Rocket Engine Applications Research Articles

Turbopump Parametric Modelling and Reliability Assessment for Reusable Rocket Engine Applications

The development of modern reusable launchers, such as the Themis project with its LOX/LCH4 Prometheus engine, CALLISTO—a reusable VTVL-launcher first-stage demonstrator with a LOX/LH2 RSR2 engine, and SpaceX’s Falcon 9 with its Merlin 1D engine, underscores the need for advanced control algorithms to ensure reliable engine operation. The multi-restart capability of these engines imposes additional requirements for throttling, necessitating an extended controller-validity domain to safely achieve low thrust levels across various operating regimes. This capability also increases the risk of component failure, especially as engine parameters evolve with mission profiles. To address this, our study evaluates the dynamic reliability of reusable rocket engines (RREs) and their subcomponents under different failure modes using multi-physics system-level modelling and simulation, with a particular focus on turbopump components. Transient condition modelling and performance analysis, conducted using EcosimPro-ESPSS software (version 6.4.34), revealed that turbopump components maintain high reliability under nominal conditions, with turbine blades demonstrating significant fatigue life even under varying thermal and mechanical loads. Additionally, the proposed predictive model estimates the remaining useful life of critical components, offering valuable insights for improving the longevity and reliability of turbopumps in reusable rocket engines. This study employs deterministic, thermally dependent structural simulations, with key control objectives including end-state tracking of combustion chamber pressure and mixture ratios and the verification of operational constraints, exemplified by the LUMEN demonstrator engine and the LE-5B-2 engine class.

A Review of Recent Developments in Hybrid Rocket Propulsion and Its Applications

This paper extensively reviews hybrid rocket propulsion-related activities from combustion engine designs to launch tests. Starting with a brief review of rocket propulsion development history, a comparison among the three bi-propellant rocket propulsion approaches, and hybrid rocket engine design guidelines, a very thorough review related to hybrid rocket propulsion and its applications is presented in this paper. In addition to propellant choice, engine design also affects the hybrid rocket performance and, therefore, a variety of engine designs, considering, e.g., fuel geometry, swirl injection, ignition designs, and some innovative flow-channel designs are also explored. Furthermore, many fundamental studies on increasing hybrid rocket engine performances, such as regression rate enhancement, mixing enhancement, and combustion optimization, are also reviewed. Many problems that will be encountered for practical applications are also reviewed and discussed, including the O/F ratio shift, low-frequency instability, and scale-up methods. For hybrid rocket engine applications in the future, advanced capabilities and lightweight design of the hybrid rocket engine, such as throttling capability, thrust vectoring control concept, insulation materials, 3D-printing manufacturing technologies, and flight demonstrations, are also included. Finally, some active hybrid rocket research teams and their plans for flight activities are briefly introduced.

Evaluation of Boron Combustion for Ducted Rocket Applications Using Condensed Product Analysis

Boron, a metalloid, produces high energy upon combustion. It is recommended as an ingredient for fuel or propellant in rocket propulsion, despite the challenge of extracting its full thermal energy. So far no one has claimed the complete energy conversion of boron upon combustion. On the other hand, the current propulsion system of the Meteor missile uses boron-loaded propellant. The boron-loaded propellant provides an approximately three-fold increase in specific impulse compared to conventional propellants. The present study focuses on boron-HTPB-based solid fuels impregnated with early ignited particles as additives, aiming to assess the combustion performance of boron particles. These additives are magnesium (Mg), titanium (Ti), and activated charcoal (C), and their effects are evaluated based on the residual active boron content in the condensed combustion products (burned residues). An economical tool commonly called stagnation flow or opposed flow burner (OFB) is used to deflagrate the fuel sample by means of pressurized oxygen gas. The condensed combustion products are examined using a field emission scanning electron microscope (FESEM), X-ray diffraction (XRD), thermogravimetric analysis (TGA), and differential thermal analysis (DTA). Among the fuel combinations investigated here, magnesium has been found to be a good burning enhancer of boron, leaving the lowest active boron content (30%) in the burned-out residue. The current research aims to develop an efficient boron-containing solid fuel for hybrid propellant ducted rocket engine applications.

Superior Al90Mg10-Based Composite Fuel Grain for a Hybrid Rocket Engine

A paraffin-based fuel coupled with a nested helical matrix structure is a low-cost, high-performing alternative for hybrid rocket engine applications. The mechanical and combustion properties of the composite fuel grain can be enhanced by replacing the conventional polymer matrix with a metal skeleton. Three-dimensional printing was used to design an Al90Mg10 skeleton embedded with the paraffin-based fuel. The combustion characteristics of the composite fuel grain, including the ignition behavior, pressure oscillations, regression rate, and combustion efficiency, were comprehensively investigated. The properties of grains with and without a secondary perforated structure were compared. The flame structure and metal burning behavior of the hybrid rocket engine were monitored by endoscopic radioluminescence imaging, and the emission spectral characteristics of the plume were analyzed simultaneously. Good flammability makes the Al90Mg10 helical skeleton a promising candidate for enhancing the combustion performance of a paraffin-based fuel grain. The combustion process with rapid ignition was relatively stable, and no additional pressure oscillation frequency was observed. The metal-based composite fuel grains had a superior regression rate to that of a paraffin-based fuel grain (up to 100% higher using the perforated skeleton). Introducing a secondary structure into the fuel grain promoted the reaction and thereby enhanced the combustion performance.

Effects of deposition sequence on microstructural evolution in additively manufactured Cu-Cr-Nb alloy / superalloy bimetallic structures

Although Cu-Cr-Nb alloys / superalloys bimetallic components for liquid rocket engine applications have been fabricated using additive manufacturing, a detailed understanding of their process-structure relation is currently lacking. To bridge this gap, GRCop-42 / Alloy 718 bimetallic structures were fabricated with two different deposition sequences, using powder feedstock with a laser-based directed energy deposition process. The deposition sequence resulted in significant variations in precipitate morphology, composition, and crystal structure within the GRCop-42, particularly at locations near the interface. When Alloy 718 was deposited first, dilution and convective mixing resulted in elevated Ni and Fe levels within the GRCop-42 deposit, which contributed to the formation of C14 (Cr, Ni, Fe)2Nb Laves and α-Cr phases. The anticipated C15 Cr2Nb precipitates were observed within the GRCop-42 material when the deposition sequence was reversed.

Reduced kinetic mechanism for methane/oxygen rocket engine applications: a reliable and numerically efficient methodology

CFD simulations of turbulent reacting flows based on finite rate chemistry often employ reduced kinetic mechanisms to decrease the computational cost, especially if the combustion of hydrocarbons is involved. This work presents a chemical-kinetic methodology, consisting of the formulation, development, testing and validation of a reduced, skeletal mechanism targeted to the Liquid Rocket Engines (LRE) combustion of CH4/O2. The reduced mechanism is generated for combustion processes involving medium-high pressures and ignition of undiluted methane-oxygen, using the 0D/1D open-source software Cantera. The presented mechanism, named Medium Pressure Rocket Burn (MPRB), is achieved from a semi-detailed kinetic scheme, i.e. Lu30, derived from the detailed mechanism GRI-Mech 3.0. Identification of the main chemical reaction paths and sensitivity analysis applied in a sequence leading to a final scheme consisting of 19 species and 51 reactions. Promising results are obtained in terms of ignition delay times and comparison with experimental measurements in high-pressure shock tube tests. The validation is extended to the turbulent case using a sub-scale single-injector combustion chamber with a gaseous injection of CH4/O2 as a benchmark. First, Improved Delayed Detached Eddy Simulations (IDDES) based on a non-adiabatic flamelet database are in good agreement with the available experimental data, although the average thermal load foreseen by MPRB is about 12.6% higher than the case with Lu30 used as reference. Secondly, RANS simulations based on the Eddy Dissipation Concept (EDC) show that accurate results can be obtained with an affordable computational cost, compared to the previously investigated detailed chemistry calculations. Overall the successful validation of the presented reduced mechanism encourages its use for CH4/O2 combustion regimes within this range of applicability.

Evaluation of Large-Eddy Simulation Coupled with an Homogeneous Equilibrium Model for the Prediction of Coaxial Cryogenic Flames under Subcritical Conditions

Large Eddy Simulations of liquid O2/gaseous H2 coaxial flames at subcritical pressure conditions are reported in this paper. These simulations reproduce the experimental Mascotte cases A1, A10 and A30, operating at 1, 10 and 30 bar, respectively, and for which temperature measurements and experimental visualisations are available. The main objective of this work is to assess the accuracy of the multi-fluid Homogeneous Equilibrium Model (HEM) described in Pelletier et al. (Computers & Fluids, 2020) for rocket engine applications. Of particular interest is the comparison with the experimental temperature measurements from Grisch et al. (Aerospace science and technology, 2003). To that purpose, numerical simulations are conducted with care, in order to ensure a proper statistical convergence and estimate the influence of the grid resolution for each case. Despite the crude assumptions—no surface tension and no atomisation model, for instance—that are made with the HEM used in this work, results are found to be in reasonable agreements with the measurements for case A10, even with the coarser grid. For case A30, a fine mesh resolution is required to capture the low intensity recirculation zone downstream of the inner jet necessary to reproduce the shape of the experimental profile. Finally, case A1 simulations, with the lowest Weber number, show large departures with the experimental measurements. This is expected to be due to a deficiency of the model to properly reproduce the two-phase dispersed flow.

Control of Lean Blowout in a Swirl-Stabilized Dump Combustor at Different Levels of Premixing Based on Flame Colour

ABSTRACT Presently, the use of lean combustion has been growing in industrial furnaces, aero-engine, and rocket engine applications due to their low N O X emission and minimum maintenance requirements. On the other hand, due to the excessively low combustion reaction rate, the operation at a low fuel–air ratio leads to the sudden shutdown of the engines, which is termed lean blowout (LBO). Therefore, to operate land-based and aircraft engines at an extremely lean fuel–air mixture, control of LBO has become indispensable. In the present work, we develop several feedback in-loop control strategies using a pixel-averaged intensity ratio between the red and blue components of the flame emission. During a lean operation, based on the quantification of the proximity of the combustion dynamics to the LBO limits using this color ratio, the control strategies make the solenoid valves active to inject a secondary pilot fuel and energize the flame base. We notice that a proper choice of the color ratio as a threshold enhances the stability of both premixed and partially premixed combustion using a low pilot fuel. Besides, we test N O X emission for unpiloted and piloted operations and find that the emission is only marginally affected during a pilot injection.

Feasibility of a Morphing Rocket Nozzle for Thrust Vector Control based on Corrugated Composite Laminates

The paper presents technological, experimental and numerical studies aimed at the development of a morphing nozzle in composite material with a deformable divergent section, which is designed to control the vector of thrust in rockets without using mechanisms or localized flexible joints. The nozzle design is characterized by corrugated composite walls and longitudinal elements that make possible a separation of load paths involved in bending response, bearing of internal pressures and related force resultants. Mechanical requirements and conditions of pressure, temperature and gas velocity in the morphing part of the nozzle are defined by selecting and investigating two potential application scenarios. Thereafter, both technological aspects and the assessment of mechanical performances are addressed by designing, manufacturing and testing a composite demonstrator with morphing capabilities. A lightweight and flexible thermal protection system is proposed, designed and numerically analyzed. Finally, a virtual demonstrator of the composite morphing nozzle is developed. Results point out the possibility of manufacturing composite elements with axial load bearing capability that may undergo significant bending deformation without failures. The integration of a flexible thermal protection systems leads to a promising concept for the development of innovative morphing nozzle for rocket engine applications.

Laminar flame speed evaluation for CH4/O2 mixtures at high pressure and temperature for rocket engine applications

Pure methane-oxygen mixtures in liquid rocket engines lead to extreme pressure and temperature conditions that are prohibitive for most of the experimental setups. Hence, there is very little data on such flames in the literature, especially concerning the laminar flame speed Su, often limited at atmospheric pressure. The recent development of methalox rocket engines, which design process often requires CFD calculations, brings this lack of data to the forefront. Indeed, the CFD simulations require valid chemical schemes in the real operating conditions. To address this problem, flame measurements have been performed in a special isochoric combustion chamber with full optical access (OPTIPRIME) developed at ICARE. An extensive database in conditions never tested before is generated for several equivalence ratios, temperature and pressure ranges. Multiple chemical mechanisms were then compared to those results, showing various levels of agreement. Hence, the best mechanism from the literature on OPTIPRIME results and other literature experimental data was selected. A sensitivity analysis was performed to identify key chemical reactions controlling the flame speed. These key reactions could later be tuned by an optimization process to perfectly match the experimental results. Finally, additional measurements were performed in order to develop a Su=f(P,T) correlation to build a future flame speed database under rocket engines relevant conditions.

Flow rate improvements in additively manufactured flow channels suitable for rocket engine application

This contribution describes the investigation of flow channels which are designed to be directly integrated into an aerospike engine by means of additive manufacturing with laser powder bed fusion (LPBF). During the experimental testing of a previous aerospike engine in 2019, it was observed that high surface roughness of such additively manufactured integrated channels caused a significant reduction in the mass flow rates of the propellants ethanol and liquid oxygen as well as the coolant due to increased pressure drop. In an extensive study within the CFDmikroSAT project, various factors influencing this surface roughness are, therefore, being investigated, which include the geometry of the channels as well as selected manufacturing parameters of the LPBF process, such as layer thickness and component orientation. To further reduce the roughness after manufacturing, suitable post-processing methods are also being investigated for internal cavities, initially analysing the abrasive flow machining process. Within the paper, the overall investigation approach is presented, such as the overview of the considered specimens, and the initial results of a various studies with selected specimens are discussed. These studies consist of the examination of surface roughness reduction, shape accuracy and flow behaviour of post-processed cooling channel specimens. Finally, a brief overview of the already manufactured aerospike demonstrator is presented.

On the understanding of a cryogenic two-phase LOX/GH2 flame: Parametric sensitivity, characteristic scaling and phase instability

This study concerns the cryogenic hydrogen combustion in subcritical pressure conditions, with practical relevance to rocket engine applications. A cryogenic two-phase flame in the counterflow configuration is calculated and analyzed with the consideration of real-fluid effects and heat/mass transfer across the liquid-gas interface. The effects of pressure, strain rate, fuel inlet temperature, and heat loss on the flame characteristic behaviors are examined. It is found that the vaporization rate of liquid oxidizer scales with the square-root of pressure (p) times strain rate (ast), and is not a limiting factor for combustion. The total heat release scales with past at lower strain rates while with p4/5ast1/3 at higher strain rates. The significant results on the phase-stability of cryogenic flame are also established. It is found that the unstable phase, in terms of vapor-liquid equilibrium, arises in the vicinity of the liquid-gas interface; however, all unstable-phase states still stay in the metastable region of the phase diagram. For all the flame solutions considered in this study, no thermochemical state enters the spinodal region.

Dataset of Wall-Resolved Large-Eddy Simulations Turbulent Pseudoboiling in Cryogenic Hydrogen Pipe Flows

In this paper, a dataset of wall-resolved large-eddy simulations of cryogenic hydrogen at supercritical pressure and different values of wall heat flux is presented. The aim is to provide a reference dataset for wall-function development under trans- and supercritical conditions, such as those found in liquid rocket engine applications. The employed numerical framework is a pressure-based segregated low-Mach-number approach based on an equation-of-state independent formulation. The wall-adapting local eddy-viscosity subgrid model is used for turbulence closure. Real-gas effects are taken from the National Institute for Standards and Technology database and stored as a function of a nondimensional temperature at the considered pressure. A validation and a grid-convergence analysis are first performed on an incompressible case without imposed heat flux. The effect of axial, radial, and azimuthal refinements on first- and second-order velocity statistics is discussed and compared with direct numerical simulation data from the literature. A parametric analysis at different wall heat fluxes is then performed by keeping the inlet mass flux, temperature, and Reynolds number constant. Particular attention is devoted to turbulent pseudoboiling and its effect on the wall temperature. The latter shows a more pronounced increment as the heat flux increases, which is attributed to the pseudochange of the phase in the core flow. Correspondingly, a flattening of the probability density function of the temperature is observed, and it is associated with the pseudoboiling interface forming close to the wall and causing a more intense stratification. First- and second-order statistics for velocity and selected scalars are then presented, and the effect of pseudoboiling is discussed. The effect of the wall heat flux on the viscous and thermal resolution of the computational grid is also assessed, and considerations on the relation between turbulent pseudoboiling and near-wall gradients is finally provided.

Thermomechanical response of out-of-autoclave infused carbon-phenolic laminates for rocket engine applications subjected to surface ablation

Cost-affordable and reliable manufacturing technologies to produce Carbon/Phenolic Composites (CPCs) are essential for competitive access to space. These materials are used as ablative Thermal Protection System (TPS) for space applications, such as rocket nozzles or the heat shields of space vehicles. Out-Of-Autoclave (OOA) infusion processes can reduce the manufacturing cost of composites: in this work, OOA produced CPC laminates were compared to Carbon/Epoxy Composites (CECs). The CPCs produced in this work were mechanically characterized to assess their quality and compared with state-of-the-art counterparts through the data available in the open literature. CPC laminates instrumented with ultra-small thermocouples were subjected to Oxy-Acetylene Torch (OAT) ablation testing and a numerical approach was assessed to predict the temperature reached in the laminate layers. Post-burning mechanical tests performed on specimens subjected to the ablation test showed that when the temperature does not exceed a given threshold, the material retains adequate properties to perform a structural role.

Development of solid propellant for the production of fire suppression aerosols

This paper documents the development of solid propellant aerosol generators (SPAGs) as alternative fire suppression systems. SPAGs mainly consist of an igniter, a solid propellant (SP), a cooling system and an outlet. The main focus of this study is on the SP as the research item. Such SP consists of oxidant, fuel and additives. This research work aimed to investigate the application of the aerosol produced by a SP as a potential fire suppression agent. The first step in this regard involved a numerical calculation, using RPA® software which is designed for and used in rocket engine applications, to define a number of combustion and thermodynamic parameters related to the SP. This was followed by a parametric study of different SP compositions and their respective fire suppression aerosols (FSAs). The SPs that produced the highest fraction of FSA were chosen for an experimental comparison. A special experimental setup was developed for this comparison and data were generated by performing more than 100 tests. The experimental data showed that the selected SP are able to suppress small n-heptane pool fires in compartments effectively, validating the results of the RPA® calculations. The experimental data proves that the suppression effect is inhibition.

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.

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.

Propellant Atomization for Porous Injectors

Porous injectors represent an alternative injection concept to coaxial injectors for rocket engine applications using gas/liquid propellant combinations such as liquid oxygen (LOX)/hydrogen (). This paper summarizes the main design features of porous injectors and proposes a mechanism of atomization for porous injectors that is considerably different to the atomization mechanism for coaxial injectors. The results of hot-fire test campaigns are presented, in which several parameters relevant to the injection process were varied: the injection velocities and momentum fluxes, the combustion chamber Mach number at the beginning of the nozzle contraction, and the LOX injector diameter. All hot-fire tests were conducted at P8 test facility for high pressure combustion research at the DLR site of Lampoldshausen with 50-mm-diameter combustion chambers operated with at sub- and supercritical chamber pressures between 30 and 100 bar. The results presented here are supporting the proposed mechanism of atomization and allow derivation of some general design guidelines for porous injectors.

Development of a high flow rate piston pump for rocket engines

There is a need for cheaper small satellite rocket launchers. One way of reducing costs is to reduce the cost of individual components. We describe here the development of an inexpensive piston pump for rocket engine applications. It demonstrates that it can deal with the combined requirements of high flow rate and operating speed at cryogenic temperatures without incurring the costs normally associated with such performance.

Spray formation and atomization characteristics of non-Newtonian impinging jets at high Carreau numbers

Over the last two decades there has been a renewed interest in employing shear-thinning non-Newtonian liquids for use with liquid rocket engines due to a desire for greater safety and performance. Since the difficulty with producing quality atomization is the main disadvantage for using non-Newtonian liquid propellants, further understanding of the breakup physics and accurate measurement of the spray characteristics is needed. The atomization produced by the like-on-like jet impingement of three different non-Newtonian liquids was experimentally investigated in this work using shadowgraphy and Phase Doppler Anemometry (PDA) at high strain rates (105–106), which correspond closely to the strain rates used in practical rocket engines. The three water-based solutions tested were of two different grades of Carboxymethylcellulose (CMC): 0.5 wt.-% CMC-7HF, 0.8 wt.-% CMC-7MF, and 1.4 wt.-% CMC-7MF. The Bird–Carreau rheological model was used to characterize the non-Newtonian behavior and a generalized Carreau number was used as the primary parameter to characterize the liquid viscosity inside the orifice at injection. A generalized Bird–Carreau jet Reynolds number Rej,gen-BC and the jet Weber number Wej were used as primary parameters to characterize the spray formation, sheet breakup length, drop size, and drop velocity. The unique spray formation a behavior of the three non-Newtonian liquids was ascribed to the relative influence of the inertial vs. viscous forces, primarily expressed through Rej,gen-BC. At many relatively high inertial force test conditions, drops in the spray were near spherical and could be reliably measured using PDA. Spray characteristics of intact sheet length, drop diameter, and drop velocity were observed to be strong functions of Rej,gen-BC. In addition to providing fundamental understanding to aid with injector design for liquid rocket engine applications using non-Newtonian liquids, the measurements presented in this work could be used to validate a future closed-form analytical non-Newtonian atomization model, which remains an open problem in non-Newtonian impinging jets literature.