Rocket Combustion Chamber Research Articles

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.

POD에 의한 하이브리드 로켓 연소실의 유동특성 해석

하이브리드 로켓 연소실 내부 유동장에 대한 수치계산 결과에 POD기법을 적용하였다. 특히, 다이어프램 설치에 따른 유동모드 변화를 분석하여, 연소특성에 미치는 영향을 해석하였다. 또한, 다이어프램이 있는 연소실에서 표면 분출유동의 유무에 따른 POD를 적용하여 분출유동이 연소실 내부 유동특성에 미치는 영향을 판단하였다. 10개의 모드를 사용하여 기본형상에 대한 POD 결과를 살펴보면 주 유동을 나타내는 모드 1과 벽면 근처의 작은 크기 유동인 2-9 모드 사이의 구분이 분명하게 나타났다. 다이어프램을 설치한 형상의 POD 결과, 모드 2부터 5의 에너지가 증가하였는데 이것은 다이어프램 주변 순환영역에서 생성되는 유동 때문인 것으로 보인다. 한편, 다이어프램 주위 영역의 유동특성을 보여주는 모드 2-5와 후류 벽면의 유동특성을 보여주는 모드 6-9의 에너지 분포가 분출유동 유무에 관계없이 비슷한 특성을 나타냈다. 따라서 연소율이 다이어프램 근처에만 국부적으로 증가하는 이유는 다이어프램 후류에 형성되는 비교적 큰 크기의 유동모드 2-5의 에너지가 증가되었기 때문인 것으로 분석된다. POD analysis has been done to investigate the internal flow characteristics using LES calculation results of hybrid rocket combustion chamber. The special emphasis was put on the change in the mode energy distribution caused by the installation of diaphragm compared to the baseline case. Also the comparison was made to investigate the effect of wall blowing on the changes in the mode energy between the regions near and far from the diaphragm. For baseline case, POD results clearly distinguish the primary mode containing most of flow energy from the rest of flow modes (2-9 mode) depicting small scale modes. Also, the increase in the energy of flow modes 2-5 is responsible for the formation of relatively large scale structures due to diaphragm. In addition, the comparison of mode energy distributions of flow fields with diaphragm shows similar patterns in both wall blowing and no blowing case. This implies that the local increase in regression rate just after the diaphragm is directly associated with the increase in energy distributions of 2-5 modes.

A pressure-based treatment for the direct numerical simulation of compressible multi-phase flow using multiple pressure variables

In this paper we present a pressure-based numerical scheme for the direct numerical simulation of compressible two-phase flows using the stiffened gas equation of state. While for many technical applications, two-phase flows can be treated as incompressible, this assumption fails in cases with high pressure and temperature as they can be found in rocket combustion chambers, for example. Our interest is in the development of a pressure-based method that aims at the extension of an incompressible two-phase code to the compressible regime. The development builds upon an asymptotic pressure decomposition using multiple pressure variables and has originally been designed for single-phase flows. Its adaptation to compressible two-phase flows is presented. This includes the possibility to resolve and track the interface as well as the description of the two phases by different equations of state. It is shown that the pressure-based scheme does not necessitate a cumbersome interface treatment in order to avoid spurious oscillations in the vicinity of the material interface. We do not yet take into account phase changes whose approximation requires a careful thermodynamic consistent procedure. Numerical examples are shown ranging from the one-dimensional transport of a multi-material contact discontinuity to the three-dimensional simulation of shock-droplet interactions. The scheme proves to be able to accurately simulate the propagation of pressure waves in gaseous and liquid phases.

Radiative heat transfer analysis in modern rocket combustion chambers

Radiative heat transfer is analyzed for subscale and fullscale rocket combustion chambers for H2/O2 and CH4/O2 combustion using the P1 radiation transport model in combination with various Weighted Sum of Gray Gases Models (WSGGMs). The influence of different wall emissivities, as well as the results using different WSGGMs, the size of the combustion chamber and the coupling of radiation and fluid dynamics, is investigated. Using rather simple WSGGMs for homogeneous systems yields similar results as using sophisticated models. With models for nonhomogeneous systems the radiative wall heat flux (RWHF) decreases by 25–30 % for H2/O2 combustion and by almost 50 % for CH4/O2 combustion. Enlarging the volume of the combustion chamber increases the RWHF. The influence of radiation on the flow field is found to be negligible. The local ratio of RWHF to total wall heat flux shows a maximum of 9–10 % for H2/O2 and 8 % for CH4/O2 combustion. The integrated heat load ratio is around 3 % for H2/O2 and 2.5 % for CH4/O2 combustion. With WSGGMs for nonhomogeneous systems, the local ratio decreases to 5 % (H2/O2) and 3 % (CH4/O2) while the integrated ratio is only 2 % (H2/O2) and 1.3 % (CH4/O2).

Numerical Simulation of Ionized Rocket Plumes

In this paper, a multiphysics numerical approach for predicting the ionization level in solid rocket engine plumes is presented. Ionization takes place in the rocket combustion chamber and in the exhaust plume. A low-temperature, high-density plasma is created, with small Debye length in most of the plume region. Ambipolar diffusion is therefore assumed for ions and electrons in the plume, and a set of conservation equations is derived to be solved by the numerical model. A number of numerical strategies to resolve this system is derived, as well as a novel scheme that enforces charge neutrality. This approach is integrated into a complex code for compressible, multispecies, turbulent flow simulations. The model is then coupled with a Maxwell’s equations solver in order to simulate the radar cross section of rocket plumes. Finally, computations of ionization levels and radar cross section of a Black Brant rocket plume are presented.

Large-eddy simulation of supercritical fluid injection

Mixing is a key point for thrust and efficiency of combustion systems. It becomes crucial in the case Liquid Rocket Engines as large investments are involved. Besides, the pressure in liquid rocket combustion chamber often exceeding the critical point of loaded propellants, mixing becomes an important scientific issue as fluid properties differ from classical ideal gas assumption. In this study, two configurations are studied to evaluate the impact of subgrid models on mixing. Firstly, Mayer's experiments of trans- and super-critical nitrogen jet injection into a warm nitrogen atmosphere have been numerically investigated with a structured numerical code called SiTCom-B. SiTCom-B solves Direct Numerical Simulations and Large Eddy Simulations equations for perfect or real gas equation of states. In this study, Soave–Redlich–Kwong (SRK) and Peng–Robinson equation of state are used with appropriated thermodynamics relations and validated against NIST data. Three-dimensional LES are conducted for two cases (cases 3 and 4 in Mayer et al.'s reference [1]) with real-gas NSCBC treatment. Several sub-grid scale models are tested and the results are compared to experimental data for density on jet axis: a very good agreement is obtained on a light mesh (11.6 million of points) with the SRK equation of state and standard Smagorinsky model. Flow structures are evidenced with Schlieren snapshots. Secondly, the Mascotte test-bench from ONERA is simulated with SiTCom-B based on Soave–Redlich–Kwong equation of state and Smagorinsky models. The simulated non-reacting case is characterized by a very short liquid oxygen dense core because of the strong density and velocity gradients boosting mixing efficiency.

Role of gaseous film cooling injector orientation in duct flow: Experimental Study

Experimental investigations are carried out to analyze the effect of coolant injector configuration on overall film cooling performance in a cylindrical test section similar to a rocket combustion chamber. Three different injector orientations are investigated: (i) holes parallel to core gas flow, (ii) holes at a tangential angle to the core gas flow, and (iii) compound angle holes inclined to the wall both in the tangential and azimuthal direction. The objective is to provide a consistent set of measurements for cooling effectiveness within the same test facility with respect to different blowing ratios ranging from 1.69 to 3.9. The results suggest that coolant injection parallel to the axis should be used in situations where far field effectiveness is of concern. The tangential injection, in general, shows lower wall protection compared to other schemes. The results show that the compound angle configuration augments film cooling effectiveness in the near injection regimes. However comparison of the effectiveness values for the compound injectors suggests the existence of an optimum compound angle configuration. The circumferential wall temperature data represent systematic changes in the coolant flow behavior with the change from straight injection to compound angle injection.

Similarity solution of fuel mass transfer, port mass flux coupling in hybrid propulsion

Boundary layer combustion is the primary mechanism of hot gas generation in hybrid rockets. The idea of a hybrid rocket is to store the oxidizer as a liquid and the fuel as a solid, producing a design that is less susceptible to chemical explosion than conventional solid and bipropellant liquid designs. The fuel is contained within the rocket combustion chamber in the form of a cylinder, with a circular channel called a port hollowed out along its axis. Upon ignition, a diffusion flame forms over the fuel surface along the length of the port. The combustion is sustained by heat transfer from the flame to the solid fuel causing continuous fuel vaporization until the oxidizer flow is turned off. Theory shows that the fuel mass transfer rate is proportional to the mass flux averaged across the port. The mass flow rate increases with axial distance, leading to coupling between the local fuel regression rate and the local mass flux. For proper design, accurate expressions are needed for both the time-dependent oxidizer-to-fuel ratio at the end of the port and the time at which all the fuel is consumed. As the fuel is depleted, the flame approaches the motor case, at which point the burn must be terminated. The purpose of this paper is to present a similarity solution for the coupled problem and a brief discussion of current practice.

Analysis and design optimization of solid rocket motors in viscous flows

The objective of this study is to develop a design tool that can be used in viscous flows. The flow analysis is based on the axisymmetric Navier–Stokes and k–ɛ turbulence equations. These coupled equations are solved using an explicit finite difference method. The accuracy of the analysis code is validated for viscous flows in solid rocket motor combustion chamber and nozzle. The gradient-based numerical optimization model is used to maximize the thrust of solid rocket motor under a constraint of propellant weight. The sensitivity analysis that measures the response of the flow with respect to a geometry perturbation is calculated by finite differencing. The optimization of design study employs a commercial optimization package. The performance of design optimization method is tested in solid rocket motor combustion chamber and nozzle design.

Investigation on the Ignition Over Pressure Related to Launch Vehicle Lift-of.

Several transient events occur during the startup of the Solid Rocket Motor (SRM) during the launch vehicle lift-off. Each event produces a complex transient signal and requires systematic assessment. The event discussed in this paper is Ignition Over Pressure (IOP). This Ignition Over Pressure is resulting from fluid dynamic compression of the accelerating plume gas, subsequent rarefaction and propagation during the pressure rise rate period in the rocket combustion chamber. These high-amplitude unsteady fluid-dynamic perturbations can adversely affect the vehicle and its surrounding structure. This wave behaves as a blast or shock wave characterized by a positive triangular shaped first pulse and negative half sine wave second pulse. The pulse travels upwards towards the propulsion system and has the potential to overload the individual elements or exciting overall vehicle dynamics. The later effect results from the phase difference of the wave from one side of the vehicle to the other due to the skew in the ignition of the strap-on boosters. In the case of the launchers, the mechanical stress due to the ignition over pressure wave comes, in addition to the acoustic constraint due to the jet noise. The over pressure phasing or ΔP environment, because of its spectral content as well as amplitudes becomes a crucial input for the design of sub-assemblies viz., thermal shields, and pay loads etc. In this paper, an attempt is made to numerically visualize the propagation of the blast wave causing the unsteady pressure oscillations during the transient pressure rise in the combustion chamber. To understand further, pressure measurements have been made at different heights along the umbilical tower during the solid rocket motor lift-off with the presence of jet deflectors, in order to capture the shock propagation phenomena. From the results, it appears that a sudden shock front that is generated during flow development within the rocket nozzle leaves a clear signature in the form of a well defined peak at typical time intervals. This overpressure amplitude seems to be related to the slope of the combustion chamber pressure rise rate during the transient period. It is also observed that the shock front initially propagates at supersonic speed but decays with time later. Similarly, the magnitude of the IOP peaks are also seen to decay with distance.

Investigation on the Ignition Over Pressure Related to Launch Vehicle Lift-off

Several transient events occur during the startup of the Solid Rocket Motor (SRM) during the launch vehicle lift-off. Each event produces a complex transient signal and requires systematic assessment. The event discussed in this paper is Ignition Over Pressure (IOP). This Ignition Over Pressure is resulting from fluid dynamic compression of the accelerating plume gas, subsequent rarefaction and propagation during the pressure rise rate period in the rocket combustion chamber. These high-amplitude unsteady fluid-dynamic perturbations can adversely affect the vehicle and its surrounding structure. This wave behaves as a blast or shock wave characterized by a positive triangular shaped first pulse and negative half sine wave second pulse. The pulse travels upwards towards the propulsion system and has the potential to overload the individual elements or exciting overall vehicle dynamics. The later effect results from the phase difference of the wave from one side of the vehicle to the other due to the skew in the ignition of the strap-on boosters. In the case of the launchers, the mechanical stress due to the ignition over pressure wave comes, in addition to the acoustic constraint due to the jet noise. The over pressure phasing or ΔP environment, because of its spectral content as well as amplitudes becomes a crucial input for the design of sub-assemblies viz., thermal shields, and pay loads etc. In this paper, an attempt is made to numerically visualize the propagation of the blast wave causing the unsteady pressure oscillations during the transient pressure rise in the combustion chamber. To understand further, pressure measurements have been made at different heights along the umbilical tower during the solid rocket motor lift-off with the presence of jet deflectors, in order to capture the shock propagation phenomena. From the results, it appears that a sudden shock front that is generated during flow development within the rocket nozzle leaves a clear signature in the form of a well defined peak at typical time intervals. This overpressure amplitude seems to be related to the slope of the combustion chamber pressure rise rate during the transient period. It is also observed that the shock front initially propagates at supersonic speed but decays with time later. Similarly, the magnitude of the IOP peaks are also seen to decay with distance.

Development of eddy current testing system for inspection of combustion chambers of liquid rocket engines

An eddy current testing (ECT) system using a high sensitive anisotropic magnetoresistive (AMR) sensor was developed. In this system, a 20 turn circular coil with a diameter of 3 mm was used to produce the excitation field. A high sensitivity AMR sensor was used to measure the magnetic field produced by the induced eddy currents. A specimen made of copper alloy was prepared to simulate the combustion chamber of liquid rocket. Scanning was realized by rotating the chamber with a motor. To reduce the influence of liftoff variance during scanning, a dual frequency excitation method was used. The experimental results proved that ECT system with an AMR sensor could be used to check liquid rocket combustion chamber.

액체로켓엔진의 연소가스와 액체질소 혼합에 의한 연소 가스 냉각 특성에 관한 연구

본 연구에서는 액체산소와 케로신을 추진제로 사용하는 액체로켓엔진의 연소 가스에 액체질소를 분사하여 연소 가스의 냉각 특성을 알아보고자 하였다. 일반적인 액체로켓엔진의 연소실 후단에 액체질소 분사기와 혼합용 연소실을 추가적으로 장착하였고, 혼합용 연소실 후단에 노즐을 연결하여 전체적인 시스템을 구성하였다. 연소실험은 안전을 고려하여 점화실험부터 순차적으로 수행하였으며, 최종적으로 20초 연소실험을 수행하였다. 그 결과 기존의 액체로켓엔진에 액체질소를 분사함으로써, 연소 가스의 온도를 저하시킬 수 있음을 확인하였다.

A new generalised model for liquid film cooling in rocket combustion chambers

A one dimensional analytical model of liquid film cooling in rocket combustion chambers operating at subcritical conditions is developed. The approach followed involves the selection of a control volume for mass and energy balance. The coolant evaporation rate per area is obtained from this energy balance. The present model incorporates mass transfer via entrainment by adapting suitable correlations from literature pertaining to annular flow conditions. The model predicted favourably with the experimental data available in open literature and produced superior results compared to all existing models. Results are presented for a mixed gas–water system under different conditions. Results indicate that convection dominates the heat transfer at the gas–liquid interface. Effects of gas Reynolds number, coolant inlet temperature, combustion chamber pressure, mass flow ratio of the liquid coolant to the free stream and the free stream turbulence on the liquid film length are presented in detail.

A0 JAXA角田宇宙センターにおける宇宙輸送系研究開発活動(宇宙輸送・飛翔体)

This paper presents activities on the rocket engines and others at JAXA Kakuda Space Center. The center has several kinds of test facilities. Research and development activities and acceptance tests are carried out with these facilities. Studies on the rocket engine turbopump, combustion chamber and nozzle are carried out. Development tests of LE-X hydrogen turbopump are planned at the center. Studies on the rocket-based combined-cycle engine are carried out for future transportation system. Re-entry technologies are also studied at the center. Kakuda Space Center has exhibition hall and accepts visitors. There are open-to-public days in spring and autumn, respectively. The center conducts all activities with attention to safety, and contributes to activities of decreasing stress to the environment.

<i>In Situ</i> Synthesis of Cr<sub>2</sub>Nb Reinforced Copper Alloy by Liquid Metallurgy Route

Cu-8Cr-4Nb alloy with high strength and high thermal conductivity have received a lot of attention over the last decades due to promising role in space application of rocket combustion chamber of reusable launch vehicle. In the present work, the copper alloy is synthesized in induction furnace by liquid metallurgy route wherein elemental powders of chromium and niobium was used added into premelted copper melt under argon atmosphere. The cast product is characterised for hardness, optical, SEM, EDS and XRD. It is possible to produce this alloy with reproducible average hardness of 135 VHN. The precipitate of Cr2Nb showed moderate distribution in and around the grain boundaries of copper with precipitate size varying from 1.33 μm to7.0 μm.

Single-stage conversion of associated petroleum gas and natural gas to syngas in combustion and auto-ignition processes

Single-stage conversion of alkane mixtures simulating associated petroleum gas (APG) to syngas is studied in a static installation and in a flow reactor based on the rocket combustion chamber. Yields of the desired reaction products close to their thermodynamically equilibrium values are obtained. A range of experimental parameters, in which ignition delays of APG-oxygen mixtures exhibit negative or zero temperature coefficients, is determined for the first time. Such a behavior of ignition delays is proved to be a fundamental property of fuel-rich APG mixtures. The range of abnormal temperature dependence of ignition delays is shown to be extended as the initial pressure rises, which makes it possible to significantly increase the reaction rate by increasing the initial working pressure.

Numerical analysis of regenerative cooling in liquid propellant rocket engines

High combustion temperatures and long operation durations require the use of cooling techniques in liquid propellant rocket engines (LPRE). For high-pressure and high-thrust rocket engines, regenerative cooling is the most preferred cooling method. Traditionally, approximately square cross sectional cooling channels have been used. However, recent studies have shown that by increasing the coolant channel height-to-width aspect ratio and changing the cross sectional area in non-critical regions for heat flux, the rocket combustion chamber gas-side wall temperature can be reduced significantly without an increase in the coolant pressure drop. In this study, the regenerative cooling of a liquid propellant rocket engine has been numerically simulated. The engine has been modeled to operate on a LOX/kerosene mixture at a chamber pressure of 60 bar with 300 kN thrust and kerosene is considered as the coolant. A numerical investigation was performed to determine the effect of different aspect ratio and number of cooling channels on gas-side wall and coolant temperatures and pressure drop in cooling channels.

Development of In-Situ Synthesis of Cr2Nb Reinforced Copper Alloy by Aluminothermic Process

Cu-8 (by %at) Cr-4 (by %at) Nb alloy with high strength and high thermal conductivity have received a lot of attention over the last decades due to promising role in space application of rocket combustion chamber of reusable launch vehicle. In the present work, the copper alloy synthesize in induction furnace by in situ reduction of Nb2O5 by aluminothermic process in presence of Cr in the premelted copper. The resultant cast product was characterised for hardness, optical, SEM, EDS and XRD analysis. It is possible to synthesise copper alloy with reproducible microstructures and an average hardness of 134 VHN. The precipitates of Cr2Nb showed moderate distribution in and around the grain boundaries of copper with precipitates size varying from 0.7 to 2.5 μm.

Simulation of a GOX–kerosene subscale rocket combustion chamber

In view of future film cooling tests at the Institute for Flight Propulsion (LFA) at Technische Universitat Munchen, the Astrium in-house spray combustion CFD tool Rocflam-II was validated against first test data gained from this rocket test bench without film cooling. The subscale rocket combustion chamber uses GOX and kerosene as propellants which are injected through a single double swirl element. Especially the modeling of the double swirl element and the measured wall roughness were adapted on the LFA hardware. Additionally, new liquid kerosene fluid properties were implemented and verified in Rocflam-II. Also the influences of soot deposition and hot gas radiation on the wall heat flux were analytically and numerically estimated. In context of reviewing the implemented evaporation model in Rocflam-II, the binary diffusion coefficient and its pressure dependency were analyzed. Finally simulations have been performed for different load points with Rocflam-II showing a good agreement compared to test data.