Rocket Engine Design Research Articles

Modeling Fuel Film Cooling on a Flat Plate

Cooling of liquid rocket engine combustion chambers and nozzles is a critical component to liquid rocket engine design. A common method of cooling is liquid-fuel film cooling. Liquid fuel is injected along the surface of the wall to act as a barrier against the core combustion gases. A numerical model is developed for simulating a physical experimental test setup meant to study liquid-fuel film cooling over a flat plate using a hydrocarbon fuel. The model incorporates turbulent multiphase flow with species transport within a commercial computational fluid dynamics software package. Conjugate heat transfer is simulated through walls containing embedded cooling channels. A novel user-defined function is written to incorporate reactions between the liquid fuel and the freestream gases. Comparisons are made between simulations with and without the phase interactions, as well as with simplifications to the cooling channel geometry. It was found necessary to model the cooling channel geometry in order to avoid artificially reducing cooling performance.

High-pressure methane-oxygen flames. Analysis of sub-grid scale contributions in filtered equations of state

Turbulent combustion modelling under high-pressure conditions is a key issue for the design of future aero or rocket engines. In the context of large-eddy simulations, the exact filtered equation of state (EoS) is generally approximated by an EoS directly computed from the Favre filtered quantities for species mass fractions and temperature. The soundness of this approximation is presently addressed through the analysis of laminar CH4-air and CH4-O2 high-pressure premixed and non-premixed flames. Their computations were performed either with the ideal gas equation of state (EoS) or with a high-pressure package that gather a cubic EoS along with additional pressure-dependent transport parameters and thermodynamic relations. Various kinetic schemes are used for the computation of laminar premixed flames and a comparison with available experimental data is provided for flame speed (SL). Laminar premixed CH4-O2 flames exhibit micro metric flame thickness for pressures above 2.0MPa as well as a high flame temperature (>3000K). These flamelets were then filtered either with a Gaussian filter or through a β-pdf technique. The resulting filtered profiles were used to approximate the filtered pressure, p¯. A negligible error is observed for CH4-air flames when comparing p¯ with the exact filtered pressure. However, significant errors were found for high-pressure CH4-O2 flames, that increase with the filter size. This behaviour is exacerbated for transcritical injections meaning that classical techniques that use tabulated thermochemistry methods in compressible codes must be revisited in that context. A correction to the tabulated chemistry approach coupled to a presumed subgrid pdf is finally proposed to comply with high-pressure considerations.

Effect of pressure on hydrogen/oxygen coupled flame–wall interaction

The design and optimization of liquid-fuel rocket engines is a major scientific and technological challenge. One particularly critical issue is the heating of solid parts that are subjected to extremely high heat fluxes when exposed to the flame. This in turn changes the injector lip temperature, leading to possibly different flame behaviors and a fully coupled system. As the chamber pressure is usually much larger than the critical pressure of the mixture, supercritical flow behaviors add even more complexity to the thermal problem. When simulating such phenomena, these thermodynamic conditions raise both modeling and numerical specific issues. In this paper, both subcritical and supercritical hydrogen/oxygen one-dimensional, laminar flames interacting with solid walls are studied by use of conjugate heat transfer simulations, allowing to evaluate the wall heat flux and temperature, their impact on the flame as well as their sensitivity to high pressure and real gas thermodynamics up to 100 bar where real gas effects are important. At low pressure, results are found in good agreement with previous studies in terms of wall heat flux and quenching distance, and the wall stays close to isothermal. On the contrary, due to important changes of the fluid transport properties and the flame characteristics, the wall experiences significant heating at high pressure condition and the flame behavior is modified.

Flow separation in rocket nozzles under high altitude condition

The knowledge of flow separation in rocket nozzles is crucial for rocket engine design and optimum performance. Typically, flow separation is studied under sea-level conditions. However, this disregards the change of the ambient density during ascent of a launcher. The ambient flow properties are an important factor concerning the design of altitude-adaptive rocket nozzles like the dual bell nozzle. For this reason an experimental study was carried out to study the influence of the ambient density on flow separation within conventional nozzles.

Derating design for optimizing reliability and cost with an application to liquid rocket engines

Derating is the operation of an item at a stress that is lower than its rated design value. Previous research has indicated that reliability can be increased from operational derating. In order to derate an item in field operation, however, an engineer must rate the design of the item at a stress level higher than the operational stress level, which increases the item׳s nominal failure rate and development costs. At present, there is no model available to quantify the cost and reliability that considers the design uprating as well as the operational derating. In this paper, we establish the reliability expression in terms of the derating level assuming that the nominal failure rate is constant with time for a fixed rated design value. The total development cost is expressed in terms of the rated design value and the number of tests necessary to demonstrate the reliability requirement. The properties of the optimal derating level are explained for maximizing the reliability or for minimizing the cost. As an example, the proposed model is applied to the design of liquid rocket engines.

Evaluation of Mechanical Properties of the Rocket Thrust Chamber Material under Flight Conditions

Thrust chamber of rocket engines often operate under conditions of rapid heating environments with temperatures approaching the melting points of the materials involved. Therefore, for the optimum design of rocket engines, it is necessary to obtain the properties of thrust chamber materials under actual operating conditions. The heating rates and loading rates in conventional (laboratory) tensile stress-strain tests, which are intended to evaluate the high temperature tensile properties of the materials, are usually very low compared to that actually encountered in rocket engines. The heating rates in conventional stress-strain tests are of the order of 0.4K per second to 0.5K per second only, whereas the combustion and aerodynamic heating rates in rockets and re-entry vehicles will usually exceed 100K per second. In this context, a very important beginning has been made to experimentally determine the high temperature tensile properties of KC20WN (a cobalt based superalloy) used in earth storable liquid rocket engine thrust chamber under conditions of rapid rates of heating which actually exists during flight. These investigations have shown that the elevated temperature strength of KC20WN depends upon the heating rate (or heating time) and can be considerably higher for rapid-heating conditions than for conventional heating conditions.

Modeling and simulation of a GOX/kerosene subscale rocket combustion chamber with film cooling

Detailed knowledge on the heat transfer mechanisms is crucial for the design of reliable and efficient rocket engines. Due to the high heat loads of the combustion chamber walls and the corrosive hot gases, film cooling is often applied supplementary or even as a primary cooling technique. Nevertheless, the dominating processes determining the film effectiveness under the conditions representative for rocket combustors are still not fully understood. In context of the national research program Transregio SFB/TRR-40, TEKAN and TARES, the Institute for Flight Propulsion (LFA) of the Technische Universitat Munchen (TUM) and Airbus Defence and Space carry out experimental and numerical investigations on heat transfer and film cooling techniques at application-relevant combustion pressures and temperatures. In this paper, results from film cooling experiments and numerical simulations with liquid and transcritical kerosene films in a water-cooled GOX/kerosene rocket combustion chamber are presented. The tests have been performed at two different combustion chamber pressures and with two different throat diameters to study the influence of Reynolds and Mach number. In the numerical investigations, a major issue has been the modeling of kerosene films in sub- and transcritical state. For the modeling Airbus Defence and Space’s in-house tool, Rocflam-II has been applied. The main goal of Rocflam-II is to provide a tool package for the simulation of a wide range of rocket combustion devices, validated against experimental data. This includes the modeling of propellant injection, atomization, mixing, combustion, wall heat transfer, film modeling as well as the conjugate heat transfer into the chamber wall and the cooling channels.

Critical Ignition Criteria for Monomethylhydrazine and Red Fuming Nitric Acid

Understanding the parameters governing ignition is crucial for safe rocket engine design, and while hypergolic propellants have been used for several decades, there is still a lack of data on the early condensed-phase reactions for many systems. The ignition and combustion of monomethylhydrazine (MMH) and red fuming nitric acid (RFNA) have been investigated in an unlike doublet impinging jet apparatus. Variation of the jet diameter and total propellant mass flow rate at constant oxidizer-to-fuel and jet momentum ratio allowed for investigation of critical ignition thresholds. Ignition probability was determined for each condition and classified as good ignition, a transition region, or failed ignition. Conditions resulting in reactive stream separation are compared with literature data obtained with hydrazine/nitrogen tetroxide. Two critical MMH/RFNA ignition thresholds were identified: a residence time of 0.2–0.4 ms and a nondimensional sheet growth rate of . Two power law correlations were identified allowing for prediction of an MMH/RFNA sheet growth rate given a residence time or Reynolds number. Thermal ignition theory was adapted to simulate the experiment, and selection of model parameters resulted in a trend that supports the experimental data and provides insight into the condensed-phase reaction rates. The results from this work are intended to provide an experimental data set for validation of MMH/RFNA combustion models and to guide the selection of injection conditions for impinging jet systems.

Starting Transient Experiments in the MEC-80 Rocket Engine Scaled Ignition Device

One of the main sources of concern within the design and manufacturing of rocket engines is the starting device, due to the relatively low reliability of the ignition of the main combustion process in these engines. Not only is necessary for the igniter to produce a sound ignition, but the starting pressure transient should manifest a tightly controlled build-up, in a very definite time interval. A specific research and development work was done by the authors in order to reliably assess the requirements for the ignition device of the newly developed compound, solid-liquid experimental rocket engine MEC-80 of the ADDA team. Computer predictions and scaled experiments on the constant volume combustion into a modified calorimeter were performed with notable results that led to the optimal design of the ignition device of the motor. The test results are presented with emphasize on the predictability of the ignition delay.

Effective Modeling of Conjugate Heat Transfer and Hydraulics for the Regenerative Cooling Design of Kerosene Rocket Engines

The present work is motivated to develop a unified framework to simulate multi-physical processes which are crucial for trade-off design of liquid rocket thrust chambers among propulsive performance, regenerative cooling, and pressure budget. In this paper, an effective modeling of conjugate heat transfer and hydraulics through the regenerative cooling passage has been performed to quantitatively evaluate detailed cooling designs, including spirally twisted channels and bidirectionally branched circuit, as well as to provide the wall heat flux to a compressible reacting flow solver in an interactively coupled manner. The kerosene fuel used as coolant is modeled by a three-component physical surrogate, and the fluid properties required for calculation of a Nusselt number correlation and empirical resistance coefficients are computed over the entire thermodynamic states from compressed liquid to supercritical fluid using the NIST SUPERTRAPP. The present method has been applied to an actual regeneratively cooled thrust chamber and validated against measurement of hot-firing tests in terms of temperature increase and pressure drop of the coolant through the cooling passages. Based on the numerical results, supplementary effects of peripheral fuel cooling injection and thermal barrier coating are addressed.

Hydrogen fuel rocket engines simulation using LOGOS code

Abstract Computer aided design of new effective and clean hydrogen rocket engines needs mathematical tools for supercomputer modeling of hydrogen–oxygen components mixing and chemically reacting in rocket combustion chambers. The paper presents the results of computer code developing, verification and validation, making it possible to simulate unsteady processes of ignition and combustion of hydrogen fuel in rocket engines. Restrictions on unsteady gas dynamics working cycles supercomputer simulations due to accumulations of errors are developed.

3D fluid–structure interaction analysis of a typical liquid rocket engine cycle based on a novel viscoplastic damage model

SUMMARYIn many space missions, expandable or reusable launch systems are used. In this context, the reliable design of liquid rocket engines (LREs) is a key issue. In the present paper, we present a novel combination of numerical schemes. It is applied to model the extreme physical phenomena a typical LRE undergoes during its loading cycles. The numerical scheme includes a partitioned fluid–structure interaction (FSI) algorithm in combination with a unified viscoplastic damage model. This allows the complex description of the material response under cyclic thermomechanical loading taking place in LREs. In this regard, we focus on the response of the cooling channel wall that is made from copper alloys. For the coupled FSI analysis, the individual domains of the rocket thrust chamber are modeled by a 3D parametrized approach. The well‐established single field solver codes, DLR TAU for the hot gas and ABAQUS FE software for the structural domain, are coupled via the inhouse developed simulation environment ifls. Ifls provides the necessary algorithms for a partitioned coupling approach such as individual code steering, data interpolation, time integration and iteration control. Finally, the results of an FSI analysis of a complete engine cycle are presented. They show the potential of the new numerical scheme for the lifetime prediction of LREs.Copyright © 2013 John Wiley & Sons, Ltd.

Design of liquid-propellant rocket engines

A design method for liquid-propellant rocket engines is proposed, on the basis of models adjusted in response to test results. This method increases the reliability of the results and reduces ground testing of the motors. The effectiveness of the method is confirmed by ground testing and flight experience with the motors developed by Salyut Design Bureau at Krunichev State Space-Research Center.

Optimization of System Parameters for Liquid Rocket Engines with Gas-Generator Cycles

System design of liquid rocket engines must consider engine performance, weight, cost, and reliability requirements. A general design optimization framework has been developed in this paper to select the best system parameters for liquid rocket engines with gas-generator cycles. The object is to maximize the specific impulse and vacuum thrust-to-weight ratio of the engine with given system requirements and design assumptions by changing thrust-chamber pressure and mixture ratio. The system analysis, along with the engine weight estimation, is based on a modular scheme. Multidisciplinary design optimization formulations including multidisciplinary feasible and collaborative optimization are used, evaluated, and compared during the optimization process. Several techniques of multi-objective processing are also used to identify the Pareto frontier and the optimal compromise solutions. A proposed cryogenic-propellant engine using liquid oxygen and hydrogen with a gas-generator cycle is studied as a specific example. Moreover, uncertainties in the engine operation, such as thrust-chamber pressure and mixture ratio, are taken into account as random variables in the reliability-based optimization. Results are presented to illustrate the tradeoff between the engine performance and reliability requirements.

Model of the design of cryogenic liquid propellant rocket engines

Model of the design of cryogenic liquid propellant rocket engines

Injection coupling with high amplitude transverse modes: Experimentation and simulation

High frequency combustion instabilities have technical importance in the design of liquid rocket engines. These phenomena involve a strong coupling between transverse acoustic modes and combustion. They are currently being investigated by combining experimentation and numerical simulations. On the experimental level, the coupling is examined in a model scale system featuring a multiple injector combustor (MIC) comprising five coaxial injectors fed with liquid oxygen and gaseous methane. This system is equipped with a novel VHAM actuator (Very High Amplitude Modulator) which comprises two nozzles and a rotating toothed wheel blocking the nozzles in an alternate fashion. This device was designed to obtain the highest possible levels of transverse oscillation in the MIC. After a brief review of the VHAM, this article reports cold flow experiments using this modulator. Velocity maps obtained under resonant conditions using the VHAM are examined at different instants during a cycle of oscillation. Experimental data are compared with numerical pressure and velocity fields obtained from an acoustic solver. The good agreement observed in the nozzle vicinity indicates that numerical simulations can be used to analyze the complex flow field generated by the VHAM. To cite this article: Y. Mery et al., C. R. Mecanique 337 (2009).

Propulsão líquida no IAE: visão das atividades e perspectivas futuras

This paper presents the activities in the area of liquid propulsion, which received an effective boost at the end of the 90s. There is a consensus among specialists on the use of liquid propulsion in the next generation of satellite launcher vehicles in order to increase significantly both the performance and the satellites insertion accuracy. Several activities in this area are in development, among which should mention, the L5, the L15 and the L75 LPRE, the specifications for engine testing facilities for components and rocket stages, and a human resources program for training in the areas of liquid rocket engine design, manufacture and testing.

Cylindrical Shell Submerged in Bounded Acoustic Media: A Modal Approach

The dynamics of a simply-supported cylindrical shell submerged in liquid hydrogen (LH 2 ) and liquid oxygen (LOX) are considered. The shell itself is bounded by a rigid outer cylinder with closed rigid ends. This conOguration gives rise to two ∞uid-Olled cavities{ an inner cylindrical cavity and an outer annular cavity. Such geometries are common in rocket engine design. This study computes the natural frequencies and modes of the ∞uid-structure system by combining the rigid wall acoustic cavity modes and the in vacuo structural modes into a system of coupled ordinary diAEerential equations. Eigenvalue veering is observed near the intersections of the curves representing natural frequencies of the rigid wall acoustic and the in vacuo structural modes. In the case of a shell submerged in LH 2 , system frequencies near these intersections are as much as 30% lower than the corresponding in vacuo structural frequencies. Due to its high density, the frequency reductions in the presence of LOX are even more dramatic. The forced response of the ∞uid-loaded shell subject to a harmonic point excitation is also presented. The forced response in the presence of ∞uid is diAEerent from the response of the structure in vacuo in a variety of ways. The frequency shifts that arise from consideration of the ∞uid alter the order of the resonant response peaks. In some cases, modes that are well separated in the in vacuo case are within close proximity in the ∞uid-loaded case (and vice-versa). The ∞uid-loaded structural responses also contain relatively small resonant peaks corresponding to system modes that are dominated by contributions from the ∞uid.

メタン燃料を用いたロケットエンジンノズル性能特性

Performance of the methane fueled rocket nozzles are numerically investigated using computational fluid dynamics approach. A simple set of chemical reactions and kinetics for methane/oxygen nozzle flow is proposed. The chamber pressure, mixture ratio and size of the nozzle are parametrically changed to study the influence of characteristic rocket engine design parameters on nozzle losses. The amount of dissociation is high when the chamber pressure is low and the kinetic loss becomes dominant compared to the other nozzle losses. The peak specific impulse is achieved at a higher mixture ratio region as the chamber pressure increases. The chemical non-equilibrium flow appears mainly at down stream region of the nozzle throat. The influence of the chemical non-equilibrium effect decreases as the chamber pressure increases. Supersonic chemically reactive gas stays longer in the nozzle as the size of the nozzle become larger and the amount of recombination increases which decreases the kinetic loss. When the chamber pressure is high, the kinetic loss becomes small and the effect of the size of nozzle also becomes small.

액체로켓 노즐의 열화학적 성능 해석

로켓 엔진 노즐의 설계에서 동결 유동 해법과 동일한 수치적 특성을 가지는 화학 평형 해석은 노즐의 열역학적 최대 성능을 예측하는 효율적인 설계 도구로 이용될 수 있다. 본 연구에서는 탄화수소 연료 로켓 엔진 설계를 위한 화학적 평형 유동 해석 코드를 개발하였다. 로켓 노즐을 통한 팽창과정에서 일어나는 화학 성분의 재결합 효과와 이에 수반하는 에너지 회복과 같은 열화학적 특징을 이해하기 위하여, KSR-III 로켓 노즐에 대하여 동결유동 해석 및 비평형 유동의 해석과 더불어 화학적 평형 유동 해석을 수행하였다. 유동 해석 결과에 기초한 KSR-III 엔진 성능 평가로부터 노즐에서의 열화학적 특징을 이해할 수 있었으며, 이와 더불어 열화학적인 효과를 고려할 때 출구 면적비를 줄여서 수정된 새로운 노즐 형상이 지상 추력을 증대시키기 위한 적절한 설계임을 확인할 수 있었다. For a design of rocket engine nozzle, chemical equilibrium analysis which shares the same numerical characteristics with frozen flow analysis can be used as an efficient design tool for predicting maximum thermodynamic performance of the nozzle. 10 this study, a chemical equilibrium flow analysis code was developed for the design of hydrocarbon fueled rocket engines. 10 oder to understand the thermochemical characteristics occurring in a nozzle through the expansion process, such as recombination of chemical components and the accompanying energy recovery, chemical equilibrium flow analysis was carried out for the KSR-III rocket engine nozzles together with frozen flow and non-equilibrium flow analyses. The performance evaluation based on the present KSR-III nozzle flow analyses has provided an understanding of the thermochemical process in the nozzle and additionally, it has confirmed that the newly designed nozzle shape modified to have a reduced exit area ratio is an adequate design for obtaining an increased ground thrust.