Liquid Propellant Rocket Engines Research Articles

Effect of Chamber Backpressure on Swirl Injector Fluid Mechanics

Fluid mechanics of a liquid swirl injector element at various chamber backpressures were investigated. The center-jet swirling element was designed using typical liquid propellant rocket engine parameters, then manufactured and tested in a high-pressure, optically accessible, cold flow facility. Water was injected into a chamber pressurized with gaseous nitrogen at a constant swirl injector flow rate of 0.09 kg/s. The chamber backpressure ranged from 0.10 to 4.81 MPa. The film thickness and spray angle near the nozzle exit were measured by shadowgraphy. The film thickness was also measured within the injector upstream of the exit through a transparent nozzle tube section. Increasing the chamber backpressure for this fixed mass flow rate increased the film thickness from predicted design values. Measured discharge coefficient values increased with increasing chamber backpressure, reflecting the observed increase in internal nozzle film thickness. The spray angle decreased for increasing chamber backpressure.

Ricardo Dyrgalla (1910–1970), pioneer of rocket development in Argentina

One of the most important developers of liquid propellant rocket engines in Argentina was Polish-born Ricardo Dyrgalla. Dyrgalla immigrated to Argentina from the United Kingdom in 1946, where he had been studying German weapons development at the end of the Second World War. A trained pilot and aeronautical engineer, he understood the intricacies of rocket propulsion and was eager to find practical applications to his recently gained knowledge. Dyrgalla arrived in Argentina during Juan Perón's first presidency, a time when technicians from all over Europe were being recruited to work in various projects for the recently created Argentine Air Force. Shortly after immigrating, Dyrgalla proposed to develop an advanced air-launched weapon, the Tábano, based on a rocket engine of his design, the AN-1. After a successful development program, the Tábano was tested between 1949 and 1951; however, the project was canceled by the government shortly after. Today, the AN-1 rocket engine is recognized as the first liquid propellant rocket to be developed in South America. Besides the AN-1, Dyrgalla also developed several other rockets systems in Argentina, including the PROSON, a solid-propellant rocket launcher developed by the Argentine Institute of Science and Technology for the Armed Forces (CITEFA). In the late 1960s, Dyrgalla and his family relocated to Brazil due mostly to the lack of continuation of rocket development in Argentina. There, he worked for the Institute of Aerospace Technology (ITA) until his untimely death in 1970. Ricardo Dyrgalla deserves to be recognized among the world's rocket pioneers and his contribution to the science and engineering of rocketry deserves a special place in the history of South America's rocketry and space flight advocacy programs.

A liquid propulsion panorama

Liquid-propellant rocket engines are widely used all over the world, thanks to their high performances, in particular high thrust-to-weight ratio. The present paper presents a general panorama of liquid propulsion as a contribution of the IAF Advanced Propulsion Prospective Group. After a brief history of its past development in the different parts of the world, the current status of liquid propulsion, the currently observed trends, the possible areas of future improvement and a summarized road map of future developments are presented. The road map includes a summary of the liquid propulsion status presented in the “Year in review 2007” of Aerospace America. Although liquid propulsion is often seen as a mature technology with few areas of potential improvement, the requirements of an active commercial market and a renewed interest for space exploration has led to the development of a family of new engines, with more design margins, simpler to use and to produce associated with a wide variety of thrust and life requirements.

Erratum to: “Numerical and experimental investigations of the effect of repeated cyclic thermal loads on the strength of materials of recoverable liquid-propellant rocket engines”

Erratum to: “Numerical and experimental investigations of the effect of repeated cyclic thermal loads on the strength of materials of recoverable liquid-propellant rocket engines”

Diagnostics of acoustic properties of the combustion zone by means of periodic changes in the nozzle throat area

A mathematical model is proposed, which allows calculating the amplitude of forced pressure oscillations in the combustor of a liquid-propellant rocket engine, caused by periodic changes in the nozzle throat area, with allowance for acoustic properties of the combustion zone. Results calculated for some particular cases are given.

Opportunities for a Liquid Rocket Feed System Based on Electric Pumps

A feed system for liquid-propellant rocket engines based on electric pumps powered by batteries is proposed. It is proven to stand as a viable alternative to the pressure-gas feed system. The dependence of the feed system mass on the different operating parameters is obtained so as to identify the conditions favoring its adoption, that is, a relatively long burning time and a fairly high chamber pressure. Under such conditions, the proposed system is shown to offer significant mass savings with respect to the pressure-gas system when advanced batteries are used. This advantage is further enhanced by the beneficial effect of chamber pressure on the engine effective exhaust velocity. A test case for a low Earth orbit to geostationary equatorial orbit transfer is also presented to identify the optimum value of the burning time, deriving from the competition between the feed system mass and the effect of gravitational losses.

Numerical and experimental investigations of the effect of repeated cyclic thermal loads on the strength of materials of recoverable liquid-propellant rocket engines

Elements of the liquid-propellant rocket engines (LPREs) are subject to the action of intense mechanical and thermal loads, which are temporally related to the start-up, normal regime, and switch-off periods of the working cycle. Recoverable LPREs experience these loads many times. Repeated (cyclic) loads lead to the manifestation of processes typical of fatigue, whereby residual stresses and (plastic) deformations are retained and accumulated in the LPRE materials. Numerical investigations using models of the nonstationary thermal and stressed-strained state and the plastic behavior of materials have been performed in order to determine the limiting numbers of loading cycles for the structural materials employed in various units of LPREs. A special computational-experimental procedure is described, which allows these materials to be studied in a stressed-strained state close to that in real LPREs.

Distinctive features of the intrachamber instability of combustion in liquid-propellant rocket engines

Self-oscillations and certain of their regularities determined by solution of a degenerate system of differential equations that is used in considering combustion instability in combustion chambers of liquid-propellant rocket engines are modeled mathematically.

Numerical simulation and investigation of working process features in high-duty combustion chambers

Basic concepts of a numerical simulation method of two-phase turbulent flows with combustion are stated. Results of computations in gas generators and combustion chambers of liquid-propellant rocket engines operating on oxygen and methane are presented. Features of the processes of evaporation, mixture, flow, and combustion of the propellant within chambers with tree types of injectors, i.e., coaxial-jet gas-liquid, liquid-liquid monopropellant and bipropellant impinging-jets, are studied.

액체 추진제 로켓엔지의 혼합비 안정기

There is provided a mixture ratio stabilizer for a liquid propellant rocket engine which comprises: a body internally having a chamber and a fuel supply hole communicating with an end of the chamber to receive fuel from a fuel supply pump, an oxidizer inlet, and a fuel outlet; a spool guide in the chamber and including a plurality of first orifices for the fuel outlet and the fuel supply hole to communicate with each other; a bellows in the chamber and having one end adjacent to the spool guide and the other end positioned adjacent to the oxidizer inlet, to expand and contract in accordance with the pressure of an oxidizer which flows into the oxidizer inlet; and a spool having a part placed in the spool guide.

ACOUSTIC CAVITIES DESIGN PROCEDURES

Combustion instability is recognized as one of the major problems frequently faced by engineers during the development of either liquid or solid propellant rocket engines. The performance of the engine can be highly affected by these high frequencies instabilities, possibly leading the rocket to an explosion. The main goal while studying combustion chambers instability, either by means of baffles or acoustic absorbers, is to achieve the stability needed using the simplest possible manner. This paper has the purpose of studying combustion chambers instabilities, as well as the design of acoustic absorbers capable of reducing their eigenfrequencies. Damping systems act on the chamber eigenfrequency, which has to be, therefore, previously known.

On porous liquid propellant rocket engine injectors

A novel injection concept for cryogenic liquid propellant rocket engines has been tested and verified successfully using an optically accessible combustion chamber. The injector consists of a porous faceplate made from sinter metal and five LOX posts arranged in a classical parallel showerhead configuration. While liquid oxygen is injected at similar velocities known from typical coaxial injector elements, the fuel enters the combustion chamber at much lower velocities. The flame stabilization for both LOX/H 2 and LOX/CH 4 propellant combination has been investigated at sub-, near-, and supercritical pressures.

Experimental investigation of high-frequency combustion instabilities in liquid rocket engine

High-frequency instabilities in liquid propellant rocket engines are experimentally investigated in a model scale research facility. Liquid oxygen and gaseous methane are injected in the combustion chamber at 0.9 MPa through three coaxial injectors vertically aligned. High-amplitude transverse pressure fluctuations are generated in the chamber at frequencies above 1 kHz by a rotating toothed wheel actuator which periodically blocks an auxiliary lateral nozzle. The chamber eigenmodes are identified in a first stage by examining the response of the system to a linear frequency sweep. In a second stage the chamber is excited at the frequency corresponding to the first transverse (1T) mode. The effect of the pressure mode on combustion is observed with intensified and high-speed cameras. Photo-multipliers and pressure sensors are also used to characterize the system behavior and examine phase relations between the corresponding signals. Flame structure modifications observed for specific injection conditions correspond to a strong coupling between acoustics and combustion which notably modifies the flow dynamics, augments the flame expansion rate and enhances heat transfer to the wall.

DYNAMIC BEHAVIOR OF A LIQUID ROCKET ENGINE FEEDING SYSTEM

This paper is the second part of a study conducted to investigate the dynamic behavior of a liquid propellant rocket engine. It is dedicated to complete the modeling, simulation and analysis of the dynamic behavior of the feeding system of a typical rocket engine. The elements involved are the gas generator, the turbo-pump assembly, and the combustion chamber. These individual elements are treated as separate units before dealing with the integrated system. The system incorporates the combustion chamber controller, the thrust controller, the gas generator controller, and the safety valve. Non-linear mathematical models and empirical relations have been deduced and used to develop computer simulation programs using MATLAB SIMULINK technique. The study was extended to investigate the static and dynamic behaviors of such components. The study showed that the thrust reaches always a new steady state value corresponding to pre-compression distance of the thrust controller spring and through regulating the combustion chamber pressure. The system, which has two power ratings, operates at a stable, self-regulated condition with slightly under-damped nature of settling time in switching from first to second power rating. The developed simulation programs could help investigate and analyze other similar liquid propellant rocket engines.

Eugen Sänger: Eminent space pioneer

In international literature on astronautics, three main space pioneers are mentioned: Konstantin E. Tsiolkovsky, Robert H. Goddard and Hermann Oberth. There are other two space pioneers that are very rarely mentioned: Robert Esnault-Pelterie and Eugen Sänger. Pelterie is known particularly in Europe, and Sänger is mentioned in the second half of the 20th century normally only in connection with space shuttle flights. Taking a look at Sänger's work and heritage, it is obvious that he greatly influenced the development of astronautics in terms of purely theoretical dissertations on achievable limits of space research as well as in terms of technical approaches to achieving the short- and long-term goals of astronautics, and in terms of setting tasks for organizing mankind to achieve these goals. Sänger's book “The Technology of Rocket Flight” was the first study based not only on basic research, but also on the applied research that he conducted and the findings of which he published in various papers. Sänger was clearly connected with and influenced the development of two experimental research groups in the US in the 1930s, which resulted in two of the most significant companies in the US in the 1950s that manufactured liquid propellant rocket engines. Basic and applied research in the field of space planes resulted in construction of rocket planes such as the US space shuttle and Soviet Buran shuttle. Sänger's research on subsonic and supersonic ramjets in combination with a turbojet engine provided a basis for developing this promising propulsion for use in subsequent space planes designed for flights into low Earth orbits. His pioneering work on the photon rocket represents human achievements in reaching almost unimaginable limits of space research. By striving for a peaceful international approach to space research, Sänger participated in establishing the non-governmental organization IAF (International Astronautical Federation) and realized his idea that space research is a concern for all mankind. He was therefore appointed the first president of the IAF. The paper presents how Sänger influenced the development of rocket technology and astronautics, which definitely ranks him with the first three space pioneers.

Research on health evaluation system of liquid-propellant rocket engine ground-testing bed based on fuzzy theory

In this paper, the theory based on multi-sensor information fusion is studied, which is used to evaluate the health condition of liquid-propellant rocket engine ground-testing bed. The concept of health degree is defined. It is used as a quantitative index for evaluating the health condition of the ground-testing bed. In order to evaluate the health condition of the ground-testing bed on different levels, health degrees of a single parameter, of a sub-system and of a system are defined. They accordingly measure the health conditions of single parameter, sub-system and system of the ground-testing bed. The method of fuzzy data fusion is used to calculate the health degree. In this method the weight of each monitoring sensor is calculated by analytic hierarchy process (AHP), then, the multi-sensor data are fused by the fuzzy comprehensive evaluation method, next, the sequence data are fused by two-grade index evaluation method, finally, the health degree is calculated by defuzzification method. Based on these, the health condition evaluation system of ground-testing bed is set up, which can evaluate the health condition of the ground-testing bed properly and quantitatively. At last, the application of the health evaluation system in fault detection and health evaluation of the ground-testing bed are discussed.

Structural features and behavior of seals in the hydrogen pump flow passages

Some features of designing seals for hydrogen pumps used in turbopump assemblies of liquid-propellant rocket engines are considered. These features are interrelated with specific physical and chemical characteristics of the working medium, namely, hydrogen.

Combustion Stability Boundaries of the Subscale Rocket Chamber with Impinging Jet Injectors

Combustion stability boundaries are investigated experimentally using a subscale chamber for screening the split-triplet impinging jet injectors of a liquid propellant rocket engine. Geometrical dimensions and operating conditions of the subscale chamber are determined using the scaling methods proposed in the previous works. From the experimental tests, two major instability regions are identified by the parameters of combustion-chamber pressure and mixture (oxidizer/fuel) ratio. A key instability mechanism can be explained by the correlation between the characteristic burning or mixing time and the characteristic acoustic time. In each instability region, the dynamic behavior of the flame is. investigated to verify the characteristic lengths of the jet injectors derived from hydrodynamic theory. It is found that pressure oscillations with large amplitude are generated by lifted-off flames. Stability margins are evaluated as a function of the impingement angle and thereby, the optimum angle is identified.

The thermal expansion of carbon/carbon composites from room temperature to 1400 °C

The coefficient of thermal expansion (CTE) is one of the most important properties for materials used under variable temperatures. Carbon/carbon composites must be capable of high geometric stability and low thermal expansion when they are used in engine turbine blades, rocket nozzles and seal rings in a liquid propellant rocket engine [1]. The thermal expansion of C/SiC composites and pyrocarbon films has been studied by Cheng et al. [2] and Taylor et al. [3], respectively. Carbon/carbon composites were developed using high-pressure impregnation/carbonization techniques and their thermophysical properties studied by Manocha et al. [4]. The thermophysical properties of five different carbon/carbon composites were discussed by Luo et al. [5]. Baxter et al. studied the effect of density on thermal expansion [6]. Piat and Schnack [7] studied pyrocarbon with different texture degrees and attempted to model the influence of microstructure on the coefficients of thermal expansion by collecting experimental data from the real physical structure using SAED and HRTEM. A material with a specific thermal expansion can be designed by adjusting the microstructure. Here we report the influence of heat treatment, preform architecture and the microstructure of the matrix carbon on the thermal expansion of carbon/carbon composites. A layer of carbon fiber cloth without a weft thread and a random short fiber layer were laminated alternately to prepare porous carbon fiber preform felt with an apparent density around 0.6 g/cm. The carbon cloth layer was stacked in such a way that each layer was oriented at 90 to the adjacent layer. Carbon/ carbon composites with a density of 1.75 g/cm were prepared by a thermal gradient CVI process using natural gas (98%CH4, 0.3%C3H8, 0.3%C4H10, 0.4% other hydrocarbons) as precursor. Specimens were subjected to heat treatment at different temperatures in an argon atmosphere. The detailed description of the preparation had been reported elsewhere [8]. The coefficient of thermal expansion of the densified composites was measured parallel and perpendicular to the carbon fiber cloth layer, or the random short fiber layer, from room temperature to 1400 C in an argon atmosphere by using a Netzsch DIL 402C dilatometer. The specimen size was 3.5 · 3.5 · 20 mm. The microstructures were analyzed using an X’Pert X-ray diffraction analyzer. Data for diffraction peaks (002) was obtained. The microstructures were examined by a Olympus PMG3 polarized light microscope. The thermal expansion of graphite materials is anisotropic and is higher perpendicular to the graphite layers than parallel. In carbon/carbon composites, the graphitic flat-layered structure crystallites of pyrocarbon are arranged on the surface of the carbon fiber, and grow around the carbon fiber in the form of a ringed multilayer structure, in which the graphitic layers are parallel to the fiber axis. Therefore, the thermal expansion is anisotropic as well if the preform J.-g. Zhao (&) K.-z. Li H.-j. Li C. Wang Y.-q. Zhai School of Materials Science, Northwestern Polytechnical University, Xi’an 710072, P.R. China e-mail: jgzhaoshi@163.com

Damage Mechanisms in Polymer Matrix Composites in Extreme Environments

In this work, potential problems with the application of polymer matrix composites (PMC) in extreme environments [1] is discussed. Then, two specific examples of the applications of PMCs in high voltage [2-7] and high temperature [8-15] situations are evaluated. The first example deals with damage evolution in high voltage composite insulators [2-7] with PMC rods subjected to a combined action of extreme mechanical, electrical and environmental stresses. These insulators are widely used in transmission line and substation applications around the world. Subsequently, advanced high temperature graphite/polyimide composites [8-15] are evaluated for aerospace applications. The composite investigated in this project were used to manufacture and successfully test a Rocket Based Combined Cycle (RBCC) third-generation, reusable liquid propellant rocket engine, which is one possible engine for a future single-stage-to-orbit vehicle [8].