Particle mass flow determination in dust laden supersonic flows by means of simultaneous application of optical measurement techniques

Particle mass flow rate and particle mass concentration are key parameters for describing two-phase flows, especially for particle-induced heating augmentation analysis. This work addresses the question of how accurate particle mass flow rate can be determined with three non-intrusive measurement approaches, based on shadowgraphy, particle tracking velocimetry (PTV), and scattered light intensity, in supersonic flows. In terms of shadowgraphy and PTV, the particle mass flow rate was determined by measuring individual particle characteristics, namely particle size, velocity, and density, as well as the measurement volume. The presented shadowgraphy procedure is based on the commercial LaVision DaVis software and additional shadowgraphy corrections. Multiple tests were conducted in the experimental test facility GBK of DLR with varying flow conditions, at a Mach number of 2.1, unit Reynolds number (Re∞) ranging from 5e7 1/m to 1.5e8 1/m, total temperature (T0) ranging from 303 to 544 K, and particle materials, namely Al2O3, MgO, and SiO2, in the size range of 1 to 60 µm. Particle size distributions of Al2O3 and MgO particles could be reproduced with shadowgraphy quite well, while the PTV procedure resulted in non-similar distributions. Pycnometer measurements indicated MgO particle density to be significantly lower than reference values. A DaVis parameter variation analysis resulted in a particle mass flow rate uncertainty of shadowgraphy of up to 30%. The particle mass flow rate uncertainty of PTV is approx. 76%, and the respective uncertainty of scaled PTV and scattered light intensity approach is 28%. The particle mass flow rate, measured with shadowgraphy, is 58% higher than those of the semi-axisymmetric scattered light intensity approach, which can be explained by a higher particle concentration at the injection plane

Experimental investigation of heating augmentation by particle kinetic energy conversion in dust laden supersonic flows

The presence of particles in supersonic flows can cause significant increases in stagnation point heat fluxes (Dunbar et al. in AIAA J 13:908–912, 1975). This effect is commonly named particle-induced heat flux augmentation or just heating augmentation. Heating augmentation can be described as the sum of the conversion of kinetic energy of the particles into thermal energy, characterized by the energy conversion efficiency, also called accommodation coefficient, and the increase of convective heat flux (Polezhaev et al. in High Temp 30:1147–1153, 1992; Vasilevskii and Osiptsov in Experimental and numerical study of heat transfer on a blunt body in dusty hypersonic flow 33rd thermophysics conference, American Institute of Aeronautics and Astronautics, 1999). Although the accommodation coefficient is fundamental for heating augmenta- tion characterization, there is only a small number of experimental datasets for it. This work focusses on the experimental determination of the accommodation coefficient in flow regimes at Mach number 2.1, Reynolds number, based on the probe nose diameter, from approx. 6e5 to 1.8e6, and nominal particle sizes of approx. 20 um. The decrease of particle velocity and kinetic energy flux in the shock layer is measured with highly resolved shadowgraphy for individual particles. The particle kinetic energy flux is decreased by 29% on average by particle deceleration in the shock layer. Negligible kinetic energy fluxes of rebounded particles were measured. The accommodation coefficient is approx. 0.36 for Al 2 O 3 and SiO 2 particles, while it is approx. 0.09 for MgO particles. Hence, it is significantly smaller than the widely used value of 0.7, based on the study of (Fleener and Watson in Convective heating in dust-laden hypersonic flows 8th thermophysics conference, 1973), but in good agreement with values given in (Hove and Shih in Reentry vehicle stagnation region heat transfer in particle environments 15th aerospace sciences meeting, 1977) and (Molleson and Stasenko in High Temp 55:87–94, 2017. https:// doi.org/10.1134/S0018151X1701014X ). No difference between erosive and elastic particle reflection mode was detected on the conversion efficiency. The data from a simplification of the modeling approach of the conversion efficiency for elastic particle reflection by Molleson and Stasenko (2017) are in poor agreement with experimental data

Shock interaction induced heat flux augmentation in hypersonic flows

This paper gives a summary of dedicated experiments on the shock interaction induced heat flux augmentation, by means of tests carried out in the hypersonic wind tunnel H2K. The first test case is devoted to the shock boundary layer interaction on a flat plate. The interaction impact has been varied by changing the free stream parameters and the position of the shock generator, i.e. shock impingement point on the plate. The heat flux distribution has been determined using surface temperature data measured by an infrared camera. The heat flux data combined with free stream flow parameters allow calculation of the Stanton number evolution. The second test case is a double sphere configuration with a variable axial and lateral distance between the spheres. This allowed measurements of the heat flux augmentation induced by a shock-shock teraction along the complete frontal surface of the second sphere, which was hit by the bow shock of the first sphere. Shock-shock and shock-boundary layer interaction effects are studied by means of experiments on the IXV flight konfiguration with double control flaps. Depending on the test configuration and flow parameters, shock interaction induced heat flux augmentation factors up to seven have been measured

Simultaneous determination of particle size, velocity, and mass flow in dust‑laden supersonic flows

The particle mass concentration and -mass flow rate are fundamental parameters for describing two-phase flows and are products of particle number, -size, -velocity, and -density. When investigating particle-induced heating augmentation, a detailed knowledge of these parameters is essential. In most of previous experimental studies considering particle-induced heating augmentation, only average particle mass flow rates are given, without any relation to measured particle sizes and -velocities within the flow or any indication of measurement uncertainty. In this work, particle number, individual particle sizes, and velocities were measured in a supersonic flow by means of shadowgraphy and particle tracking velocimetry (PTV). The goals are to determine measurement uncertainties, a particle velocity-size relation, and the spatial distribution of number, size, velocity, and mass flow rate across the nozzle exit. Experiments were conducted in a facility with a nozzle exit diameter of 30 mm, at Ma_inf = 2.1 and Re_inf = 8.2e7 1/m. Particles made of Al2O3 and up to 60 µm in size were used for seeding. Particle mass flow rates up to 50 kg/m2 s were achieved. It is shown that an additional correction procedure reduced common software uncertainties regarding shadowgraphy particle size determination from 14% to less than 6%. Discrepancies between calculated particle velocities and experimental data were found. In terms of spatial distribution, larger particles and a higher mass flow rate concentrate in the flow center. The determined particle mass flow rate uncertainty was up to 50% for PTV; for shadowgraphy, it was less than 17%

System Analyses Driving Improved Aerothermodynamic Lay-out of the SpaceLiner Configuration

The revolutionary ultrafast passenger transport SpaceLiner is under investigation at DLR since 2005. The two-stage, fully reusable vehicle is powered by rocket engines. The maximum achieved velocity, depending on the configuration or mission type, is beyond 6.5 km/s putting some challenging aerothermal requirements on the vehicle. At the lower end of the speed-range, the SpaceLiner should have the smallest possible flight velocity for landing with an acceptable angle of attack. The focus of the paper is on all system aspects including the SpaceLiner’s flight performance which have an impact on the aerodynamic configuration. A preliminary sizing of both stages’s passive TPS is done. At the vehicle’s leading edges heat fluxes and hence equilibrium temperatures temporarily reach excessive values requiring advanced active transpiration cooling. An experimental campaign is run at the DLR arc-heated facility to increase the TRL of this promising cooling technology. An aerodynamic shape optimization taking into account trim drag aspects and latest status of the vehicle design and flight profile is described

EXPERIMENTAL FLOW CHARACTERIZATION AND HEAT FLUX AUGMENTATION ANALYSIS OF A HYPERSONIC TURBULENT BOUNDARY LAYER ALONG A ROUGH SURFACE

Surface roughness increases skin friction drag and convective heat transfer along high-speed flight vehicles. Although the corresponding heat flux augmentation is usually lower compared to increased friction, careful consideration in the prediction of the resulting heat loads is required to define suitable margins in the design of thermal protection systems. In the present study, the response of a hypersonic turbulent boundary layer to a smooth and rough surface along a sharp right-circular cone is examined. Tests were conducted at an inflow of Ma = 6 and Re = 16 Million per meter in the hypersonic wind tunnel H2K at DLR Cologne. The testing time was in the order of 20 seconds. The model consisted of three segments with exchangeable parts to consider smooth or rough surfaces. The roughness topology consisted of square bar elements to enable comparisons to previous experimental campaigns. The roughness-element wavelength was four times the depth of the elements. The model was made of a specific material with low thermal conductivity, in order to measure the surface temperature distribution by means of global quantitative infrared thermography and to avoid lateral heat dissipation. The flow field along the smooth and rough cone was measured in selected regions of interest by Particle Image Velocimetry (PIV). This technique was successfully applied for the first time in the high-speed environment of the H2K. The data is compared and discussed based on comparison to analytical and numerical predictions. The analytical calculations include classical turbulent smooth cone relations as well as correlations for rough surfaces. The data for numerical comparisons was derived by a boundary layer code and full CFD. In case of the boundary layer code a modified Krogstad model was applied to account for the rough wall

Experiments on a smooth wall hypersonic boundary layer at Mach 6

The turbulent boundary layer along the surface of high-speed vehicles drives shear stress and heat flux. Although essential to the vehicle design, the understanding of compressible turbulent boundary layers at high Mach numbers is limited due to the lack of available data. This is particularly true if the surface is rough, which is typically the case for all technical surfaces. To validate a methodological approach, as initial step, smooth wall experiments were performed. A hypersonic turbulent boundary layer at Ma = 6 (Ma_e = 5.4) along a 7° sharp cone model at low Reynolds numbers Re_theta = 3000 was characterized. The mean velocities in the boundary layer were acquired by means of Pitot pressure and particle image velocimetry (PIV) measurements. Furthermore, the PIV data were used to extract turbulent intensities along the profile. The mean velocities in the boundary layer agree with numerical data, independent of the measurement technique. Based on the profile data, three different approaches to extract the skin friction velocity were applied and show favorable comparison to literature and numerical data. The extracted values were used for inner and outer scaling of the van Driest transformed velocity profiles which are in good agreement to incompressible theoretical data. Morkovin scaled turbulent intensities show ambiguous results compared to literature data which may be influenced by inflow turbulence level, particle lag and other measurement uncertainties

MODELLING CAPSULE STABILITY ACCOUNTING FOR SHAPE CHANGE

Earth return missions or exploration missions use mostly capsule-like shapes, which enter the atmosphere at very high velocities. Some of these missions, like sample return, do not use any parachute or other stabilizing aerodynamic or RCS devices. Therefore, the capsule stability has to be guaranteed solely by the spacecraft configuration from hypersonic conditions down to the subsonic regime at landing. This task is very challenging and requires reliable design tools. However, both experimental and numerical tools still have shortcomings in full simulation or modelling of the flight environment. Therefore, further improvement of these tools by means of complementary application is essential. Most of the exploration missions use an ablative thermal protection system, which experiences shape changes during the hypersonic flight regime. This may lead to a change of the pressure distribution and movement of the center of gravity of the vehicle. Since the vehicle does not have control devices, it can lose its aerodynamic stability and the situation may become critical. The prediction of the Thermal Protection System (TPS) recession over the complete surface with the existing tools is not possible. Therefore, the aerodynamic design should consider it in the margin policy and the flight qualities and risk analysis need to be performed accordingly. The ESA TRP MODSHAPE (Modelling Capsule Stability accounting for Shape Change) addresses the aforementioned challenges and this paper gives an overview of the planned activities and summarizes the main challenges and goals