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Über dieses Buch

This open access book presents the findings of Collaborative Research Center Transregio 40 (TRR40), initiated in July 2008 and funded by the German Research Foundation (DFG). Gathering innovative design concepts for thrust chambers and nozzles, as well as cutting-edge methods of aft-body flow control and propulsion-component cooling, it brings together fundamental research undertaken at universities, testing carried out at the German Aerospace Center (DLR) and industrial developments from the ArianeGroup. With a particular focus on heat transfer analyses and novel cooling concepts for thermally highly loaded structures, the book highlights the aft-body flow of the space transportation system and its interaction with the nozzle flow, which are especially critical during the early phase of atmospheric ascent. Moreover, it describes virtual demonstrators for combustion chambers and nozzles, and discusses their industrial applicability. As such, it is a timely resource for researchers, graduate students and practitioners.

Inhaltsverzeichnis

Open Access

Collaborative Research for Future Space Transportation Systems

Abstract
This chapter book summarizes the major achievements of the five topical focus areas, Structural Cooling, Aft-Body Flows, Combustion Chamber, Thrust Nozzle, and Thrust-Chamber Assembly of the Collaborative Research Center (Sonderforschungsbereich) Transregio 40. Obviously, only sample highlights of each of the more than twenty individual projects can be given here and thus the interested reader is invited to read their reports which again are only a summary of the entire achievements and much more information can be found in the referenced publications. The structural cooling focus area included results from experimental as well as numerical research on transpiration cooling of thrust chamber structures as well as film cooling supersonic nozzles. The topics of the aft-body flow group reached from studies of classical flow separation to interaction of rocket plumes with nozzle structures for sub-, trans-, and supersonic conditions both experimentally and numerically. Combustion instabilities, boundary layer heat transfer, injection, mixing and combustion under real gas conditions and in particular the investigation of the impact of trans-critical conditions on propellant jet disintegration and the behavior under trans-critical conditions were the subjects dealt with in the combustion chamber focus area. The thrust nozzle group worked on thermal barrier coatings and life prediction methods, investigated cooling channel flows and paid special attention to the clarification and description of fluid-structure-interaction phenomena I nozzle flows. The main emphasis of the focal area thrust-chamber assembly was combustion and heat transfer investigated in various model combustors, on dual-bell nozzle phenomena and on the definition and design of three demonstrations for which the individual projects have contributed according to their research field.
Oskar J. Haidn, Nikolaus A. Adams, Rolf Radespiel, Thomas Sattelmayer, Wolfgang Schröder, Christian Stemmer, Bernhard Weigand

Open Access

A Coupled Two-Domain Approach for Transpiration Cooling

Abstract
Transpiration cooling is an innovative cooling concept where a coolant is injected through a porous ceramic matrix composite (CMC) material into a hot gas flow. This setting is modeled by a two-domain approach coupling two models for the hot gas domain and the porous medium to each other by coupling conditions imposed at the interface. For this purpose, appropriate coupling conditions, in particular accounting for local mass injection, are developed. To verify the feasibility of the two-domain approach numerical simulations in 3D are performed for two different application scenarios: a subsonic thrust chamber and a supersonic nozzle.
Valentina König, Michael Rom, Siegfried Müller

Open Access

Innovative Cooling for Rocket Combustion Chambers

Abstract
Transpiration cooling in combination with permeable ceramic-matrix composite materials is an innovative cooling method for rocket engine combustion chambers, while providing high cooling efficiency as well as enhancing engine life time as demanded for future space transportation systems. In order to develop methods and tools for designing transpiration cooled systems, fundamental experimental investigations were performed. An experimental setup consisting of a serial arrangement of four porous carbon fiber reinforced carbon (C/C) samples is exposed to a hot gas flow. Perfused with cold air, the third sample is unperfused in order to assess the wake flow development over the uncooled sample as well as the rebuilding of the coolant layer. Hereby, the focus is on the temperature boundary layer, using a combined temperature/pitot probe. Additionally, the sample surface temperature distribution was measured using IR imaging. The experiments are supported by numerical simulations which are showing a good agreement with measurement data for low blowing ratios.
Jonas Peichl, Andreas Schwab, Markus Selzer, Hannah Böhrk, Jens von Wolfersdorf

Open Access

Film Cooling in Rocket Nozzles

Abstract
In this project supersonic, tangential film cooling in the expansion part of a nozzle with rocket-engine like hot gas conditions was investigated. Therefore, a parametric study in a conical nozzle was conducted revealing the most important influencing parameter on film cooling for the presented setup. Additionally, a new axisymmetric film cooling model and a method for calculating the cooling efficiency from experimental data was developed. These models lead to a satisfying correlation of the data. Furthermore, film cooling in a dual-bell nozzle performing in altitude mode was investigated. The aim of these experiments was to show the influence of different contour inflection geometries on the film cooling efficiency in the bell extension.
Sandra Ludescher, Herbert Olivier

Open Access

Numerical Simulation of Film Cooling in Supersonic Flow

Abstract
High-order direct numerical simulations of film cooling by tangentially blowing cool helium at supersonic speeds into a hot turbulent boundary-layer flow of steam (gaseous H2O) at a free stream Mach number of 3.3 are presented. The stagnation temperature of the hot gas is much larger than that of the coolant flow, which is injected from a vertical slot of height s in a backward-facing step. The influence of the coolant mass flow rate is investigated by varying the blowing ratio F or the injection height s at kept cooling-gas temperature and Mach number. A variation of the coolant Mach number shows no significant influence. In the canonical baseline cases all walls are treated as adiabatic, and the investigation of a strongly cooled wall up to the blowing position, resembling regenerative wall cooling present in a rocket engine, shows a strong influence on the flow field. No significant influence of the lip thickness on the cooling performance is found. Cooling correlations are examined, and a cooling-effectiveness comparison between tangential and wall-normal blowing is performed.
Johannes M. F. Peter, Markus J. Kloker

Open Access

Heat Transfer in Pulsating Flow and Its Impact on Temperature Distribution and Damping Performance of Acoustic Resonators

Abstract
A numerical framework for the prediction of acoustic damping characteristics is developed and applied to a quarter-wave resonator with non-uniform temperature. The results demonstrate a significant impact of the temperature profile on the damping characteristics and hence the necessity of accurate modeling of heat transfer in oscillating flow. Large Eddy Simulations are applied to demonstrate and quantify enhancement in heat transfer induced by pulsations. The study covers wall-normal heat transfer in pulsating flow as well as longitudinal convective effects in oscillating flow. A discussion of hydrodynamic and thermal boundary layers provides insight into the flow physics of oscillatory convective heat transfer.
Simon van Buren, Wolfgang Polifke

Open Access

Effects of a Launcher’s External Flow on a Dual-Bell Nozzle Flow

Abstract
Previous research on Dual-Bell nozzle flow always neglected the influence of the outer flow on the nozzle flow and its transition from sea level to altitude mode. Therefore, experimental measurements on a Dual-Bell nozzle with trans- and supersonic external flows about a launcher-like forebody were carried out in the Trisonic Wind Tunnel Munich with particle image velocimetry, static pressure measurements and the schlieren technique. A strongly correlated interaction exists between a transonic external flow with the nozzle flow in its sea level mode. At supersonic external flow conditions, a Prandtl–Meyer expansion about the nozzle’s lip decreases the pressure in the vicinity of the nozzle exit by about 55%. Therefore a new definition for the important design criterion of the nozzle pressure ratio was suggested, which considers this drastic pressure drop. Experiments during transitioning of the nozzle from sea level to altitude mode show that an interaction about the nozzle’s lip causes an inherently unstable nozzle state at supersonic free-stream conditions. This instability causes the nozzle to transition and retransition, or flip-flop, between its two modes. This instability can be eliminated by designing a Dual-Bell nozzle to transition during sub-/transonic external flow conditions.
Istvan Bolgar, Sven Scharnowski, Christian J. Kähler

Open Access

Interaction of Wake and Propulsive Jet Flow of a Generic Space Launcher

Abstract
This work investigates the interaction of the afterbody flow with the propulsive jet flow on a generic space launcher equipped with two alternative nozzle concepts and different afterbody geometries. The flow phenomena are characterized by experimental measurements and numerical URANS and LES simulations. Investigations concern a configuration with a conventional truncated ideal contour nozzle and a configuration with an unconventional dual-bell nozzle. In order to attenuate the dynamic loads on the nozzle fairing, passive flow control devices at the base of the launcher main body are investigated on the configuration with TIC nozzle. The nozzle Reynolds number and the afterbody geometry are varied for the configuration with dual-bell nozzle. The results for integrated nozzles show a shift of the nozzle pressure ratio for transition from sea-level to altitude mode to significant lower levels. The afterbody geometry is varied including a reattaching and non-reattaching outer flow on the nozzle fairing. Investigations are performed at supersonic outer flow conditions with a Mach number of $$Ma_\infty =3$$. It turns out, that a reattachment of the outer flow on the nozzle fairing leads to an unstable nozzle operation.

Open Access

Rocket Wake Flow Interaction Testing in the Hot Plume Testing Facility (HPTF) Cologne

Abstract
Rocket wake flows were under investigation within the Collaborative Research Centre SFB/TRR40 since the year 2009. The current paper summarizes the work conducted during its third and final funding period from 2017 to 2020. During that phase, focus was laid on establishing a new test environment at the German Aerospace Center (DLR) Cologne in order to improve the similarity of experimental rocket wake flow–jet interaction testing by utilizing hydrogen–oxygen combustion implemented into the wind tunnel model. The new facility was characterized during tests with the rocket combustor model HOC1 in static environment. The tests were conducted under relevant operating conditions to demonstrate the design’s suitability. During the first wind tunnel tests, interaction of subsonic ambient flow at Mach 0.8 with a hot exhaust jet of approx. 920 K was compared to previously investigated cold plume interaction tests using pressurized air at ambient temperature. The comparison revealed significant differences in the dynamic response of the wake flow field on the different types of exhaust plume simulation.
Daniel Kirchheck, Dominik Saile, Ali Gülhan

Open Access

Numerical Analysis of the Turbulent Wake for a Generic Space Launcher with a Dual-Bell Nozzle

Abstract
The turbulent wake of an axisymmetric generic space launcher equipped with a dual-bell nozzle is simulated at transonic ($$Ma_\infty = 0.8$$ and $$Re_D = 4.3\cdot 10^5$$) and supersonic ($$Ma_\infty = 3$$ and $$Re_D = 1.2\cdot 10^6$$) freestream conditions, to investigate the influence of the dual-bell nozzle jet onto the wake flow and vice versa. In addition, flow control by means of four in circumferential direction equally distributed jets injecting air encountering the backflow in the recirculation region is utilized to determine if the coherence of the wake and consequently, the buffet loads can be reduced by flow control. The simulations are performed using a zonal RANS/LES approach. The time-resolved flow field data are analyzed by classical spectral analysis, two-point correlation analysis, and dynamic mode decomposition (DMD). At supersonic freestream conditions, the nozzle counter pressure is reduced by the expansion of the outer flow around the nozzle lip leading to a decreased transition nozzle pressure ratio. In the transonic configuration a spatio-temporal mode with an eigenvalue matching the characteristic buffet frequency of $$Sr_D=0.2$$ is extracted by the spectral and DMD analysis. The spatial shape of the detected mode describes an antisymmetric wave-like undulating motion of the shear layer inducing the low frequency dynamic buffet loads. By flow control this antisymmetric coherent motion is weakened leading to a reduction of the buffet loads on the nozzle fairing.
Simon Loosen, Matthias Meinke, Wolfgang Schröder

Open Access

Numerical Investigation of Space Launch Vehicle Base Flows with Hot Plumes

Abstract
The flow field around generic space launch vehicles with hot exhaust plumes is investigated numerically. Reynolds-Averaged Navier-Stokes (RANS) simulations are thermally coupled to a structure solver to allow determination of heat fluxes into and temperatures in the model structure. The obtained wall temperatures are used to accurately investigate the mechanical and thermal loads using Improved Delayed Detached Eddy Simulations (IDDES) as well as RANS. The investigated configurations feature cases both with cold air and hot hydrogen/ water vapour plumes as well as cold and hot wall temperatures. It is found that the presence of a hot plume increases the size of the recirculation region and changes the pressure distribution on the nozzle structure and thus the loads experienced by the vehicle. The same effect is observed when increasing the wall temperatures. Both RANS and IDDES approaches predict the qualitative changes between the configurations, but the reattachment location predicted by IDDES is up to 7% further upstream than that predicted by RANS. Additionally, the heat flux distribution along the nozzle and base surface is analysed and shows significant discrepancies between RANS and IDDES, especially on the nozzle surface and in the base corner.
Jan-Erik Schumann, Markus Fertig, Volker Hannemann, Thino Eggers, Klaus Hannemann

Open Access

On the Consideration of Diffusive Fluxes Within High-Pressure Injections

Abstract
Mixing characteristics of supercritical injection studies were analyzed with regard to the necessity to include diffusive fluxes. Therefore, speed of sound data from mixing jets were investigated using an adiabatic mixing model and compared to an analytic solution. In this work, we show that the generalized application of the adiabatic mixing model may become inappropriate for subsonic submerged jets at high-pressure conditions. Two cases are discussed where thermal and concentration driven fluxes are seen to have significant influence. To which extent the adiabatic mixing model is valid depends on the relative importance of local diffusive fluxes, namely Fourier, Fick and Dufour diffusion. This is inter alia influenced by different time and length scales. The experimental data from a high-pressure n-hexane/nitrogen jet injection were investigated numerically. Finally, based on recent numerical findings, the plausibility of different thermodynamic mixing models for binary mixtures under high pressure conditions is analyzed.
Fabian Föll, Valerie Gerber, Claus-Dieter Munz, Berhand Weigand, Grazia Lamanna

Open Access

Numerical Investigation of Injection, Mixing and Combustion in Rocket Engines Under High-Pressure Conditions

Abstract
The design and development of future rocket engines severely relies on accurate, efficient and robust numerical tools. Large-Eddy Simulation in combination with high-fidelity thermodynamics and combustion models is a promising candidate for the accurate prediction of the flow field and the investigation and understanding of the on-going processes during mixing and combustion. In the present work, a numerical framework is presented capable of predicting real-gas behavior and nonadiabatic combustion under conditions typically encountered in liquid rocket engines. Results of Large-Eddy Simulations are compared to experimental investigations. Overall, a good agreement is found making the introduced numerical tool suitable for the high-fidelity investigation of high-pressure mixing and combustion.
Christoph Traxinger, Julian Zips, Christian Stemmer, Michael Pfitzner

Open Access

Large-Eddy Simulations for the Wall Heat Flux Prediction of a Film-Cooled Single-Element Combustion Chamber

Abstract
In order for modern launcher engines to work at their optimum, film cooling can be used to preserve the structural integrity of the combustion chamber. The analysis of this cooling system by means of CFD is complex due to the extreme physical conditions and effects like turbulent fluctuations damping and recombination processes in the boundary layer which locally change the transport properties of the fluid. The combustion phenomena are modeled by means of Flamelet tables taking into account the enthalpy loss in the proximity of the chamber walls. In this work, Large-Eddy Simulations of a single-element combustion chamber experimentally investigated at the Technical University of Munich are carried out at cooled and non-cooled conditions. Compared with the experiment, the LES shows improved results with respect to RANS simulations published. The influence of wall roughness on the wall heat flux is also studied, as it plays an important role for the lifespan of a rocket engine combustors.
R. Olmeda, P. Breda, C. Stemmer, M. Pfitzner

Open Access

Calculation of the Thermoacoustic Stability of a Main Stage Thrust Chamber Demonstrator

Abstract
The stability behavior of a virtual thrust chamber demonstrator with low injection pressure loss is studied numerically. The approach relies on an eigenvalue analysis of the Linearized Euler Equations. An updated form of the stability prediction procedure is outlined, addressing mean flow and flame response calculations. The acoustics of the isolated oxidizer dome are discussed as well as the complete system incorporating dome and combustion chamber. The coupling between both components is realized via a scattering matrix representing the injectors. A flame transfer function is applied to determine the damping rates. Thereby it is found that the procedure for the extraction of the flame transfer function from the CFD solution has a significant impact on the stability predictions.
Alexander Chemnitz, Thomas Sattelmayer

Open Access

Experimental Investigation of Injection-Coupled High-Frequency Combustion Instabilities

Abstract
Self-excited high-frequency combustion instabilities were investigated in a 42-injector cryogenic rocket combustor under representative conditions. In previous research it was found that the instabilities are connected to acoustic resonance of the shear-coaxial injectors. In order to gain a better understanding of the flame dynamics during instabilities, an optical access window was realised in the research combustor. This allowed 2D visualisation of supercritical flame response to acoustics under conditions similar to those found in European launcher engines. Through the window, high-speed imaging of the flame was conducted. Dynamic Mode Decomposition was applied to analyse the flame dynamics at specific frequencies, and was able to isolate the flame response to injector or combustion chamber acoustic modes. The flame response at the eigenfrequencies of the oxygen injectors showed symmetric and longitudinal wave-like structures on the dense oxygen core. With the gained understanding of the BKD coupling mechanism it was possible to derive LOX injector geometry changes in order to reduce the risks of injection-coupled instabilities for future cryogenic rocket engines.
Wolfgang Armbruster, Justin S. Hardi, Michael Oschwald

Open Access

Pseudo-transient 3D Conjugate Heat Transfer Simulation and Lifetime Prediction of a Rocket Combustion Chamber

Abstract
Rocket engine nozzle structures typically fail after a few engine cycles due to the extreme thermomechanical loading near the nozzle throat. In order to obtain an accurate lifetime prediction and to increase the lifetime, a detailed understanding of the thermomechanical behavior and the acting loads is indispensable. The first part is devoted to a thermally coupled simulation (conjugate heat transfer) of a fatigue experiment. The simulation contains a thermal FEM model of the fatigue specimen structure, RANS simulations of nine cooling channel flows and a Flamelet-based RANS simulation of the hot gas flow. A pseudo-transient, implicit Dirichlet–Neumann scheme is utilized for the partitioned coupling. A comparison with the experiment shows a good agreement between the nodal temperatures and their corresponding thermocouple measurements. The second part consists of the lifetime prediction of the fatigue experiment utilizing a sequentially coupled thermomechanical analysis scheme. First, a transient thermal analysis is carried out to obtain the temperature field within the fatigue specimen. Afterwards, the computed temperature serves as input for a series of quasi-static mechanical analyses, in which a viscoplastic damage model is utilized. The evolution and progression of the damage variable within the regions of interest are thoroughly discussed. A comparison between simulation and experiment shows that the results are in good agreement. The crucial failure mode (doghouse effect) is captured very well.
Oliver Barfusz, Felix Hötte, Stefanie Reese, Matthias Haupt

Open Access

Lifetime Experiments of Regeneratively Cooled Rocket Combustion Chambers and PIV Measurements in a High Aspect Ratio Cooling Duct

Abstract
This paper aims at experimental investigations of the life limiting mechanisms of regeneratively cooled rocket combustion chambers, especially the so called doghouse effect. In this paper the set up of a cyclic thermo-mechanical fatigue experiment and its results are shown. This experiment has an actively cooled fatigue specimen that is mounted downstream of a subscale GOX-GCH$$_{\text {4}}$$ combustion chamber with rectangular cross section. The specimen is loaded cyclically and inspected after each cycle. The effects of roughness, the use of thermal barrier coatings, the length of the hot gas phase, the oxygen/fuel ratio and the hot gas pressure are shown. In a second experiment the flow in a generic high aspect ratio cooling duct is measured with the Particle Image Velocimetry (PIV) to characterize the basic flow. The main focus of the analysis is on the different recording and processing parameters of the PIV method. Based on this analysis a laser pulse interval and the window size for auto correlation is chosen. Also the repeatability of the measurements is demonstrated. These results are the starting point for future measurements on the roughness effect on heat transfer and pressure loss in a high aspect ratio cooling duct.
Felix Hötte, Oliver Günther, Christoph von Sethe, Matthias Haupt, Peter Scholz, Michael Rohdenburg

Open Access

Mechanical Integrity of Thermal Barrier Coatings: Coating Development and Micromechanics

Abstract
To protect the copper liners of liquid-fuel rocket combustion chambers, a thermal barrier coating can be applied. Previously, a new metallic coating system was developed, consisting of a NiCuCrAl bond-coat and a Rene 80 top-coat, applied with high velocity oxyfuel spray (HVOF). The coatings are tested in laser cycling experiments to develop a detailed failure model, and critical loads for coating failure were defined. In this work, a coating system is designed for a generic engine to demonstrate the benefits of TBCs in rocket engines, and the mechanical loads and possible coating failure are analysed. Finally, the coatings are tested in a hypersonic wind tunnel with surface temperatures of 1350 K and above, where no coating failure was observed. Furthermore, cyclic experiments with a subscale combustion chamber were carried out. With a diffusion heat treatment, no large-scale coating delamination was observed, but the coating cracked vertically due to large cooling-induced stresses. These cracks are inevitable in rocket engines due to the very large thermal-strain differences between hot coating and cooled substrate. It is supposed that the cracks can be tolerated in rocket-engine application.
Torben Fiedler, Joachim Rösler, Martin Bäker, Felix Hötte, Christoph von Sethe, Dennis Daub, Matthias Haupt, Oskar J. Haidn, Burkard Esser, Ali Gülhan

Open Access

Assessment of RANS Turbulence Models for Straight Cooling Ducts: Secondary Flow and Strong Property Variation Effects

Abstract
We present well-resolved RANS simulations of two generic asymmetrically heated cooling channel configurations, a high aspect ratio cooling duct operated with liquid water at $$Re_b = 110 \times 10^3$$ and a cryogenic transcritical channel operated with methane at $$Re_b = 16 \times 10^3$$. The former setup serves to investigate the interaction of turbulence-induced secondary flow and heat transfer, and the latter to investigate the influence of strong non-linear thermodynamic property variations in the vicinity of the critical point on the flow field and heat transfer. To assess the accuracy of the RANS simulations for both setups, well-resolved implicit LES simulations using the adaptive local deconvolution method as subgrid-scale turbulence model serve as comparison databases. The investigation focuses on the prediction capabilities of RANS turbulence models for the flow as well as the temperature field and turbulent heat transfer with a special focus on the turbulent heat flux closure influence.
Thomas Kaller, Alexander Doehring, Stefan Hickel, Steffen J. Schmidt, Nikolaus A. Adams

Open Access

Experiments on Aerothermal Supersonic Fluid-Structure Interaction

Abstract
Mastering aerothermal fluid-structure interaction (FSI) is crucial for the efficient and reliable design of future (reusable) launch vehicles. However, capabilities in this area are still quite limited. To address this issue, a multidisciplinary experimental and numerical study of such problems was conducted within SFB TRR 40. Our work during the last funding period was focused on studying the effects of moderate and high thermal loads. This paper provides an overview of our experiments on FSI including structural dynamics and thermal effects for configurations in two different flow regimes. The first setup was designed to study the combined effects of thermal and pressure loads. We investigated a range of conditions including shock-wave/boundary-layer interaction (SWBLI) with various incident shock angles leading to, in some cases, large flow separation with high amplitude temperature dependent panel oscillations. The respective aerothermal loads were studied in detail using a rigid reference panel. The second setup allowed us to study the effects of severe heating leading to plastic deformation of the structure. We obtained severe localized heating resulting in partly plastic deformations of more than 12 times the panel thickness. Furthermore, the effects of repeated load cycles were studied.
Dennis Daub, Sebastian Willems, Burkard Esser, Ali Gülhan

Open Access

Numerical Modelling of Fluid-Structure Interaction for Thermal Buckling in Hypersonic Flow

Abstract
Experiments have shown that a high-enthalpy flow field might lead under certain mechanical constraints to buckling effects and plastic deformation. The panel buckling into the flow changes the flow field causing locally increased heating which in turn affects the panel deformation. The temperature increase due to aerothermal heating in the hypersonic flow causes the metallic panel to buckle into the flow. To investigate these phenomena numerically, a thermomechanical simulation of a fluid-structure interaction (FSI) model for thermal buckling is presented. The FSI simulation is set up in a staggered scheme and split into a thermal solid, a mechanical solid and a fluid computation. The structural solver Abaqus and the fluid solver TAU from the German Aerospace Center (DLR) are coupled within the FSI code ifls developed at the Institute of Aircraft Design and Lightweight Structures (IFL) at TU Braunschweig. The FSI setup focuses on the choice of an equilibrium iteration method, the time integration and the data transfer between grids. To model the complex material behaviour of the structure, a viscoplastic material model with linear isotropic hardening and thermal expansion including material parameters, which are nonlinearly dependent on temperature, is used.
Katharina Martin, Dennis Daub, Burkard Esser, Ali Gülhan, Stefanie Reese

Open Access

Experimental and Numerical Investigation of CH/O Rocket Combustors

Abstract
The experimental investigation of sub-scale rocket engines gives significant information about the combustion dynamics and wall heat transfer phenomena occurring in full-scale hardware. At the same time, the performed experiments serve as validation test cases for numerical CFD models and for that reason it is vital to obtain accurate experimental data. In the present work, an inverse method is developed able to accurately predict the axial and circumferential heat flux distribution in CH$$_4$$/O$$_2$$ rocket combustors. The obtained profiles are used to deduce information about the injector-injector and injector-flame interactions. Using a 3D CFD simulation of the combustion and heat transfer within a multi-element thrust chamber, the physical phenomena behind the measured heat flux profiles can be inferred. A very good qualitative and quantitative agreement between the experimental measurements and the numerical simulations is achieved.
Nikolaos Perakis, Oskar J. Haidn

Open Access

Rocket Combustion Chamber Simulations Using High-Order Methods

Abstract
High-order spatial discretizations significantly improve the accuracy of flow simulations. In this work, a multi-dimensional limiting process with low diffusion (MLP$$^\text {ld}$$) and up to fifth order accuracy is employed. The advantage of MLP is that all surrounding volumes of a specific volume may be used to obtain cell interface values. This prevents oscillations at oblique discontinuities and improves convergence. This numerical scheme is utilized to investigate three different rocket combustors, namely a seven injector methane/oxygen combustion chamber, the widely simulated PennState preburner combustor and a single injector chamber called BKC, where pressure oscillations are important.
Timo Seitz, Ansgar Lechtenberg, Peter Gerlinger

Open Access

Dual-Bell Nozzle Design

Abstract
The dual-bell nozzle is an altitude adaptive nozzle concept that offers two operation modes. In the framework of the German Research Foundation Special Research Field SFB TRR40, the last twelve years have been dedicated to study the dual-bell nozzle characteristics, both experimentally and numerically. The obtained understanding on nozzle contour and inflection design, transition behavior and transition prediction enabled various follow-ups like a wind tunnel study on the dual-bell wake flow, a shock generator study on a film cooled wall inflection or, in higher scale, the hot firing test of a thrust chamber featuring a film cooled dual-bell nozzle. A parametrical system study revealed the influence of the nozzle geometry on the flow behavior and the resulting launcher performance increase.
Chloé Génin, Dirk Schneider, Ralf Stark

Open Access

Definition and Evaluation of Advanced Rocket Thrust Chamber Demonstrator Concepts

Abstract
Since the beginning of the German collaborative research center SFB-TRR 40 in 2008 ArianeGroup has been involved as industrial partner and supported the research activities with its expertise. For the final funding period ArianeGroup actively contributes to the SFB-TRR 40 with the self-financed project K4. Within project K4 virtual thrust chamber demonstrators have been defined that allow the application of the attained knowledge of the entire collaborative research center to state-of-the-art numerical benchmark cases. Furthermore, ArianeGroup uses these testcases to continue the development of its in-house spray combustion and performance analysis tool Rocflam3. Unique within the collaborative research center fully three-dimensional conjugate heat transfer computations have been performed for a full-scale 100 kN upper stage thrust chamber. The strong three-dimensionality of the temperature field in the structure resulting from injection element and cooling channel configuration is displayed.
Daniel Eiringhaus, Hendrik Riedmann, Oliver Knab
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