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The sensitivity of hybrid RANS-LES methods like Improved Delayed Detached Eddy Simulation (IDDES) to numerical model parameter variations related to generic space launch vehicle aft-body flows is investigated. In particular, the changes resulting from the choice of the time-step size, the turbulence model, the fluid modelling, the circumferential grid resolution, the filter length definition, and the data collection period is considered. The results are also compared to experimental and numerical data taken from the available literature. The sensitivity to the time-step size and the turbulence model is minuscule with respect to the obtained mean flow field, wall pressure distributions, azimuthal modes, and wall pressure frequency spectra. However, circumferential resolution, fluid model, and filter length definition affect the solution to a higher extent. Buffeting spectra are very sensitive to the data collection period.

A subscale flight experiment configuration propelled by a Mach 8 supersonic combustion ramjet (scramjet) was designed within the framework of the European Commission co-funded Long Term Advanced Propulsion Concepts and Technologies II project. The focus of this design exercise was to verify by ground testing the ability of the proposed scramjet engine to produce adequate thrust for hypersonic level flight. Experiments performed in DLR’s High Enthalpy Shock Tunnel Göttingen confirmed precedent CFD predictions of the total thrust and demonstrated the operability of the vehicle. Yet, significant discrepancies between the CFD analyses, which were performed to design the vehicle, and subsequent detailed measurements of the pressure distribution in the combustor were observed and could not be resolved so far. This paper focuses on a further analysis of these residual discrepancies. It was found that the CFD predictions of the combustor pressure distribution are sensitive to the configuration of the intake boundary layer. Particularly, different assumptions for the location of the laminar to turbulent boundary layer transition strongly influence cross-flow structures which develop on the intake and which are able to trigger different combustion modes in the combustor. While the effect on total vehicle performance remains limited, a significant impact on the structure and magnitude of the surface pressure distribution was observed. i.e., the large combustor peak pressures, which occur in the experiment, can be explained by the occurrence of a strong shock train in the vicinity of the combustor wall.

Experiments are conducted on a planar ramp in high-enthalpy flow, with gaseous hydrogen injected through a porthole at the ramp surface. The ramp is covered in a thin layer of temperature-sensitive luminescent paint, originally intended to allow surface heat-flux measurements; however, anomalous transient features are identified in high-speed recorded images of the paint emission, which are postulated as originating from excitation by OH* chemiluminescence produced during hydrogen combustion. By assuming the OH*-induced emission to occur on much shorter timescales than surface temperature changes, it is possible to extract this pseudo-OH* signal, allowing time-resolved combustion features to be investigated. Two cases are examined, the first exhibiting marginal combustion and the second with more robust ignition. For the first, the weak OH* excitation allows concurrent temperature and heat-flux distributions to be derived. In both cases, the time-averaged combustion distribution shows intense ignition in the immediate vicinity of the injection hole, with extended fingers of weaker combustion extending downstream; the latter are assumed to be associated with the main counter-rotating vortex pair produced by the injection. By examining shorter timescales, however, it is revealed that these combustion fingers are not steady features but rather the time-integrated result of more localized ignition streaks convecting downstream. Overall, the presence of combustion is associated with higher surface heat fluxes. The pseudo-OH* visualization method is a possible alternative to conventional OH* chemiluminescence for wall-bounded combusting flows, allowing a conventional visible-wavelength camera to be employed in place of a UV-sensitive imaging system.

Depending on the cone half-angle and the inverse normal-shock density ratio $\unicode[STIX]{x1D700}$, hypersonic flow over a spherically blunted cone exhibits two regimes separated by an almost discontinuous jump of the body end of the sonic line from a point on the spherical nose to the shoulder of the cone, here called sphere behaviour and cone behaviour. The inflection point of the shock wave in sphere behaviour is explained. In Part 1 we explore the two elements of the capsule shape, the sphere and the sharp cone with detached shock, theoretically and computationally, in order to put the treatment of the full capsule shape on a sound basis. Starting from the analytical expression for the shock detachment angle of a cone given by Hayes & Probstein (Hypersonic Flow Theory, 1959, Academic Press) we make a hypothesis for the sharp cone, about the functional form of the dependence of dimensionless quantities on $\unicode[STIX]{x1D700}$ and a cone angle parameter, $\unicode[STIX]{x1D702}$. In the critical part of atmospheric entry the shock shape and drag of the capsule are insensitive to viscous effects, so that much can be learned from inviscid studies. Accordingly, the hypothesis is tested by making a large number of Euler computations to cover the parameter space: Mach number, specific heat ratio and cone angle. The results confirm the hypothesis in the case of the dimensionless shock stand-off distance as well as for the drag coefficient, yielding accurate analytical functions for both. This reduces the number of independent parameters of the problem from three to two. A functional form of the shock stand-off distance is found for the transition from the $90^{\circ }$ cone to the sphere. Although the analysis assumes a calorically perfect gas, the results may be carried over to the high-enthalpy real-gas situation if the normal-shock density ratio is replaced by the density ratio based on the average density along the stagnation streamline (see e.g. Stulov, Izv. AN SSSR Mech. Zhidk. Gaza, vol. 4, 1969, pp. 142–146; Hornung, J. Fluid Mech., vol. 53, 1972, pp. 149–176; Wen & Hornung, J. Fluid Mech., vol. 299, 1995, pp. 389–405).

We investigate the flow in planar microscale nozzles and find that design and analysis paradigms based on the assumption of a dominant isentropic core with moderate viscosity corrections are not valid. Instead, the flow downstream of the throat is dominated by boundary layers that may choke the flow to subsonic velocities. The geometrical expansion ratio is found to be essentially irrelevant, instead, the length from throat to exit plane is found to be a much more important design parameter. Full 3D simulations are required to predict the flow topology; thermophysical modeling of the expanding gas has a noticeable impact on predicted performance. An analytical estimation of the Knudsen number in the expanding flow is given, allowing to determine its values from the expansion pressure ratio. An axial thrust analysis suggests truncation of the nozzle, resulting in a predicted 30% increase in thrust and 30% increase in specific impulse compared to the baseline configuration. The work has been carried out within the European Commission co-funded PRECISE project which was focused on designing and testing a micro chemical propulsion system thruster prototype using catalytically decomposed hydrazine as propellant.

The accurate numerical simulation of high-pressure rocket combustion chambers is a key element for the design of future space transportation vehicles. In recent years investigations focus more and more on methane-based combustion chambers for which many thermodynamic and combustion issues are still unknown. Experimental investigations indicate that we might have to deal with a close interaction of non-ideal equation-of-state effects together with complex kinetic schemes.

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.

International audience; The transition from old space to new space along with increasing commercialization has a major impact on space flight, in general, and on electric propulsion (EP) by ion thrusters, in particular. Ion thrusters are nowadays used as primary propulsion systems in space. This article describes how these changes related to new space affect various aspects that are important for the development of EP systems. Starting with a historical overview of the development of space flight and of the technology of EP systems, a number of important missions with EP and the underlying technologies are presented. The focus of our discussion is the technology of the radio frequency ion thruster as a prominent member of the gridded ion engine family. Based on this discussion, we give an overview of important research topics such as the search for alternative propellants, the development of reliable neutralizer concepts based on novel insert materials, as well as promising neutralizer-free propulsion concepts. In addition, aspects of thruster modeling and requirements for test facilities are discussed. Furthermore, we address aspects of space electronics with regard to the development of highly efficient electronic components as well as aspects of electromagnetic compatibility and radiation hardness. This article concludes with a presentation of the interaction of EP systems with the spacecraft.

The flow through a thin slit and a thin orifice is studied with the Direct Simulation Monte Carlo (DSMC), cubic Fokker–Planck (FP), and a coupled FP-DSMC hybrid method. Pressure driven monatomic gas flows through a slit and an orifice with various values of degree of rarefaction and pressure ratio are computed. The DSMC method is physically accurate for all flow regimes; however it is computationally expensive in high density, near continuum regions. An alternative stochastic particle scheme, the cubic FP kinetic model has addressed this issue by approximating the particle collisions involved in the Boltzmann collision integral with continuous stochastic processes. The ability of the cubic FP method to reproduce breakdown of translational equilibrium is discussed. In addition, a coupled FP-DSMC hybrid scheme is employed aiming at an efficient and accurate solution. The FP-DSMC hybrid scheme employs DSMC in rarefied regions and FP method in near continuum flow regions. Numerical procedures of the cubic FP method are implemented within the framework of an existing DSMC-solver, SPARTA. The FP-DSMC hybrid solution reproduces pure DSMC solution with improved computational efficiency up to a factor of eight for vacuum flow through a thin slit.

The sensitivity to grid changes of a detaching-reattaching shear layer flow over a generic space launch vehicle geometry is investigated with Improved Delayed Detached Eddy Simulation (IDDES). First and second order statistical moments of the flow are compared and show good agreement with experimental and numerical data from literature. A sufficient axial resolution in the initial shear layer region is found to be crucial for an accurate capturing of the spectral content of the flow. Results on a grid with prismatic elements compare well to those on grids with hexahedral elements. A finer circumferential resolution changes the flow field drastically and leads to a merging of the first and second recirculation region, also affecting the pressure field and spectral features. Additionally, an absolute and a directional grid sensors are used that are found to be well suited to support anisotropic grid refinement.

In this work we investigate the kinetic Fokker-Planck (FP) model [1, 2] concerning its applicability to simulate hyper-sonic, rarefied gas flow. As in the direct simulation Monte Carlo (DSMC) algorithm, the FP model simulates the stochastic motion of a set of particles through the computational domain which results in a simple coupling of both methods. No collisions must be calculated in the FP approach and so larger cell and time step sizes than by DSMC calculations can be used. Because of this, the FP model holds the potential to be more efficient than DSMC where the Knudsen number is small. Further we present and investigate a hybridisation scheme for the FP and DSMC algorithms. We derive a non-equilibrium parameter from the Fokker-Planck- and Boltzmann-operator which indicates the validity of the FP model. In hybrid simulations this parameter is used to partition the computational domain in DSMC and FP regions. The particle motion is processed similarly in both regions, only the procedure for assigning new particle velocities differs. To achieve an equal number of particles per cell, weighting factors depended on the cell size are used. The hybridisation scheme is tested for one-dimensional shock and two-dimensional cylinder flows and gives results in agreement to pure DSMC simulations. An efficiency study shows, that the hybrid scheme is roughly five times faster than a pure DSMC simulation for a hypersonic cylinder flow test case.

An important task in the development of electric propulsion devices is ground-based testing. In order to characterize electric thrusters, their generated ion beams are regarded as figures of merit and have to be well-known and understood. We present an approach to measure DC ion beam currents generated by electric thrusters by means of an inductive direct-current current transformer. In contrast to commonly used devices like Faraday cups, this device can measure the beam current non-intrusively by means of magnetic field coupling rather than charge collection. This paper shows the development of a prototype sensor which can measure currents in the typical range for electric thrusters with high resolution and satisfactory accuracy. A detailed SPICE model to assist during hardware development is also introduced and verified by test case measurements. The prototype has been tested and validated with a radio-frequency ion thruster. Its readout shows very good agreement with the output of an analytical model which computes a Gaussian-shaped ion beam in the far-field based on experimental input data.

Global heat-flux measurements are performed using a newly developed temperature-sensitive paint (TSP) on an inclined ramp with sidewalls in a hypersonic shock tunnel. The paint response and image acquisition rate are sufficiently fast to allow flow phenomena on timescales of around $$100\,\upmu \mathrm{{s}}$$ to be resolved. Although a priori calibration of the new TSP proves inaccurate, in situ calibration allows the recovery of heat fluxes that agree well with embedded thermocouple measurements on both short and long timescales. At low unit Reynolds numbers, the flow on the main ramp surface is entirely laminar, but transition occurs in the corner-flow regions, causing a turbulent region to spread inwards from each sidewall and producing weak, unsteady features in the heat-flux distribution of the main laminar region. Within this laminar region, roughly steady streamwise streaks with a period of approximately ten times the boundary-layer thickness are also observed. At higher unit Reynolds numbers, the boundary layer on the main ramp surface transitions to turbulence. The fast-response TSP allows tracking of the time-resolved transition front: significant unsteadiness is observed, which appears to be only weakly correlated to unsteadiness in the freestream flow conditions. Based on the heat-flux signature in the transition region, the breakdown mechanism seems to be quite different from that observed in earlier measurements on a slender cone at similar conditions.

This paper derives from the cooperation between DLR and Airbus DS within the work package “CFD Modeling of Combustion Chamber Processes” conducted in the frame of the Propulsion 2020 project. In a joint strategy, DLR Göttingen and Airbus DS Ottobrunn have identified a number of test cases with gradually growing complexity where adequate test data are available for proper successive validation of the computational fluid dynamics (CFD) tools to be used in an industrial environment. This work highlights the simulation results for the Mascotte A-10 and A-60 test cases as presented at the 2nd International Workshop on Rocket Combustion Modeling in Lampoldshausen 2001 by ONERA and SNECMA [1]. These two test cases are characterized by different chamber pressures (10 and 60 bar) and, consequently, by oxygen injection conditions which are subcritical in one case and transcritical in the other case. The test cases are treated with three different CFD codes: the DLR TAU Code (only A-60 case), the Airbus DS in-house tool Rocflam3, and the commercial CFD tool ANSYS CFX incorporating several modeling extensions by Airbus DS. To the knowledge of the authors, this paper is the first one which covers both the A-10 and the A-60 test cases.

The impact of two geometric configurations of sonic orifices on the free plume expansion into high vacuum is studied both experimentally and numerically. One configuration is a flat-faced cylinder, into whose face a square edged orifice is centrally machined. This simple type of orifice is often used for ease of manufacturing. The other source features a sharp-edged orifice, whose outer walls are inclined backwards 135o from the axis. Nitrogen plumes emanating at two source Reynolds numbers, Re* ∈ {3850, 15400}, are compared here. Mass flux measurements with a free-molecular impact pressure probe (“Patterson probe”) show, that the plume originating from the flat-faced orifice expands freely within about 45 o from the axis. Numerical analysis of the plume near-field suggests, that molecules scattered back from the walls of a conical orifice only affect the plume expansion at angles greater than about 80o from the axis.The impact of two geometric configurations of sonic orifices on the free plume expansion into high vacuum is studied both experimentally and numerically. One configuration is a flat-faced cylinder, into whose face a square edged orifice is centrally machined. This simple type of orifice is often used for ease of manufacturing. The other source features a sharp-edged orifice, whose outer walls are inclined backwards 135o from the axis. Nitrogen plumes emanating at two source Reynolds numbers, Re* ∈ {3850, 15400}, are compared here. Mass flux measurements with a free-molecular impact pressure probe (“Patterson probe”) show, that the plume originating from the flat-faced orifice expands freely within about 45 o from the axis. Numerical analysis of the plume near-field suggests, that molecules scattered back from the walls of a conical orifice only affect the plume expansion at angles greater than about 80o from the axis.

Spherically blunted cones are commonly utilized as re-entry or entry capsules. Depending on the cone angle, the shock shape is either dominated by the spherical nose or the conical part. In flight a transition between these two situations might occur, and this can result in undesirable effects on the aerodynamic stability. To understand the shock shape transition behaviour in more detail, a systematic numerical and experimental study is ongoing. In the present article, the influence of vibrational non-equilibrium compared to frozen flow is discussed for sphere-cone configurations at zero-degree angle of attack.

Within the context of supersonic combustion modeling in air-breathing propulsion, the turbulent Prandtl and turbulent Schmidt numbers have a profound effect on the resulting predictions. This work considers a Scalar Fluctuation Model (SFM) introducing two extra transport equations in order to specify variable values for turbulent Prandtl and turbulent Schmidt numbers within a Reynolds-Averaged Navier-Stokes (RANS) framework relying on a gradient diffusion hypothesis for both the turbulent heat and mass flux. The SFM is applied in conjunction with a two-equation linear eddy viscosity model. In order to study the effect of the different modeling choices, the supersonic reacting wall jet experiment of Burrows and Kurkov [24] is considered. A strong effect of Turbulence Chemistry Interaction (TCI) through an assumed PDF approach on the SFM is observed and needs to be enabled in order for predictions to be within experimental observation bounds in terms of ignition delay. Overall, the presented SFM methodology could be a viable candidate for supersonic combustion studies.

Kinetic Fokker–Planck (FP) methods for modeling rarefied gas flows have received increasing attention over the last few years. However, formulating such models for realistic multi-species gases is still an open subject of research. Therefore, in this letter, we develop a kinetic FP model for describing gas mixtures with particles interacting according to the variable hard-sphere interaction potential. In accordance with the kinetic FP framework, a stochastic solution algorithm is employed in order to solve the model on a particle level. Different test cases are carried out, and the performance of the proposed method is compared with the direct simulation Monte Carlo algorithm.

Since its first introduction, it has always been a subject of research to find models for a meaningful approximation of the highly accurate but complex Boltzmann equation. In the kinetic Fokker-Planck (FP) approach, a FP operator in velocity space is employed to approximate the collision integral of the Boltzmann equation. Instead of directly solving the resulting FP equation, a Monte Carlo technique is used to model an associated random process. This approach leads to an efficient stochastic solution algorithm. In recent years, the FP ansatz has become increasingly popular. Nevertheless, the modeling of gas mixtures in the context of kinetic FP has so far only been addressed in a very few papers. This article introduces a kinetic FP model that is capable of describing gas mixtures with particles interacting according to the hard-sphere collision model. The model is constructed to reproduce Grad's 13 moment equations on a Navier-Stokes level of accuracy for gas mixtures with an arbitrary number of constituents. A stochastic simulation algorithm is derived that ensures a correct evolution of the species diffusion velocities and the species temperatures for a homogeneous gas, regardless of the applied time step size. It is shown that the proposed model is capable of correctly predicting shear stresses, heat fluxes, and diffusion velocities for different test cases, employing a He-Ar mixture.

The High Enthalpy Shock Tunnel Göttingen (HEG) of the German Aerospace Center (DLR) is one of the major European hypersonic test facilities. It was commissioned for use in 1991 and was utilized since then extensively in a large number of national and international space and hypersonic flight projects. Originally, the facility was designed for the investigation of the influence of high temperature effects such as chemical and thermal relaxation on the aerothermodynamics of entry or re-entry space vehicles. Over the last years its range of operating conditions was subsequently extended. In this framework the main emphasis was to generate test conditions which allow investigating the flow past hypersonic flight configuration from low altitude Mach 6 up to Mach 10 in approximately 33 km altitude. The studies performed in HEG focused on the external as well as internal aerodynamics including combustion of hydrogen in supersonic combustion and the investigation of transition from laminar to turbulent hypersonic flow.

We present a methodology to measure the power delivered to sustain an inductive plasma discharge inside radio-frequency ion thrusters. This measurement is performed in real time near the coil within the resonant circuit using a patented low-noise radio-frequency voltage and current sensor system. With knowledge of the actual forwarded power, the efficiency of the generator under consideration, including the power feed cable, is assessed in real time. The variation of the power-to-thrust ratio can thus be understood more clearly since its value is given by physical changes of the plasma properties as well as the quality of the impedance bridge between the generator and plasma load. The latter can directly be determined by the proposed methodology, while the influence of the plasma parameters on the power-to-thrust ratio is obtained by means of a numerical model that is used in postprocessing to assess the power distribution within the thruster.

Experimental shock wave boundary layer interactions (SBLI) at a compression corner of a flat plate model with deflectable flap and variable surface temperature were studied in the High Enthalpy Shock Tunnel Gottingen (HEG) of the German Aerospace Center (DLR) at Mach 7.4 and \({6.65\times 10^6}\,{\text {m}^{-1}}\) unit Reynolds number. The present paper focuses on the effect of the leading edge bluntness and surface temperature. The effects are discussed based on surface heat flux and surface pressure measurements as well as schlieren visualizations. The study confirms that the state of the incoming boundary layer is crucial for the interaction. A small change of the leading edge bluntness was found to affect the boundary layer stability and thus the SBLI. In contrast, surface heating was found to destabilize the boundary layer. Tests at different flap deflection angles and transitional incoming boundary layers revealed that the transition process was completed in the separated shear layer, leading to a fully turbulent reattachment on the flap. The experimental effort is complemented by Reynolds-Averaged Navier–Stokes (RANS) computations providing insights into fundamental trends. The 2D simulations underpredict the size of the separation bubble in the compression corner but qualitatively mirrored the experimental trends. The wall heating resulted in an increased boundary layer thickness, reduced the extent of the separation bubble and elevated the peak boundary layer temperature on the flap.

Transonic flow in the wake of a generic space launch vehicle is investigated using Improved Delayed Detached-Eddy simulations on quasi-structured, partially unstructured and hybrid grids. It is shown that a \(90^\circ \) segment of the full domain can already capture the main afterbody flow dynamics and give a good agreement with experimental results. In order to assess the effect of unstructured grids, simulation results on a grid with an unstructured recirculation region are compared to a quasi-structured and a true hybrid grid. Finally, a comparison between IDDES and zonal DES results are presented and discussed.

Previous investigations in the High Enthalpy Shock Tunnel Göttingen of the German Aerospace Center (DLR) show that carbon fiber reinforced carbon ceramic (C/C) surfaces can be utilized to damp hypersonic boundary layer instabilities resulting in a delay of boundary layer transition onset. Linear stability analyses were performed using the DLR stability code NOLOT, NOnLocal Transition analysis code. To adapt the boundary condition to account for the characteristics of porous C/C material, the ultrasonic absorption properties of C/C were investigated experimentally and theoretically. Therefore, a test rig was set up to directly measure the reflection coefficient in the frequency and pressure range corresponding to the test conditions in HEG. In this frame, the reflection of ultrasonic waves from flat plate test samples with different porous layer thicknesses was investigated and compared to an ideally reflecting surface.The obtained results were used to improved the boundary condition used for stability analysis above porous surfaces. The numerical results, using the original as well as the improved boundary condition, were compared with wind tunnel tests. These experiments were performed at Mach 7.5 and different unit Reynolds numbers. A 7∘ half-angle cone model with a nose radius of 2.5mm and a total length of 1077mm was used. One-third of the metallic model surface in circumferential direction was replaced by C/C ceramics. The comparison between numeric and experiments includes the investigations of the second modes, the damping of the these modes and the resulting transition shift.

Experiments to measure the forces and moments acting on a blunted-cone capsule model are conducted in the High Enthalpy Shock Tunnel Gottingen. A free-flying arrangement is employed, whereby the mo...

For a spacecraft traveling at hypersonic speeds, the design of the vehicle’s thermal protection system is crucial [2] and the accurate quantification of the surface heat flux is mandatory to aid this design. Traditional surface-mounted temperature sensors, such as thermocouples and thin film gauges, have been widely used as reported already over 40 years ago by [19]. These sensors have response times on the order of a few microseconds, and thus data can be sampled with high temporal resolution; typically 1 MHz. However, the spatial resolution is less satisfactory: the packing density is limited by the sensor size and a great amount of expertise is needed to mount sensors on narrow geometries (e.g. sharp leading edges or corners). In addition, the installation of large numbers of transducers results in complicated wiring and high costs. The clear advantages of temperature sensitive paints (TSPs) over the classical techniques are the non-intrusive measurement approach and the high spatial resolution. A feasibility application of fast-response TSPs in Europe’s major hypersonic facility, the High Enthalpy Shock Tunnel Gottingen (HEG), on transition measurements has already been published [18]. These measurements demonstrated the ability of the TSP method to be used for measurements of time resolved surface temperatures and consequently the determination of surface heat fluxes; the TSP used in the publication was based on the luminophore Ru(phen). This paper will discuss the prerequisites, the requirements and the calibration work which has to be accomplished to progress the development of ultra-fast temperature sensitive paints towards a more versatile application in hypersonic high enthalpy flows in the HEG. Additional information on the HEG is reported by Hannemann and Martinez Schramm [7]. The calibration procedure is shown for the laser dye 4-Methylumbelliferone (4MU) which has been determined to be the next promising candidate for the application in the HEG [15].

Two phase ow is a challenging subject in the design of solid rocket motors. Not only does it impact the motor performance characteristics, but also inuences the thermal and chemical loads on the motor components. The main particulate found in hybrid and solid rocket combustion is made of aluminum. This study focuses on the continuous development of the TAU Lagrangian particle tracer, including eorts in modeling droplet evaporation, droplet breakup and condensation processes to describe the behavior of the solid particles within the combustion chamber. By combining computationally aordable models, an engineering model able to predict thermal, chemical and mechanical loads is presented. Application to a solid rocket motor engine indicates that this particular modeling approach underpredicts the combustion chamber temperature. Implementation of nucleation and subsequent condensation as a gas phase reaction resulted in chamber temperatures close to the temperature found from equilibrium combustion. While the temperature and species distribution within the chamber largely depend on the modeling of the combustion process, it is shown that the droplet breakup model has a much stronger inuence on the accurate prediction of droplet diameters in the nozzle and subsequent plume.

The numerical prediction of the processes in rocket combustion chambers is still challenging due to a variety of problems that are not completly investigated nor completly understood within computational fluid dynamics (CFD): First to mention is the accurate prediction of the jet behavior for sub- and supercritical conditions of propellents (e.\,g. liquid oxygen LOx) that influences the whole combustion process as well as the flame structure. Another challenge for the numerical method are the dissimilar length scales and material properties for the liquid and gaseous parts in combination with the chemical reactions for the LOx-GH2 combustion. In this study we use the TAU numerical flow solver for the simulation of single-injector experiments conducted at the German Aerospace Center (DLR) in Lampoldshausen. Based on experimental investigations in a windowed DLR subscale thrust chamber ``C'' (designated BKC) the numerical method in TAU for the simulation of sub- und supercritical LOx-GH2 combustion and atomization model is validated. Liquid oxygen and gaseous hydrogen have been injected through a single coaxial nozzle injector element. For validation of the numerical codes extensive measurements of OH emissions and shadowgraph images of the jet structure are available. Furthermore, wall pressure and temperature measurement data are available that are useful for the validation of the numerical framework. With this study we aim to validate an extension to the TAU flow solver that handles efficiently cryogenic fluids. This solver is able to predict the atomization and combustion processes within rocket combustion chambers at typical operating conditions. For validation we compare the numerical results for one super- and one subcritical load step measured in the experimental BKC campaign.

The free flight force measurement technique is a very attractive tool to determine forces and moments in particular in short duration ground based test facilities. With test times in the order of a few milliseconds, conventional force balances cannot be applied. The free flight technique has been applied in a number of shock tunnels utilizing models up to approximately 300 mm in length and looking at external aerodynamics. Test campaigns in the High Enthalpy Shock Tunnel Gottingen (HEG) of the German Aerospace Center (DLR) conducted in the framework of the EU co-funded LAPCAT II project demonstrated that this technique, in combination with optical tracking, can successfully be applied for thrust measurements of a complex fully integrated scramjet engine. In addition to the engine components intake, combustor and thrust nozzle, the hydrogen fueling system and gauge measurement infrastructure were integrated in the 1.5 m long model with a mass of 85 kg. The test campaigns in HEG together with vehicle aerodynamics predictions based on computational fluid dynamics demonstrated experimentally that the considered small scale flight experiment configuration is able to generate net thrust, i.e., has a positive aero propulsive balance. In the present article, the results of a new HEG test campaign utilizing an optimized optical set-up are presented. Motivated by uncertainty analyses results, the optical resolution and the camera frame rate were increased. It could be shown that by these modifications the uncertainty of the thrust increment measurements could be reduced by a factor of approximately 2. Further, the new results demonstrated the repeatability of the scramjet thrust measurement technique in HEG.

The High-Enthalpy Shock Tunnel Gottingen, HEG, is a free piston-driven shock tunnel, which was developed and constructed from 1989 to 1991. In a free piston-driven shock tunnel, the conventional driver of a shock tunnel is replaced by a free piston driver. This concept was proposed by Prof. Ray Stalker and, therefore, the tunnels are also referred to as Stalker tubes. HEG was commissioned for use in 1991, at that time being the largest facility of its type worldwide. Since then it was extensively used in a large number of national and international space and hypersonic flight projects. The research activities which were always strongly linked with CFD investigations range from the calibration process of the facility and the study of basic aerodynamic configurations, which are well suited to investigate fundamental aspects of high-enthalpy flows, to the investigation of complex entry, reentry, hypersonic flight, and integrated scramjet configurations. In the present article, an overview of selected research projects conducted in HEG is given.

Generally, the shape of streamline-traced intakes is defined by stream surfaces based upon an inviscid template flowfield but corrected by appropriate local boundary-layer displacement thicknesses....

An experimental and numerical investigation of the unsteady phenomena induced in a hydrogen-fuelled scramjet combustor under high-equivalence-ratio conditions is carried out, focusing on the processes leading up to unstart. The configuration for the study is the fuelled flow path of the HyShot II flight experiment. Experiments are performed in the HEG reflected-shock wind tunnel, and results are compared with those obtained from unsteady numerical simulations. High-speed schlieren and OH∗ chemiluminescence visualization, together with time-resolved surface pressure measurements, allow links to be drawn between the experimentally observed flow and combustion features. The transient flow structures signalling the onset of unstart are observed to take the form of an upstream-propagating shock train. Both the speed of propagation and the downstream location at which the shock train originates depend strongly on the equivalence ratio. The physical nature of the incipient shock system, however, appears to be similar for different equivalence ratios. Both experiments and computations indicate that the primary mechanism responsible for the transient behaviour is thermal choking, though localized boundary-layer separation is observed to accompany the shock system as it moves upstream. In the numerical simulations, the global choking behaviour is dictated by the limited region of maximum heat release around the shear layer between the injected hydrogen and the incoming air flow. This leads to the idea of ‘local’ thermal choking and results in a lower choking limit than is predicted by a simple integral analysis. Such localized choking makes it possible for new quasi-steady flow topologies to arise, and these are observed in both experiments and simulation. Finally, a quasi-unsteady one-dimensional analytical model is proposed to explain elements of the shock-propagation behaviour.

Detailed understanding of the physical processes occurring in the combustor of a scramjet engine is crucial for enabling this technology and is considered the most promising for hypersonic flight. Laboratory, ground testing, and flight-testing experiments, together with simulations, comprise the tools available to fill the current gap in scramjet combustor knowledge. Here, a computational study has been carried out for the HyShot II scramjet combustor using Large Eddy Simulation (LES) models together with one skeletal and one comprehensive reaction mechanism. Based on a survey of the experimental data available for the HyShot II combustor, we focus on the High Enthalpy Shock Tunnel Gottingen (HEG) operating condition XIII, emulating flight conditions at 28 km altitude. To account for slight experimental run-to-run variations, two simulations are performed at different equivalence ratios. The LES are found to capture the experimental wall-pressure and heat-flux data well compared to the measurement data, with marginal influence of the reaction mechanism. The LES results are subsequently used to analyze the flow, mixing, and combustion processes involved. The tools employed include conventional, as well as recently developed methods, such as the Takeno flame index and Chemical Explosive Mode Analysis. These methods give a concise description of the interactions between flow, fuel–air mixing, and combustion, and it is discovered that supersonic combustion is a combination of auto-ignition and non-premixed flame regions and self-igniting fronts. Furthermore, ignition is enabled by shocks, and the supersonic flame is very different in nature to subsonic turbulent flames.

Experimental investigations of the free-stream disturbance spectrum by means of a slender wedge probe were conducted in the DLR High Enthalpy Shock Tunnel Gottingen (HEG) at Mach 7.4, the DNW Ludwieg tube (RWG) at Mach 3 and Mach 6 as well as in the TU Braunschweig Ludwieg tube (HLB) at Mach 6. The identical measurement chain, including the probe, the amplifier and the type of transducer were used in all facilities. A wide range of fast response pressure, temperature and heat flux transducers were used to measure the corresponding fluctuations on the probe surface behind the oblique shock. Quantitative results were obtained as function of frequency taking into account the limitations imposed by each transducer. The wedge probe was shown to be a robust and easy to implement probe particularly suited for use in harsh test environments which do not allow the application of hot wire techniques or flush mounted transducers in a Pitot probe configuration.

Visualization experiments are performed to investigate the development of instability waves within the boundary layer on a slender cone under high Mach number conditions. The experimental facility is a reflected-shock wind tunnel, allowing both low (Mach-8 flight equivalent) and high-enthalpy conditions to be simulated. Second-mode instability waves are visualized using a high-speed schlieren set-up, with pulse bursting of the light source allowing the propagation speed of the wavepackets to be unambiguously resolved. This, in combination with wavelength information derived from the images, enables the calculation of the disturbance frequencies. At the lower-enthalpy conditions, we concentrate on the late laminar and transitional regions of the flow. General characteristics are revealed through time-resolved and ensemble-averaged spectra on both smooth and porous ceramic surfaces of the cone. Analysis of the development of individual wavepackets is then performed. It is found that the wavepacket structures evolve from a ‘rope-like’ appearance to become more interwoven as the disturbance nears breakdown. The wall-normal disturbance distributions of both the fundamental and first harmonic, which initially have local maxima at the wall and near $y/{\it\delta}=0.7$–0.75, exhibit an increase in signal energy close to the boundary-layer edge during this evolution. The structure angle of the disturbances also undergoes subtle changes as the wavepacket develops prior to breakdown. Experiments are also performed at high-enthalpy ($h_{0}\approx 12~\text{MJ}~\text{kg}^{-1}$) conditions in the laminar regime, and the visualization technique is shown to be capable of resolving wavepacket propagation speeds and frequencies at such conditions. The visualizations reveal a somewhat different wall-normal distribution to the low-enthalpy case, with the disturbance energy concentrated much more towards the wall. This is attributed to the highly cooled nature of the wall at high enthalpy.

This paper introduces a new model for real gas thermodynamics, with improved accuracy, performance, and robustness compared to state-of-the-art models. It is motivated by the physical insight that in non-premixed flames, as encountered in high pressure liquid propellant rocket engines, mixing takes place chiefly in the hot reaction zone among ideal gases. We developed a new model taking advantage of this: When real fluid behavior only occurs in the cryogenic oxygen stream, this is the only place where a real gas equation of state (EOS) is required. All other species and the thermodynamic mixing can be treated as ideal. Real fluid properties of oxygen are stored in a library; the evaluation of the EOS is moved to a preprocessing step. Thus decoupling the EOS from the runtime performance, the method allows the application of accurate high quality EOS or tabulated data without runtime penalty. It provides fast and robust iteration even near the critical point and in the multiphase coexistence region. The model has been validated and successfully applied to the computation of 0D phase change with heat addition, and a supercritical reactive coaxial LOX/GH 2 single injector.

Two identical, interacting plumes emanating from model thrusters with parallel axes separated from 50 to 150 throat diameters are studied numerically. The nozzle throat Reynolds number is set to nearly 15, 000 to match that of a small bi-propellant attitude control thruster, but the simulated gas is nitrogen with a stagnation temperature of 300 K. The near-isentropic, dense plume core is computed with the DLR Navier-Stokes solver TAU and the conditions at a suitably defined interface are then used on the inflow boundary of a separately conducted direct simulation Monte Carlo (DSMC) simulation. The results are shown to agree favorably with particle flux measurements performed in the DLR high-vacuum plume test facility for chemical thrusters (STG-CT). Varying the nozzle separation distance alters the degree of rarefaction in the interaction plane, and by tagging DSMC particles according to their origin, the effect on the individual plume may be investigated. The impact of nozzle axis separation on mass flux...

The paper summarizes technical results and first highlights of the cooperation between DLR and Airbus Defence and Space (DS) within the work package “CFD Modeling of Combustion Chamber Processes” conducted in the frame of the Propulsion 2020 Project. Within the addressed work package, DLR Gottingen and Airbus DS Ottobrunn have identified several test cases where adequate test data are available and which can be used for proper validation of the computational fluid dynamics (CFD) tools. In this paper, the first test case, the Penn State chamber (RCM1), is discussed. Presenting the simulation results from three different tools, it is shown that the test case can be computed properly with steady-state Reynolds-averaged Navier-Stokes (RANS) approaches. The achieved simulation results reproduce the measured wall heat flux as an important validation parameter very well but also reveal some inconsistencies in the test data which are addressed in this paper.

At hypersonic flight the kinetic energy of the flow is large enough that high temperature effects such as vibrational excitation or dissociation of the fluid molecules occur. Due to the extremely high power requirement and the severe flow environment, the required test conditions can only be achieved in ground based impulse facilities. The most successful types of facility which are able to generate high enthalpy and high pressure hypersonic flows are shock tunnels and shock expansion tunnels. The principle of operation of these facilities is to store the energy over a long period of time, therefore reducing the necessary power requirement and subsequently releasing the stored energy rapidly. In free piston driven shock tunnels, the conventional driver of a shock tunnel is replaced by a free piston driver. This concept was proposed by Prof. R.J. Stalker in the 1960th and the facilities are referred to as Stalker tubes. In the present article, four major Stalker tubes, the High Enthalpy Shock Tunnel Gottingen, HEG, at the German Aerospace Center, the High Enthalpy Shock Tunnel, HIEST, at the Japan Aerospace Exploration Agency, Kakuda, T4 at The University of Queensland, Brisbane, Australia and T5 at the Graduate Aeronautical Laboratories, California Institute of Technology, United States are presented. In addition to facility overviews, selected research activities are discussed.

The applicability of the particle image velocimetry (PIV) and the background-oriented schlieren (BOS) techniques in the high-enthalpy shock tunnel Gottingen of the German Aerospace Center, DLR is demonstrated. As a part of this feasibility study two different experiments are performed. The velocity field past a wedge in a Mach 6 flow at a total specific enthalpy of 1.5 MJ/kg is determined by means of PIV and the results are compared to numerical predictions. The BOS technique is applied to investigate the density field in the shock layer of a sphere at 12 and 22 MJ/kg total specific enthalpies. Using a ray tracer method, the BOS results are compared to the data obtained by corresponding numerical computations.

A stress wave internal force balance for the High Enthalpy Shock Tunnel Gottingen (HEG) of the German Aerospace Center (DLR) to measure lift, pitching moment and drag was designed, calibrated and tested. The balance is designed to measure forces in ground based test facilities with test times in the order of milliseconds on models additionally instrumented with surface pressure and wall heat flux gauges from angles of attack of −40° to 20°. Experiments in HEG were performed on a 303 mm long, 10° half angle blunt cone at angles of attack from −20° to 0°. The tests were conducted utilizing two different operating conditions at total specific enthalpies of 3.0 and 3.5 MJ/kg and dynamic pressures of 30 and 72 kPa. The performance of the balance was assessed by comparing the measured force and moment coefficients with computational fluid dynamics (CFD) predictions.