Rocket Thrust Research Articles

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.

Autophage Engines: Method to Preset Gravity Load of Solid Rockets

A retardation effect, which is the decrease of the feed pressure and the thrust of an autophage rocket, is considered. The effect is caused by inertial forces acting on the rocket sliding engine and has an exponential decay pattern. It is proposed to apply the influence of the effect on the rocket thrust as a method of throttling and presetting gravity-load (g-load) behavior. The method consists of providing the rocket with an engine of proper mass and proper retardation capability as determined by a retardation coefficient. It was shown that the preset g-load pattern includes a decay, increase, or constancy as particular cases. To calculate the rocket characteristics by taking into account the retardation effect, several kinds of the sets of equations are proposed for numerical, exact analytical, and dimensionless solutions. The nature and domains of the retardation coefficient are investigated. The general case of the g-load behavior is illustrated with plots. The novelty of the subject does not allow omitting the rocket engineering feasibility and validation with experimental results; however, to make the paper concise, these paragraphs are presented in short just to give overall insight into the subjects. The other problems (such as flight control, strength of structure, mass breakdown, and economical efficiency) will be discussed in future publications.

Computer-aided design of low-thust rocket engines using the domain-specific knowledge database and CAE / CAD systems

The paper presents approaches to computer-aided design of low-thrust thrust rocket engines using an extensive knowledge base that allows making basic technical decisions that determine the conceptual design of the engine, based on the developed algorithm of this process. The procedure of creating an electronic 3D-model of a low-thrust rocket engine fueled by gaseous oxygen-hydrogen in the environment of the graphical complex UNIGRAPHICS is described. 3D electronic models of the main elements of a rocket engine with a thrust of P=25N were obtained, with subsequent virtual assembly of all components, including the components comprised in the knowledge base, providing the development, among other things, of design documentation, creation of a production environment based on an electronic engine model, preparation for the product manufacturing and the manufacturing proper.

Real-metric spacetime own-surfaces hosting nongeodesic radar paths crossing ‘hemix’ own-lines and shared velocity helices

Pondering Enrique Loedel’s 1948 symmetrical spacetime chart whose angle’s sine reflects scaled velocity between inertial reference frames, led in 2004 to a most curious discovery: Angles of a unit sines law ratio spherical triangle geometricise relativistic velocity composition. Overlooked in a paper by Einstein’s colleague Sommerfeld and kept under wraps until publication of a recent book, the germane criterion points to a seemingly unknown hemispherical ‘hemix’ spiral which encapsulates the Gudermannian relating a fixed thrust rocket’s velocity and clocked own-time. This opens new avenues for analysing relativistic noninertial length contexts whereby hemix-generated real-metricλ|τ‘own-surfaces’ visualisable in R3 are corroborated by radar path attributes, oddly a strategy unexploited in relativity. The Euclidean models epitomise not only Born’s (one off) ‘rigid motion’ but also non-Minkowski extended medium acceleration scenarios such as the previously unresolved Bell’s spaceships problem where the ±[+++-] signature is misplaced, increment ‘own-lines’ and radar paths turn out to be nongeodesics, and Gauss curvature remains the same along each shared own-time likewise nongeodesic ‘medium curve’.

Performance analysis for aerial-towing system

The dynamics of open chain systems have been widely investigated as a result of its increased applications such as aerial towing system, submerged towed cable applications, robot manipulators, and bio-mechanical applications. In this study, a practical engineering system, Mine Clearing Line Charge MICLIC system is taken as an example for an open chain system with accelerated end and varying number of segments. To validate the proposed model, a comparison between two different finite segment methods namely Kane’s method and Riccati transfer matrix method of linear multibody system (Riccati-MS-TMM) have been performed for a large open chain system. Results show that computational speed of Riccati-MS-TMM outperforms Kane’s method due to its huge matrix size to be solved. Then, using a predefined rocket motor total impulse, a parametric study has been performed on MICLIC system to illustrate the impact of different rocket motor thrust profiles on the resulted flight performance including the missile range, flight time, maximum tension applied on the rocket towing bridle, and the missile stability.

Heat Transfer and Combustion Simulation of Seven-Element / Rocket Combustor

In the present Paper, the simulation of the flow inside an experimental rocket thrust chamber is undertaken. The combustor’s injector consists of seven individual coaxial injector elements, while the chamber and nozzle segments are water cooled. The results presented in this Paper are obtained with three-dimensional Reynolds-averaged Navier–Stokes simulations using an adiabatic flamelet formulation for the chemistry modeling. The main focus is placed on examining the effect of the different turbulence models on the flame structure and on the resulting pressure and wall heat flux. The obtained numerical values are compared to experimental measurements, delivering good agreement in the heat flux profile at the combustion chamber wall and a slight underestimation of the pressure profile of approximately 2.5%. Greater discrepancies are observed in the heat flux of the nozzle segment but are largely attributed to the experimental setup. A conjugate heat transfer simulation of the structure and cooling channel flow confirms this assumption, and results for both one-way and two-way couplings are shown. It is demonstrated that a one-way coupling between hot gas and structure is sufficient due to the low sensitivity of the wall heat flux on the wall temperature. The azimuthal variation of the heat flux is also examined, and interestingly the heat flux showcases a local minimum at the position directly above the injector element. It is shown that an increased concentration of colder fuel-rich gas directly above the injector due to a strong vortex system leads to the local minimum in heat flux values and is strongly influenced by the injector/injector interaction near the face plate.

A discrete, energetic approach to rocket propulsion

Most rockets convert the energy stored in their propellant mass into the mechanical energy required to expel it as exhaust. The ‘rocket equation’, which describes how a rocket’s speed changes with mass, is usually derived by assuming that this fuel is expelled at a constant relative velocity. However, this is a poor assumption for cases where the rocket promptly loses a large fraction of its mass. Instead, I derive the change in speed for a rocket that emits its fuel in N discrete pellets, with a constant mechanical energy produced per unit fuel mass. In this model I find that the rocket’s speed change is greatest when all the fuel is expelled at once (N = 1). In the limit of many small pellets (N → ∞), I show that the velocity change approaches, from above, the prediction of the continuous thrust rocket equation. For this model of rocket propulsion, I quantify how the fuel’s total available energy is divided between the rocket and exhaust. In the limit of continuous thrust, the rocket can utilise no more than 65% of the available mechanical energy as its kinetic energy. In an online supplement, I compare this model of discrete propulsion with those previously published. This topic uses momentum conservation and mechanical energy concepts at the introductory undergraduate level.

Experimental study on dielectric barrier discharge plasma-assisted combustion for nozzle

Aiming at the structure of small thrust rocket engine injector, a coaxial shear-type dielectric barrier discharge plasma nozzle was designed, and the effects of methane discharge and air discharge on combustion effect under different excitation and different mixing ratio conditions are studied. The results show that: the flashback occurs when the methane discharge voltage is high, the air discharge can broaden the blow off limit of the air/methane diffusion flame, and has the effect of stabilizing the flame and reducing the lift height.

Rotation Calibration Method for Thrust Based on Force Error Analysis

Accurate measurement of thrust in rocket engine, as a core parameter of attitude control of rocket engine, is an important guarantee for reliable operation of rocket engines. Therefore, it is very important for accurate calibration. In this paper, a novel rotation calibration method for rocket thrust engine is proposed based on piezoelectric force measurement system. Based on the calibration matrix, sensitivity analysis model of calibration accuracy to coupling relationship between axes is established, and error mechanism and offset mechanism of force are studied. Starting from the deviation principle between actual axis of thrust system and theoretical one, an elliptical geometrical distribution law of lateral force caused by main force is obtained. According to the periodicity of ellipse distribution of calibration force, a new calibration method, rotation calibration method, is obtained based on an idea of homogenization error. Finally, vector force verification is designed, and experimental data are calibrated with the rotation calibration method compared with the traditional linear calibration method, verifying feasibility and rationality of the calibration method. The novel calibration can be also used in force measurement in any rotation devices like manipulator in robot,thrust vector engine, manufacturing force, rocket engine,etc.

Methodology of development and test of the electrical rocket propulsion system for telecommunication spacecraft Yamal-200 (to the 15th anniversary of operation in space)

The article summarizes the results of a 15-year operation of the Yamal-202 telecommunication spacecraft in geostationary orbit. The review of using electrical rocket engines in spacecraft in domestic and foreign projects is made. The questions of methodology for developing and testing the electrical rocket propulsion system for the Yamal-200 spacecraft are considered including features of the RSC Energia stand base, specific features of the equipment and methodology for filling tanks with working fluid such as xenon. Special attention is paid to the preliminary joint test of the power supply and control equipment with thruster modules in order to justify the long-term operation of electrical rocket engines. In the analysis of the 15-year operation of the electrical rocket propulsion system the initial operation phase of the spacecraft is shown with the cruise mode of electrical rocket thruster modules for installation at operating points 49 and 90 EL with the corresponding operating time of all thruster modules. The final data on the operating time of thruster modules KA-201 for 10.5 and KA-202 – for 15 years of operation is presented that gives significant statistics on the use of engines СПД-70. The estimate of the remaining mass of the working fluid and the capability of further operation of the electrical rocket propulsion system is made. Key words: geostationary orbit, electrical rocket engine, thruster module, united propulsion system, predelivery checkout tests, filling with working fluid.

Numerical Investigation on Radiative Heat Loads in Liquid Rocket Thrust Chambers

In liquid rocket engines, wall heat transfer analyses are of primary importance for the thrust chamber design. Numerical investigations usually derive the convective heat flux from computational fluid dynamics simulations, whereas the radiative counterpart is often neglected. The goal of the present paper is a systematic numerical investigation on the role of radiative heat transfer on the overall thermal loads occurring in the thrust chamber of liquid rocket engines. To this purpose, a numerical code for thermal radiation computations is coupled to a computational fluid dynamics solver. The results of the numerical setup are discussed and compared to the available literature’s numerical data. Then, the radiative contribution to the overall heat flux in the thrust chamber of oxygen/hydrogen and oxygen/methane liquid rocket engines is investigated through a wide parametric analysis. The results show that, in typical operating conditions, the radiative heat load is proportionally greater in methane-fueled engines than in hydrogen-fueled engines, and that the radiative contribution to the overall wall heat flux becomes increasingly relevant for large-scale low-chamber-pressure engines. A methodology for a quick preliminary evaluation of the radiative wall heat flux in liquid rocket thrust chambers is finally presented and verified against coupled simulations.

Numerical investigation of flame appearance and heat flux and in a deep-throttling variable thrust rocket engine

Much attention has been drawn to liquid-propellant rocket engine with thrust that can be varied on demand in recent years. A GO2/kerosene deep-throttling variable thrust rocket engine using the newly proposed gas-liquid pintle injector is presented. Three-dimensional numerical simulations are conducted to investigate the flame appearance and heat flux distribution of the engine. The SST k-omega model is used for modeling turbulence and single-step finite-rate kinetics are used for modeling combustion. A grid sensitivity study is performed to examine the validity of the simulation results. The effects of total mass flow-rate, kerosene droplet injection velocity and kerosene droplet size distribution on flame appearance and heat flux distribution of the combustor are studied in detail. Three entirely different flame appearances are observed and it correspondingly causes different heat flux distributions. Details of the flow field structure and mass fraction contour inside the combustor are discussed to explore the cause of heat flux distribution. The work will provide a reference for the design of thermal protection system of the variable-thrust rocket engine operating with a wide range of space.

Design and Analysis of Missile Nozzle

In this paper CFD analysis of pressure and temperature for a rocket nozzle with two inlets at Mach 2.1 is analyzed with the help of fluent software. When the fuel and air enter in the combustion chamber according to the x and y plot, it is burning due to high velocity and temperature and then temperature increases rapidly in combustion chamber and convergent part of the nozzle and after that temperature decreases in the exit part of the nozzle. It is concluded in this paper that two inlet rocket nozzle is having better performance than single inlet. We know that the driving forces such as conversation, pressure and temperature gradient can causes species transport, momentum transport and energy transport respectively. Scientist have been worked on “Comparison of the rocket engines efficiency in the case of low thrust orbit-to-orbit transfers” and their findings are the following: The main task of this paper is to compare two types of low thrust rocket engines: constant thrust vs. variable-thrust engines. They are concerned with efficiency, where efficiency is evaluated in the case of the orbit-to-orbit transfer with maximum payload mass in the central Newtonian gravity field.

Inverse heat transfer method applied to capacitively cooled rocket thrust chambers

Measurements of the heat loads in experimental lab-scale rocket combustors are essential in order to obtain information about the mixing and energy release of the propellants, the injector/injector interaction as well as the injector/wall interaction. Usually the hardware used for single-element rocket thrust chambers is capacitively cooled in order to reduce the complexity of the system. The present work demonstrates an efficient method for estimating the time- and spatially resolved heat flux distribution at the hot gas wall of such engines using the information provided by temperature measurements in the material. The method is implemented in the code Roq̇FITT (Rocket q̇ Flux Inverse Thermal Tool) and is applied for the evaluation of test data of CH4/O2 and H2/O2 experiments. Three separate capacitive combustors are investigated, which are operated at the Chair of Turbomachinery and Flight Propulsion (LTF) of the Technical University of Munich (TUM): a single-element cylindrical, a single-element rectangular and a multi-element rectangular chamber. The use of the 3D inverse method for different load points gives significant information about the effect of the different propellant combinations, the choice of mixture ratio and pressure level, the spanwise heat flux distribution and hence injector/injector interaction as well as the transient effects during the igniter operation.

RX-320 Rocket Static Pressure Combustion Chamber Prediction and Validation by Using Invers Method

The static pressure data of the combustion chamber which can generally be obtained by performing direct measurements when static test is performed on the rocket is an important parameter in predicting the thrust and design of the combustion chamber of the rocket. However, there is a model rocket for flight test that is used in static test. Thus, there is no mounting for static pressure sensors (for measurement) are made. To solve the problem, then the inverse method is used as an iterative solution for the basic equations of the rocket thrust force in the nozzle by guessing the value of the static pressure of the combustion chamber firstly and calculate the iteration by including the value of the rocket thrust from static test data and the efficiency variation of the nozzle. The results of this calculation are then validated by using a 3D-CFD numerical simulation to obtain a more detailed comparison on the nozzle. In this research RX 320 LAPAN rocket nozzle with focus on maximum static thrust data of static test results is used. The 3-D numerical simulation is performed using Numeca CFD software, with k-extended wall extended turbulent model, numerical multigrid level 3 scheme, center based, and convergence criteria of 10 e-05. The result of calculation by inverse method and its comparison with numerical simulation shows that the smallest difference of the combustion chamber static pressure between inverse method and numerical simulation is 0.017%, that is achieved at 92% nozzle efficiency. At this point, the static pressure of the combustion chamber is 57.94 bar. From this point of view, the results of this comparison indicate that the inverse method can be used accurately for static pressure of the combustion chamber prediction, if the nozzle efficiency is given correctly. Furthermore, with given static pressure of the combustion chamber correctly, it will be very helpful in the design of the more optimum combustion chamber.

Effects of heat treatments to inner liner material, thermal barrier coating, and outer shell material on lifetime of a combustion chamber

Key factors were investigated to extend the lifetime of a combustion chamber of a large thrust rocket engine. Two-dimensional finite element simulations were conducted to estimate creep damage and low cycle fatigue damage to the inner liner of a combustion chamber wall. In these simulations, the effects of difference of heat treatments to the inner liner material, heat shield by a thermal barrier coating, and thermal expansion of the outer shell on lifetime were examined. It was found that heat shield by a thermal barrier coating had the largest effect in extension of the lifetime though the effect depended on the physical characteristic of the inner liner material. Expansion or contraction of the outer liner during combustion affects the restriction of the inner liner resulting in the extension of the lifetime. By combination of these effects, the lifetime can be extended up to about a factor of 10 longer than that of a combustion chamber without any effects.

Solid Rocket Motor Propellant Optimization with Coupled Internal Ballistic–Structural Interaction Approach

This research aims to optimize the geometric design of slotted propellant grains for solid rocket motors with respect to coupled internal ballistic performance and structural strength criteria. In-house codes such as a zero-dimensional internal ballistic solver and an analytical burnback solver are implemented to compute the variation of chamber pressure and the rocket thrust transiently. Structural analysis of the solid propellant is achieved by using a parametric linear viscoelastic model and a parametric cooldown heat transfer model, both of which are based on the finite element method. The transient temperature distribution data derived from the cooldown process are required inputs for the material properties to be used in the viscoelastic structural analysis. To enable an efficient optimization process, a surrogate heat transfer model that predicts the cooldown time of the system by eliminating expensive iterations is also implemented and validated. Within a coupled analysis approach, the pressure data obtained from the internal ballistic performance analysis are used for the ignition step of the linear viscoelastic analysis. The structural analysis results are evaluated by using a deterministic approach based on the margin of safety with respect to the stress and strain criteria. Finally, the optimum geometrical parameters for a slotted grain subjected to both structural and internal ballistic performance constraints are investigated through multidisciplinary optimization techniques.

Ignition characteristics and combustion performances of a LO2/GCH4 small thrust rocket engine

A 500 N model engine filled with LO2/GCH4 was designed and manufactured. A series of ignition attempts were performed in it by both head spark plug and body spark plug. Results show that the engine can be ignited but the combustion cannot be sustained when head spark plug applied as the plug tip was set in the gaseous low-velocity zone with thin spray. This is mainly because flame from this zone cannot supply enough ignition energy for the whole chamber. However, reliable ignition and stable combustion can be achieved by body spark plug. As the O/F ratio increases from 2.61 to 3.49, chamber pressure increases from 0.474 to 0.925 MPa and combustion efficiency increases from 57.8% to 95.1%. This is determined by the injector configuration, which cannot produce the sufficiently breakup of the liquid oxygen on the low flow rate case.

Technology for Transient Simulation of Vibration during Combustion Process in Rocket Thruster

The article describes the technology for simulation of transient combustion processes in the rocket thruster for determination of vibration frequency occurs during combustion. The engine operates on gaseous propellant: oxygen and hydrogen. Combustion simulation was performed using the ANSYS CFX software. Three reaction mechanisms for the stationary mode were considered and described in detail. The way for obtaining quick CFD-results with intermediate combustion components using an EDM model was found. The way to generate the Flamelet library with CFX-RIF was described. A technique for modeling transient combustion processes in the rocket thruster was proposed based on the Flamelet library. A cyclic irregularity of the temperature field like vortex core precession was detected in the chamber. Frequency of flame precession was obtained with the proposed simulation technique.

Some Calculated Research Results of the Working Process Parameters of the Low Thrust Rocket Engine Operating on Gaseous Oxygen-Hydrogen Fuel

The paper presents the calculating results of the combustion products parameters in the tract of the low thrust rocket engine with thrust P ∼ 100 N. The article contains the following data: streamlines, distribution of total temperature parameter in the longitudinal section of the engine chamber, static temperature distribution in the cross section of the engine chamber, velocity distribution of the combustion products in the outlet section of the engine nozzle, static temperature near the inner wall of the engine. The presented parameters allow to estimate the efficiency of the mixture formation processes, flow of combustion products in the engine chamber and to estimate the thermal state of the structure.