This paper presents simulation activities performed in the frame of a partnership program between the German Aerospace Center (DLR) and the launcher directorate of the French National Center for Space Studies (CNES). The specific test case presented in this paper is a supercritical LOx-methane combustion test on the penta-injector combustion chamber BhpHrM (high pressure and high mixture ratio) of ONERA's test bench Mascotte. The current paper presents numerical works on the hot gas-side analysis of a LOx-methane combustion chamber. Two different simulation approaches are compared in this paper. Both simulations are Reynolds-averaged Navier-Stokes (RANS) simulations. The DLR simulation is computed with the DLR computational fluid dynamics (CFD) code TAU. It uses a flamelet model to solve the combustion coupled to the Soave-Redlich-Kwong (SRK) equation of state. It also uses a low Reynolds approach to model the heat transfer. The CNES simulations have been performed with CPS_C by CT Ing & eacute;nierie. CPS_C uses a finite rate chemistry model with a 9-species and 21-reaction kinetic scheme coupled to the SRK equation of state. The paper investigates the temperature and species distribution within the flow field of the combustion chamber. It shows the difference in oxygen cores length, flame shapes, and recirculation zones within the major axis of the flow field and several axial cuts. The upsides and downsides of the specific modeling approaches are presented and compared. Finally, the experimental measurements of the wall temperature are compared to the wall temperature obtained from the simulation. The complex three-dimensional shape of the wall heat flux and temperatures are not captured in their entirety, especially in the front half of the combustion chamber. In contrast, in the latter half of the combustion chamber the values for temperature and wall heat flux correspond quite well between simulation and experiment.

A shear-coaxial-element is a prominent choice in terms of injection technology for bipropellants with a high density ratio. While this type of injector performed well for liquid-oxygen and hydrogen combustion in an optically accessible single-element combustor and a multi-injector thrust chamber, the same combustion devices featured instabilities using liquid-oxygen/natural-gas. Both experiments showed short-lived events of oscillatory combustion and high-amplitude and high-frequency limit-cycle combustion instabilities within the analyzed tests. The stochastic distributed occurrence of such intermittent events of heightened excitation are typically related to operation of the combustion system close to its stability boundary. Analyzing the high-speed imaging of the optical accessible experiment during these phases showed injector generated hydrodynamic phenomena preceding the short-lived combustion instabilities. Finally, depending on the chamber configuration and operating conditions - primarily the momentum flux ratio - these events triggered high-frequency combustion instabilities. Consolidation of the data obtained from different hardware configurations highlighted the significant role of recessed injector elements in developing combustion instabilities.

This study introduces a tailored postprocessing model for in-depth assessment of critical failure mechanisms within the inner liner of regeneratively cooled combustion chambers in reusable liquid rocket engines (RLREs). The approach integrates ductile and brittle damage using a damage accumulation framework based on methodologies from Bonora et al. ("Identification of the Parameters of a Non-Linear Continuum Damage Mechanics Model for Ductile Failure in Metals," Journal of Strain Analysis for Engineering Design, Vol. 39, No. 6, 2004, pp. 639-651) and Dufailly and Lemaitre ("Modeling Very Low Cycle Fatigue," International Journal of Damage Mechanics, Vol. 4, No. 2, 1995, pp. 153-170). By incorporating high-temperature material properties of the copper-chromium-zirconium alloy, the model undergoes validation against experimental data from thermomechanical fatigue tests and detailed thermal-structural finite element analysis. The results demonstrate the model's robustness in predicting damage progression, pinpointing crack initiation zones, and quantifying ratcheting effects, thereby providing a comprehensive tool for assessing fatigue life across diverse RLRE architectures and enhancing engine reusability.

The scope of this work comprises the experimental investigation of film cooling with gaseous hydrogen (GH2) in the vicinity of the faceplate and the cylindrical region of a subscale GH2/LOx-rocket combustion chamber. The experimental setup consists of a novel experimental combustion chamber with a rectangular cross section and interchangeable measurement devices, allowing for optical access and simultaneous measurement of wall heat fluxes and combustion chamber pressures. A multi-injector faceplate prevents window film cooling from interacting with the investigated near-wall cooling film and injector reacting flow. With the use of a novel patented calorimetric measurement device, 2D wall heat fluxes were evaluated. Based on these data, a new correlation for the film cooling efficiency is elaborated and verified for operating points in the supercritical, transcritical, and subcritical regimes of oxygen. Furthermore, the influence of LF instabilities on wall heat input is experimentally investigated and presented. All in all, this work provides a wide and accurate database for validation of numerical tools and gives insight into the processes of film cooling in a rocket combustion chamber.

A shear-coaxial element is a prominent choice in terms of injection technology for bipropellants with a high density ratio. While this type of injector performed well for liquid-oxygen and hydrogen combustion in an optically accessible single-element combustor and a multi-injector thrust chamber, the same combustion devices featured instabilities using liquid oxygen/natural gas. Both experiments showed short-lived events of oscillatory combustion and high-amplitude and high-frequency limit-cycle combustion instabilities within the analyzed tests. The stochastic distributed occurrence of such intermittent events of heightened excitation is typically related to the operation of the combustion system close to its stability boundary. Analyzing the high-speed imaging of the optically accessible experiment during these phases showed injector-generated hydrodynamic phenomena preceding the short-lived combustion instabilities. Finally, depending on the chamber configuration and operating conditions-primarily the momentum flux ratio-these events triggered high-frequency combustion instabilities. Consolidation of the data obtained from different hardware configurations highlighted the significant role of recessed injector elements in developing combustion instabilities.

Forced motion simulations of an overexpanded subscale rocket nozzle were performed to investigate the transient mechanisms that lead to self-exciting fluid–structure interaction as observed in preceded studies. The pressure response to the deformation could be separated into two regions upstream and downstream the flow separation position. Within these regions the transient part of the pressure was analyzed using fast Fourier transform based on the method of generalized aerodynamic forces. The amplitude spectrum and phase shift distribution of the pressure response could be explained by superposition of three independently acting mechanisms: the inclination effect, the existence of a moving axial pressure wave, and intrinsic oscillations caused by the turbulence created by the strong shock system. Simplified simulation setups using a bent flat plate and a detailed unsteady simulation of the flow in the undeformed nozzle were analyzed to validate these assumptions.

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.

The DLR Liquid Upper Stage Demonstrator Engine (LUMEN) is an experimental expanderbleed cycle rocket engine designed for LOX/LNG propulsion in the 25 kN thrust range. Its primary purpose is to serve as a flexible testing platform for industrial and institutional partners, focusing on addressing the challenges associated with rocket engine reusability like health monitoring and intelligent control algorithms. The LUMEN demonstrator allows for sub-component exchange and testing. This approach simplifies for example the turbopump design by incorporating two turbopumps instead of a single shaft configuration, providing greater flexibility for technology development. In addition to component technology maturation, the modular system approach enables the testing of technologies that would otherwise be challenging to assess in full-scale or simulation environments, such as health management and its feedback to engine control. The closed-loop neural network-based control improves the accuracy of approximating the nonlinear behavior of the engine. Finally, the upcoming development steps for LUMEN are outlined.

Fluid hammer occurs when a flow in a pipeline is rapidly stopped due to valve closure or shutting down a pump. The resulting pressure wave can reach very high amplitudes, depending on the fluid properties and the flow velocity. In spacecrafts and launchers, where the reduction of weight is always an important design goal, it is not possible to build the structure arbitrarily robust. Understanding the transient behaviour of the fluids is necessary to predict mechanical loads on the structure and align the design to them. Since a lot of launchers use reactive, cryogenic propellants, it is a common approach to use inert substitute fluids for on ground testing like water (H2O) or liquid nitrogen (LN2). LN2 comes with the advantage of being cryogenic like real propellants, but ground testing is consequently more complex than with H2O. For this purpose, several fluid hammer experiments with both fluids were performed and compared to each other to provide a foundation for deciding which substitute fluid would be useful.

Hot fire tests of a multi-injector research combustor were performed with liquid-oxygen and liquefied-natural-gas (LOX/LNG) propellants at chamber pressures from 30 up to 67 bar, hence at conditions similar to an upper stage rocket engine. Within these tests shear coaxial injectors were tested with and without a recessed LOX post. In both configurations, operating conditions with flames anchored at the LOX post tip and thus, if available, pre-combustion in the recess volume as well as lifted flames were observed. Flame anchoring was indirectly detected via acoustic measurements, using mean speed of sound to indicate the presence of flame in the head end of the combustion chamber. While the injector without recess showed only stable combustion irrespective of the flame anchoring behavior, the recessed injector featured short-lived bursts of oscillatory combustion and sustained combustion instabilities. Analysis of the test data showed that stable flame anchoring could not be ensured at momentum flux ratios below 20 for a non-recessed and below 45 for a recessed injector.

Flame radiation and acoustic measurements were performed in a single-element, liquidoxygen/compressed- and liquefied-natural-gas (LOX/CNG/LNG) rocket combustor with large optical access for sub- and supercritical injection conditions. The operating conditions of this 5 MW combustor are relevant for main and upper stage engines with the propellant combination of LOX/methane. The large optical access enables synchronized high-speed imaging of OH* and CH* radiation covering a great portion of the flame. The experiments showed intermittent, short-lived events of heightened excitation and dangerous high-amplitude and high-frequency combustion instabilities within the analyzed tests. The occurrence of those short-lived events of heightened excitation seem to feature a stochastic distribution. Finally, depending on the chamber configuration, these events triggered high-frequency combustion instabilities. Consolidation of the data gained from different hardware configurations indicated relation of the thermoacoustic behavior to chamber length, oxygen temperature, pressure drop over the oxygen injector and the momentum flux ratio between oxidizer and fuel. Combination with flame radiation measurements showed drastically modified flame shapes in presence of thermoacoustic oscillations. Analysis of the image data focusing on the section near the injection plane reveals an unstable intensity distribution in axial direction.

Reusability of LREs in Europe is increasingly attracting the attention of scientific community and industry with leading projects such as THEMIS, CALLISTO (reusable demonstrators for vertical take-off and landing (VTVL)) and Ariane Next - all powered by the reusable cryogenic Oxygen/Methane (LOX/LCH4) engine "Prometheus". To enable further expansion and cost-effectiveness of the reusability technology for future liquid rocket engines (LREs), research on critical engine components such as turbopumps is crucial. Therefore, within our research we focus on the turbine blade investigation for reusable LRE applications including high cycle fatigue (HCF) and low cycle fatigue (LCF). Validation of defined applied analytical and numerical techniques is established through the Liquid Upper stage deMonstrator ENgine (LUMEN)'s, developed at DLR Lamplodshausen for enhanced expertise in the complete cycle operation for various engine applications, as well as to empower validation studies of the operational conditions to which turbopump components, such as turbine blades, are subjected. Turbine blades are exposed to large thermo-mechanical cyclic strains emerging from an increased temperature driving gas combined with a fast start-up sequence as well as a large rotational speed - essential for acquiring high performance and structural mass efficiency for LREs. Therefore, in addition to bending & torsion as well as thermal gradient and centrifugal forces, it is critical to consider creep effects in durability studies. To forecast the turbine blade fatigue life, analytical (0-D) and numerical approaches for a selected test case are studied. Within the proposed method, a BLISK is assessed for the most severe loading condition considering HCF load by a modified Goodman method, along with a Coffin-Manson based approach for LCF contribution. Each operational cycle under constant maximum loading condition is applied to study the creep effect. As a result, an enhanced fatigue life prediction method including both creep and fatigue conditions for a turbine blade is obtained.

For the development of upper stage rocket engines with laser ignition, the transition of oxidizer and fuel from the pure cryogenic liquid streams to an ignitable mixture needs to be better understood. Due to the near vacuum conditions that are present at high altitudes and in space, the injected fuel rapidly atomizes in a so-called flash boiling process. To investigate the behavior of flashing cryogenic jets under the relevant conditions, experiments of liquid nitrogen have been performed at the DLR Lampoldshausen. The experiments are accompanied by a series of computer simulations and here we use a highly resolved LES to identify 3D effects and to better interpret results from the experiments and existing 2D RANS. It is observed that the vapor generation inside the injector and the evolution of the spray in the combustion chamber differ significantly between the two simulation types due to missing 3D effects and the difference in resolution of turbulent structures. Still, the observed 3D spray dynamics suggest a suitable location for laser ignition that could be found in regions of relative low velocity and therefore expected low strain rates. Further, measured droplet velocities are compared to the velocities of notional Lagrangian particles with similar inertia as the measured droplets. Good agreement between experiments and simulations exists and strong correlation between droplet size and velocity can be demonstrated.

Hot fire tests were performed using a single-injector research combustor featuring a large optical access window ( 255 × 38 mm) for flame visualisation. Three test campaigns were conducted with the propellant combination of liquid oxygen and hydrogen, liquid oxygen and compressed-natural-gas, as well as liquid oxygen and liquefied-natural-gas at conditions relevant for main- and upper-stage rocket engines. The large optical access enabled synchronised flame imaging using ultraviolet and blue radiation wavelengths covering a large portion of the combustion chamber for various sets of sub- and supercritical operating conditions. Combined with temperature, pressure and unsteady pressure measurements, this data provides a high-quality basis for the validation of numerical modelling. Flame width, length and opening angle as features describing the flame topology were extracted from the imaging. The suitability of flame imaging using ultraviolet and blue radiation wavelengths as qualitative markers of heat release was evaluated. Two-dimensional distributions of the Rayleigh Index were calculated for intervals with and without high-amplitude, self-excited oscillations of the longitudinal acoustic resonance modes. The calculated Rayleigh Index values properly reflect the thermoacoustic state of the chamber, indicating that both types of imaging may be used for qualitative study of thermoacoustic coupling of high-pressure cryogenic flames. Keywords Rocket engine , subcritical combustion , supercritical combustion , optical access , liquid oxygen/natural gas , liquid oxygen/hydrogen , single-injector

Hot fire tests were performed using a single-injector research combustor featuring a large optical access window ([Formula: see text] mm) for flame visualisation. Three test campaigns were conducted with the propellant combination of liquid oxygen and hydrogen, liquid oxygen and compressed-natural-gas, as well as liquid oxygen and liquefied-natural-gas at conditions relevant for main- and upper-stage rocket engines. The large optical access enabled synchronised flame imaging using ultraviolet and blue radiation wavelengths covering a large portion of the combustion chamber for various sets of sub- and supercritical operating conditions. Combined with temperature, pressure and unsteady pressure measurements, this data provides a high-quality basis for the validation of numerical modelling. Flame width, length and opening angle as features describing the flame topology were extracted from the imaging. The suitability of flame imaging using ultraviolet and blue radiation wavelengths as qualitative markers of heat release was evaluated. Two-dimensional distributions of the Rayleigh Index were calculated for intervals with and without high-amplitude, self-excited oscillations of the longitudinal acoustic resonance modes. The calculated Rayleigh Index values properly reflect the thermoacoustic state of the chamber, indicating that both types of imaging may be used for qualitative study of thermoacoustic coupling of high-pressure cryogenic flames.

The paper on hand presents the experimental and numerical results of a fatigue life study conducted with Cu-HCP being used as the inner liner material for regeneratively cooled cryogenic rocket combustion chambers. This material is suitable for combustion chambers which are prone to moderate levels of hot-gas surface temperature and wall heat flux. The experimental part of this study uses Thermomechanical Fatigue (TMF) panels made of the high-conductivity Cu-HCP that were cyclically tested to failure at the TMF test bench at DLR Lampoldshausen Institute of Space Propulsion. A TMF panel represents a small section of a regeneratively cooled rocket combustion chamber. It typically consists of 7 cooling channels. The coolant being used is supercritical nitrogen instead of hydrogen or methane due to safety and cost concerns. The TMF test bench also incorporates a high power diode laser radiating onto the TMF panel surface. This provides realistic amounts of heat flux and surface temperature. The laser is cyclically powered on and off to represent the multiple load cycles that liquid rocket engines have to endure particularly in reusable rocket engines. The setup of the TMF test bench provides data regarding the pure mechanical behavior of the material without the influences of any combustion or chemicals. Furthermore, heat flux, surface temperature and mass flow rate can be easily determined, hence providing precise input data for a numerical simulation and validation. The test conditions of the TMF panel were a heat flux of $\dot{\mathrm{q}}=24.25$ MW / m 2 and a maximum surface temperature of $\mathrm{T}_{\mathrm{s}}=800$ K. For numerical modeling of the experiment, ANSYS Mechanical was used. Herein the nonlinear material models implemented within the software package were used. In particular, the kinematic hardening model according to Chaboche, the isotropic hardening and softening model after Voce, Norton's creep fatigue law and the strain rate dependency model according to Peirce were used. The mechanical material parameters for the numerical model were determined by means of an extensive test series including low cycle fatigue (LCF) and tensile tests with uni-axial specimen covering a range of temperatures starting from cryogenic conditions at 77 K towards 1000 K. The numerical fatigue life is estimated by a post processing tool.

To increase liquid rocket engines (LREs) lifetime capability and allow for reusability applications, the efficient evaluation of the most critical subcomponents' remaining useful life plays a vital role. Regeneratively cooled combustion chamber (CC) wall must withstand extremely high loads emerging from a massive temperature gradient between the hot gas and the low temperature of the coolant. The combined loading and unloading operations, together with high temperature and rate dependent inelastic strain, significantly lessen the combustion chamber inner liner life. Within the presented research, the post-processing model was developed for low cycle fatigue (LCF) evaluation of the reusable LRE’s combustion chamber walls. The proposed damage accumulation model is based on the amalgamation of Bonora-Gentile-Pirondi (2004) and Dufailly-Lemaitre (1995) methods, and it incorporates ductile and brittle damage components which are embedded in the post-processing method. Moreover, the required numerical calculation time is further decreased on account of the proposed routine which allows for analysis of only two initial numerically acquired FE cycles. The obtained results based on the developed method combined with coupled thermal-structural quasi 2D Finite Element Analysis (FEA) of the nozzle throat cross-section, were confirmed to be in good agreement with the validation data acquired from the M51 thermo-mechanical laboratory site at DLR Lampoldshausen. The proposed model can be successfully applied for a quick evaluation of the remaining useful life of the CC wall for various rocket engine architectures.

The excitation mechanism of a thermoacoustic instability in a 42-element research rocket thrust chamber with representative operating conditions with respect to European cryogenic rocket engines is investigated in detail. From previous research it was known that the chamber 1T mode can be excited by persistent heat release rate oscillations which are modulated by the resonant modes of the liquid oxygen injectors. The excitation source of the longitudinal injector eigenmodes is investigated in this study. Fibre-optical probes measuring the OH* dynamics from the recess volume of two injectors showed additional frequency content which could neither be explained by the chamber acoustics, nor the acoustics of the injection system. Instead, the temporal evolution of these frequencies correlate with the oxidizer flow velocity. In this work we show that the additional flame modulation originates from a hydrodynamic effect in the injection system. Even though the exact process cannot be precisely identified, an effect designated orifice whistling at the injector inlet orifice seems to be a likely candidate. Combining the new results with previous publications about this combustor, it is now possible to explain past and present observations in terms of the hydrodynamic and thermoacoustic conditions which are necessary for the combustion instability to appear. The conditions, which lead to an injection-driven excitation of the 1T mode are matching frequencies of the 2L mode of the injectors and the chamber 1T mode as well as a Strouhal number between 0.2 and 0.4 based on the length and flow velocity of the injector inlet orifice.

AbstractThe substitution of the toxic hydrazine in current high-altitude rocket engines like upper stages or reaction control thrusters by green propellants is a major key driver in the current technology development of rocket propulsion systems. Operating these kind of rocket engines at high-altitude leads to a sudden pressure drop in the liquid propellants during their injection into the combustion chamber with a near-vacuum atmosphere prior to ignition. The resulting superheated thermodynamic state of the liquid causes a fast and eruptive evaporation which is called flash boiling. The degree of atomisation is important for a successful ignition and a secure operation of the rocket engine. The development and operation of a cryogenic high-altitude test bench at DLR Lampoldshausen enables the systematical experimental characterization of cryogenic flash boiling due to its ability to adjust and control the injection parameters like temperature, pressure or geometry. Several test campaigns with liquid nitrogen (LN2) were performed using two optical diagnostic methods: First, flash boiling LN2 spray patterns were visualised by means of high-speed shadowgraphy and, secondly, we determined the droplet size and velocity distributions in strongly superheated LN2 sprays with the help of a laser-based Phase Doppler system (PDA). The experimental data generated within these measurement campaigns provide defined boundary conditions as well as a broad data base for the numerical modelling of cryogenic flash boiling like e.g. the publications [8, 9].