Thrust Chamber Research Articles

Two case studies on acoustic characteristics in combustion chamber using Helmholtz-type acoustic cavity: FASTRAC and KSR-III combustors

This paper presents two case studies to evaluate passive stabilization in combustion chambers by quantifying damping capacity characteristics of acoustic cavities in combustion chambers using a three-dimensional linear acoustic analysis that has been applied in a previous study. A feasible interpretation was performed for a FASTRAC thrust chamber, where both successful and failed cases were reported. Regarding the temperature and sonic velocity in an acoustic cavity, the possibility of fine-tuning is evaluated in terms of acoustic absorption and conductance characteristics using the acoustic impedance concept. For a given acoustic cavity, the sonic velocity in the cavity required to achieve an optimal tuning is determined, which agrees with previous experimental results. The results show that the adopted cavities are not tuned appropriately to the frequency of combustion instabilities. Hence, it can be concluded that using an acoustic cavity with a small volume or an orifice having a large cross-sectional area can result in sufficient acoustic damping and the suppression of instabilities. In addition, the feasible estimation of sonic velocity in the acoustic cavity is proposed to ensure sufficient acoustic absorption. The results show similar trends without significant deviations, which can be applied in the design and verification of an acoustic cavity. From the present results, a preliminary fine-tuning of the acoustic cavity can be performed using this pure acoustic analysis.

Measuring the velocity field of a shear-coaxial, cryogenic flame in a high-pressure rocket thrust chamber

High-speed imaging was used to visualize one of the transcritical flames in a multi-injector, sub-scale rocket thrust chamber at pressures up to 80 bar. Image correlation velocimetry (ICV) was applied to the imaging to obtain quantitative information on the flow field from the shear-coaxial injectors. ICV was used to track surface irregularities on the liquid oxygen (LOX) jet in imaging filtered to blue wavelengths. By choosing the interrogation area carefully, only the LOX jet was effectively tracked, excluding the coaxial H2 flow, and the time-averaged velocity field was reconstructed. Due to the transient nature of the tracked features, which frequently change shape or disappear, the averaged ICV result underestimates the absolute values of velocity. Therefore, the averaged values were scaled by the mean of the instantaneous velocity maxima. A second, reference measurement of LOX jet propagation speed was calculated using dynamic mode decomposition (DMD). The results were consistent following the aforementioned correction of the ICV values. Comparing the ICV fields for two different operating conditions showed a marked difference in the axial velocity distribution and lateral expansion of the LOX jet, demonstrating the potential of the method in studying injection in rocket combustion chambers.

Numerical evaluation of acoustic characteristics of a thrust chamber with quarter-wave resonators

Acoustic characteristics of a thrust chamber with quarter-wave resonators are numerically studied based on the unsteady Reynolds-averaged Navier-Stokes (URANS) method. Organized pressure disturbance model and constant-volume bomb model are applied as artificial disturbances to excite pressure oscillations in the chamber. Eigenfrequencies and amplitudes of acoustic modes of the chamber are obtained by fast fourier transform (FFT) analysis, while damping characteristics are evaluated by the half-power bandwidth method. Predicted damping capacities of the chamber with and without quarter-wave resonators agree well with experimental results. Pressure oscillations can be controlled by a quarter-wave resonator mainly through reducing the amplitude of target acoustic mode, rather than increasing damping capacity of the chamber. Major damping mechanism of the resonator is cutting down pressure peak of target acoustic mode and raising up its pressure trough (CPRT); therefore the amplitude of target acoustic mode is reduced significantly. Moreover, acoustic energy can be dissipated by vortex at the orifice and by viscosity on the surface of a resonator, which increase damping capacity of the chamber slightly. Under the condition with multi-modes pressure oscillations, a resonator can still suppress pressure oscillations of target acoustic mode through CPRT. However, it may enhance pressure oscillations of other modes due to redistribution of oscillation energy among all acoustic modes.

Flame-acoustic response measurements in a high-pressure, 42-injector, cryogenic rocket thrust chamber

This work presents measurements of acoustically driven flame dynamics in a 42-element, cryogenic oxygen-hydrogen rocket thrust chamber under supercritical injection conditions. The experiment shows self-excited combustion instabilities for certain operating conditions, and this work describes the nature of the flame dynamics driving the acoustic field, as far as it can be ascertained from state-of-the-art optical measurements. Optical access has been realized in the combustion chamber with both fibre-optical probes and a viewing window. The probes collect point-like measurements of filtered OH* radiation. Their signals were used to calculate the gain and phase of intensity oscillations with respect to acoustic pressure for both stable and unstable operating conditions. Through the window, synchronized high-speed imaging of the flame in filtered OH* and blue radiation wavelengths was collected. The 2D flame response was related to the local acoustic pressure to investigate the distributed intensity and phase relationships. The flame response from OH* measurements is in agreement with the theory of Rayleigh. For stable conditions the oscillations of combustion and pressure were out of phase, whereas for an excited chamber 1T mode the oscillations were closely in phase. The integrated Rayleigh index from blue imaging was not consistent with the OH* results. The reason lies in the depth of field captured by this type of imaging, and must be used in a complementary fashion together with OH* imaging. The flame response values and 2D visualization presented in this work are expected to be of value for the validation of numerical modelling of combustion instabilities.

Dynamic reliability analysis for the reusable thrust chamber: A multi-failure modes investigation based on coupled thermal-structural analysis

To evaluate the dynamic reliability of the thrust chamber in the reusable rocket engine (RRE) under multi-failure modes, a reliability analysis based on the coupled thermal-structural finite element model and the multiple response surface (MRS) is proposed. Through the deterministic thermal-structural simulation, two major failure modes, the static strength failure and cyclic cumulative damage, are examined. While the random variables are introduced concerning working loads, geometrical dimensions, and material properties. The reliability of a typical reusable LOX/H2 rocket engine thrust chamber is examined by the proposed method. The feasibility of the proposed method is validated by comparing our results with that from the traditional Monte Carlo simulation and the influences of sampling methods as well as set sizes are investigated. This study provides an efficient and steadfast approach to evaluate the reliability of thrust chambers in the reusable rocket engine and quantifies the importance ranking of working loads, geometrical dimensions, and material properties for the chamber reliability, which provides a valuable insight into the reliability-based design and optimization process for reusable rocket engines

Numerical Simulation of Shock Wave-Boundary Layer Interaction with Supersonic Film Cooling

As an effective external cooling method, film cooling can significantly reduce the wall temperature of hot end components of aero engines and large liquid rocket engines. With the development of supersonic turbine blade, liquid rocket engine thrust chamber and ramjet thermal protection technology, the impact of shock wave-boundary layer interaction on the cooling effectiveness of film cooling is a problem that must be considered in the engine heat transfer design. In this paper, the two-dimensional supersonic film cooling is simulated based on the plane slot model, and the blowing ratio is similar to the actual engine. Impact of wave structures on film cooling was studied, and flow loss was evaluated by total pressure and entropy parameter to investigate the film cooling efficiency and uniformity distribution. Physical mechanism of shock wave-boundary layer interaction and its impact on film form was revealed. It shows that the shock wave leads to dramatic changes in flow parameters. Boundary layer separation is found due to the shock, and significantly affects the cooling efficiency. The blowing ratio is an important factor affecting the cooling efficiency. With blowing ratio increasing, local film cooling effectiveness and average cooling efficiency are improved and cooling uniformity decreases. However, the shock wave will cause the temperature of the insulation wall to rise, resulting in decreased cooling efficiency.

Numerical Evaluation of the Effects of Inclusions on Solid Rocket Motor Performance

This work describes the application of original computer graphics methods to examine the effect of inclusions within the solid rocket motor thrust chamber. The adoptance of a dynamic three-dimensional triangular mesh with a self-intersection removal algorithm, as the core of the burning surface regression module, allows the aforementioned target to be reached. Many phenomena not measurable in an analytical closed form, like different-shaped voids within the grain, can be investigated. Indeed, both local heterogeneities and surface imperfections could generate unexpected variations in the combustion chamber pressure pattern and early thermal protections exposure. The aforementioned techniques are developed and tested on a ballistic simulator developed using computer graphics methods to improve mesh handling. Simulation results are carried out and discussed.

Wall heat transfer prediction in CH4/O2 and H2/O2 rocket thrust chambers using a non-adiabatic flamelet model

The current work presents the extension of the flamelet model for turbulent combustion calculations to account for deviations from adiabatic conditions. The aforementioned extension is expected to significantly improve the prediction of the chemical processes occurring in the vicinity of cooled walls in rocket engine applications. A lower enthalpy level leads to an increase of the recombination reactions, which is of particular interest in the case of methane/oxygen combustion. In the present approach, the flamelet equations are solved in mixture fraction space and the energy equation is replaced by a prescription of the enthalpy profile in order to include non-adiabatic effects. To avoid the over-prediction of the recombination reactions, a local ”freezing” of the chemical reactions is introduced based on the Damkoehler number close to the cold wall boundary. A pre-tabulation of the chemical time-scales in the flamelet tables enables a fast calculation of the Damkoehler number. The model is verified both for CH4/O2 and H2/O2 using the simulation of a cooled reacting boundary layer. The extended hybrid model is employed for the simulation of a single-element rocket thrust chamber using CH2/O2 and H2/O2 and is compared to the non-adiabatic and frozen flamelet models. A more accurate wall heat transfer and pressure level prediction is achieved with the hybrid model for both propellant combinations leading to great agreement with the available experimental measurements.

Analysis of plasma detachment in the magnetic thrust chamber using full particle-in-cell simulation

A magnetic nozzle, which is a convergent-divergent magnetic field to control a plasma flow, has been investigated for application to plasma propulsion systems in spaceships. In the magnetic nozzle, plasma thermal energy is converted to one-directional kinetic energy by Lorentz force to generate thrust. Although magnetic field structure and strength are optimized for improvement of the thrust performance, it is essential to understand physical processes of plasma ejection from the nozzle, because the plasma may flow back along lines of magnetic field if the directed plasma continues being strongly magnetized. It is assumed, for one of the scenarios to explain a plasma detachment, that a plasma detaches from magnetic field when a cyclotron radius exceeds a scale length of magnetic field in size. Hence, we investigate a plasma detachment condition by analyzing parameters of the plasma for unmagnetization. We assumed individual particle motion of a fully ionized plasma in the magnetic nozzle and conducted a full particle-in-cell simulation in a two-dimensional coordinate system. We calculated a ratio of cyclotron radius to a scale length of the magnetic field. The ratio increased in a downstream due to variation of magnetic field induced from a diamagnetic cavity, suggesting unmagnetization.

Effect of coaxial injector parameters on LOX/methane engines: A numerical analysis

The liquid oxygen/methane engine is a desirable power for reusable launch vehicle. It is generally believed that the efficiency of the liquid oxygen/methane engine is mainly influenced by the injection conditions of the thrust chamber. However, due to the complexity of combustion experiments, the exact influence of specific injectors on the combustion performance is difficult to acquire. By analyzing a typical sub-scale thrust chamber, this paper presents a three-dimensional numerical model to evaluate the effects of three major injector parameters, i.e. the thickness of liquid oxygen post, the number of injectors and the type of injectors, on liquid oxygen/methane engine combustion performance. The results of the simulation are consistent with existing experimental data in the chamber pressure, combustion stability, as well as wall heat flux. This paper's results have shown that the larger liquid oxygen post thickness and swirl coaxial injector can significantly reduce the heat flux while influence of the injector number is relatively small. Furthermore, the decrement of the liquid oxygen post thickness can improve the combustion efficiency. And the less injectors lead to the more stable combustion. The presented simulation model gives a reliable and cost-effective way to analyze specific injector parameters and provides new principles for future liquid oxygen/methane engine design.

Damping device to reduce the risk of injection-coupled combustion instabilities in liquid propellant rocket engines

A new countermeasure against injection-coupled combustion instabilities in liquid propellant rocket engines is presented. Whereas the problem is usually addressed by adding damping elements such as baffles or resonators to the combustion chamber, this approach directly damps the acoustic eigenmodes of the injector instead. The principle of the damping method is described in this article, as well as the implementation of such a device in a sub-scale rocket thrust chamber operated with liquid oxygen and hydrogen at conditions representative of upper stage engines. Test results are presented which show that flame and pressure oscillations were successfully reduced by the modification. The absorbers had no measurable influence on thrust chamber performance, and so the solution lends itself to retrofitting in existing engines, as well as integration during the design phase.

Injector-coupled thermoacoustic instabilities in an experimental LOX-methane rocket combustor during start-up

This paper reports the investigation of acoustic combustion instability experienced during repetitive ignition testing of a sub-scale LOX-methane rocket thrust chamber. The occurrence of resonant coupling between the LOX injectors and the combustion chamber acoustic modes was assessed from the experimental data recorded during the highly transient phase of operation from ignition up to around 2 s. A method was developed to model the evolution of acoustic properties in both the combustion chamber and the injectors during the transient period. For the LOX injectors, the Woods equation was used to estimate the speed of sound in the two-phase flow. The models were used to identify the corresponding mode frequencies in the unsteady pressure measurements, and show that the high-amplitude instability occurred when they intersected. Very close coupling of less than 3% frequency difference is required for high amplitudes to be observed. However, the condition was necessary but not sufficient for high amplitudes to be reached.

Investigation of a Radial Turbines Compatibility for Small Rocket Engine

In the radial turbine preliminary design for an expander rocket engine, a comparison was made with axial turbine used in Pratt & Whitney RL10 engine. One of the primary requirements of a liquid propellant rocket engine is the generation of a high thrust, which depends on both the mass flow rate of the propellant and the pressure in the thrust chamber. In expander-cycle engines, which are the subject of the present study, the liquid propellant is first compressed using centrifugal turbo-pumps, then it is used to cool the combustion chamber and the nozzle and, once vaporized, it flows through the turbines used to drive the turbo-pumps. The aim was to demonstrate the greater efficiency of the radial turbine with a reduction of the pressure ratio with respect to the axial turbine.

Design of the thrust chamber: dimensional analysis of the combustion chamber and the nozzle of rocket engine using LOX/LCH4 propellants

This paper presents a preliminary dimensional study of small thrust chamber using LOX/LCH4 as propellants. In this study, the mixture ratio of the propellants, the temperature and the pressure at the combustion chamber are used to determinate the geometric parameters of the nozzle and the combustion chamber such as the volume, surfaces, lengths, diameters, etc. our work is based on the most important required output of the propulsion engine which is the thrust force. The results of our work provides the flow areas at the exit of the nozzle and at the throat for different values of the thrust force (10-20-30 KN) and different mixture ratio of the propellants (2.7-2.8-2.9). Moreover, the volume of the combustion chamber is calculated for different proposed characteristic length (80-160-240 cm). Our results offer decisive guidelines for the design of these key parts of the thrust chamber which is required for the design of the whole rocket engine as future work.

Numerical Analysis and Modelling of a 100 N Hypergolic Liquid Bipropellant Thruster

This study focuses on the stepwise procedure involved in the development of a numerical model of a bi-propellant hypergolic chemical propulsion system using key features and performance characteristics of existing and planned (near future) propulsion systems. The study targets specific impulse of 100 N delivery performance of thrust chambers which is suitable for primary propulsion and attitude control for spacecraft. Results from numerical models are reported and validated with the Rocket Propulsion Analysis (RPA) computation concept. In the modelling process, there was proper consideration for the essential parts of the thruster engine such as the nozzle, combustion chamber, catalyst bed, injector, and cooling jacket. This propulsion system is designed to be fabricated in our next step in advancing this idea, using a combination of additive manufacturing technology and commercial off the shelf (COTS) parts along with non-toxic propellants. The two non-toxic propellants being considered are Hydrogen Peroxide as the oxidiser and Kerosene as the fuel, thus making it a low-cost, readily available and environmentally-friendly option for future microsatellite missions.

Properties of Kerosene-Aluminium Nanofluid used to Estimate the Overall Heat Transfer Rates during Regenerative/Film Cooling of Thrust Chambers

Large heat transfer rates are always desired for rocket propulsion applications as high heat loads are associated at the nozzle exit. Different strategies have been employed in order to have high heat transfer coefficients including use of liquid nitrogen, spray cooling etc. ISRO has planned to use aluminium based nano-particles with kerosene in order to cool launching vehicles including GSLV Mk III as it is the heaviest rocket that can carry large payloads. Recently, ISRO has announced to install its own International Space Station (ISS) in future and in such applications larger payloads are to be carried by the rocket. In this work, an analytical study on the thermodynamic properties of the aluminium nano-particles based kerosene nanofluid has been done and an attempt has also been made to develop a temperature and pressure dependent correlation that can be used in computational analysis of thrust chambers while film/regenerative cooling.

Density of Kerosene Aluminium Nanofluid used for Regenerative Cooling Applications of Thrust Chambers

Efficient cooling of thrust chamber is always remaining the key area of research due to the increase in the interest of space exploration programmes. Agencies like NASA, ISRO, SpeceX, BlueOrigin are planning to start commercial flights into the space for common people. To achieve this accuracy and safety both are required and efficient cooling of the thrust chambers is one of them as it has to handle large heat loads twice, first at the time of take-off and secondly, when the vehicle has to land on the ground. ISRO has planned to use kerosene with aluminium nanoparticles for the cooling purposes of the thrust chambers. To achieve higher accuracies, the property variations as a function of temperature and pressure are required so that point to point variation can be visualized comparatively easily. In the present work, density of kerosene and its nanofluid has been studied and in order to use this property variation directly into the simulation software, curve fitting has been done for density of the nanofluid as a function of temperature.

An Algebraic Model for Structural and Life Analysis of Regeneratively-Cooled Thrust Chambers

An algebraic model to evaluate the structural behavior and the cyclic life of regeneratively cooled thrust chambers is presented. The model is based on the evaluation of the stresses, displacements, and elastoplastic strains induced by the severe thermomechanical environment under which the regeneratively cooled thrust chambers operate. The evaluation of the strain range accumulated during an engine duty cycle is then used to estimate the cyclic life by considering the so-called doghouse effect: that is, the peculiar deformation of the chamber inner wall that accumulates cycle after cycle. The cyclic life model is taken and rearranged from an already existing model. Both low-cycle fatigue and tensile ductile ruptures are considered as possible failure mechanisms. A minimum number of inputs are required for the application of the model: the geometry of the investigated section; the hot-gas and coolant pressures; and the wall temperature in the inner and outer walls. Of course, material characterization (mainly in terms of the stress–strain behavior and low cycle fatigue) must also be provided. After the model description and setup, a wide database of experimentally tested thrust chambers (which is almost unique in the technical and scientific literature) is reproduced for the sake of model verification.

Effects of angled film-cooling on cooling performance in a GO2/GH2 subscale thrust chamber

A numerical methodology is employed to simulate the flow and coupled heat transfer in the regeneratively cooled thrust chamber with angled film-cooling. Film-cooling performance is predicted for the conditions with different tangentially angled orifices and azimuthal angled orifices at various mass flow rates of the film coolant. Meanwhile, the experiment is performed in a heat-sink thrust chamber with tangentially angled film-cooling to validate the numerical methodology. The results show that the effects of positive tangential angle on the structure of the diffusion flames and hot-gas temperatures are limited. However, the structures of the flame are obviously affected by the azimuthal angle. In addition, the positive tangential angle could obviously lower the average temperatures in the recirculation region and improve the film-cooling performance in the near-injector region. Besides, the azimuthal angle could reduce the non-uniformity levels of hot-gas-side wall temperatures in the circumferential direction. The temperatures of the face plate increase with the tangential angle but decrease with the azimuthal angle. The application of angled film-cooling could increase the temperature rise of regenerative coolant. The results would be useful for the analysis and optimization design of the film-cooling orifices in the thrust chamber.

Numerical Analysis of Combustion and Film Cooling in LOx-Methane Rocket Engine

Nowadays methane is considered as an alternative fuel for rocket engine application. It is abundant in the outer solar system that can be harvested from mars, Titan, Jupiter, many other planets and moon. The properties of cryogenic methane like higher density, higher vaporization temperature, less challenge in storage requirement compared to cryogenic hydrogen and its higher specific impulse, superior cooling capability and higher coking limit compared to kerosene created a renewed interest on space scientist in choosing methane as a propellant. Combustion and film cooling involving supercritical and transcritical regimes in a high pressure rocket engine combustion chamber is a challenging problem for the rocket engine designers. At supercritical and transcritical conditions, the thermophysical and transport properties exhibit tremendous variations which makes modeling combustion and cooling a demanding task. The present study is based on the oxygen methane thrust chamber model B at DLR P6.1 test bench. The finite volume based software ANSYS Fluent 15.0 is used for the numerical studies. The combustion is simulated using single step eddy dissipation approach. The real gas effects are incorporated using Soave Redlich Kwong cubic equation of state. The numerical scheme is validated by comparing the results obtained with the experimental results reported in literature. Investigations have been carried out to study the effect of film cooling on chamber wall temperature and wall heat flux for different oxidizer inlet temperature. The result revel that the hot gas temperature near the wall is higher in the transcritical combustion compared to supercritical combustion.