phenomena necessary for sustained combustion. Successful modeling has necessitated the use of a nonequilibrium approach incorporating a detailed chemical model. Comparison of modeling results to experimental data has shown good agreement, clearly supporting nonequilibrium operation. 21 refs.