Tobias Langener

Experimental Investigations on Transpiration Cooling for Scramjet Applications Using Different Coolants

The extremely high heat loads within a scramjet combustor require the use of high-temperature materials combined with efficient cooling concepts. A promising technique is the application of transpiration cooling to ceramic matrix composite materials. A supersonic hot-gas-flow test facility is used to investigate this cooling method. The carbon/carbon samples tested have porosities of about e = 11%. The airflow is electrically heated up to 1120 K total temperature with a total pressure of ≈3 bar and is accelerated to a Mach number of 2.1 within the test channel. Air, argon, and helium are used as coolants for blowing ratios from 0 to 1 %. The surface temperature of the porous wall is measured via thermocouples and infrared thermography. Pressure and mass-flow measurements are used to analyze the throughflow characteristics of the porous materials at various temperature conditions. An approach based upon simplified analytical models is presented to analyze the experimental data of throughflow behavior and cooling efficiency. The simplified thermal model is used to analyze the effect of fluid property variations with temperature on pressure loss for different coolants and shows good agreement with the experimental data.

Experimental Investigation of Cooling Techniques and Materials for Highspeed Flight Propulsion Systems

The multidisciplinary design process of future supersonic and hypersonic flight vehicles implies the development of propulsion systems using sophisticated cooling techniques in combination with advanced materials. Within the framework of the European research project ‘Aerodynamic and Thermal Load Interactions with Lightweight Advanced Materials for High Speed Flight’, in short ATLLAS, different cooling techniques using both metallic and ceramic materials for propulsion systems are under investigation. The hot fire experiments deal with the application of film and transpiration cooling in operating conditions beyond the scope of conventional aeroengines. Different newly developed ceramic materials are studied with respect to their applicability in oxidizer rich as well as fuel rich combustion atmospheres and operating conditions typical for the high-pressure turbojet and ram-based lower pressure engines. The information gained in the different test programs is fed back to the project partners engaged in the development of design and simulation tools and used as an input for the MDO design process of the overall propulsion system.

Aerodynamic Characterization of HEXAFLY Scramjet Propelled Hypersonic Vehicle

This paper deals with the design trade-off activities undertaken to provide a trim-able, a statically and dynamically stable vehicle configuration able to perform a nominal experimental scramjet-propelled flight. The flight control activities and their impacts on vehicle layout and global aerodynamic performance are also addressed. In particular, different competing aeroshapes have been investigated to assess the best one camplyant with project requirements. In this framework, trade-off results in terms of: setup and/or analysis of aero-propulsive databases; design loops for the aileron (shape, span, length); design loops for the vertical tail (shape, size, toe-angle); analysis of aerodynamic performances; analysis of longitudinal trimming conditions; sensitivity to centre of gravity position of static longitudinal stability and trimming conditions; static stability analysis (for longitudinal and lateral-directional flight) in clean and flapped configuration, static margins (pitch, roll, yaw); dynamic stability analysis with a focus on roll-yaw coupling (linearized model analysis and Dutch-Roll period evaluation); characterization of hinge-moments; and preliminary selection of flight control equipment are provided and described in detail in the paper.

Transpiration Cooling with Supersonic Flows and Foreign Gas Injection

The extremely high heat loads within a Scramjet combustor require the use of high-temperature materials combined with efficient cooling concepts. A promising technique is the application of transpiration cooling to CMC materials. A supersonic hot-gas flow test facility is used to investigate this cooling method. The Carbon/Carbon (C/C) samples tested have porosities of about 11 %. The air-flow is electrically heated up to 1060 K total temperature with a total pressure of 3 bar and is accelerated to a Mach number of 2.1 within the test channel. Air, argon and helium are used as coolants for blowing ratios from 0 to 1 %. The surface temperature of the porous wall is measured via thermocouples and infrared thermography. An approach based upon simplified analytical models is being presented to analyze the experimental data of the cooling efficiency. This is shown for elevated temperature conditions in supersonic flows, including the specific thermal situations, which are typically encountered in such high-temperature tests.

Overview of a Supersonic Probe for Solid Propellant Rocket CCP Collection

The present work aims at giving a comprehensive overview of the current development status of an intrusive probe, capable of collecting the condensed combustion products present in the exhaust of a solid rocket motors. The innovative technique was conceived in the EMAP (Experimental Modelling of Alumina Particulate in Solid Booster) framework, a project aiming at characterizing the alumina in terms of size, temperature and spatial distribution to gain detailed information for climatological impact assessment. A supersonic probe was sized to handle a progressive deceleration and cooling of the exhaust gas, as well as the quenching and collection of the suspended particles in a pressure-controlled chamber. The task was achieved by through a quasi 1D gas dynamics code based on the Shapiro method and the normal shock wave theory, which was verified against a hybrid 2D axial-symmetric mesh whose turbulent flow field was solved using the DLR-TAU CFD code. The robustness of the system has been investigated performing a sensitivity and an uncertainty analysis, exploring uncertainties propagation through the numerical code based on Shapiro equations. The sensitivity analysis enabled to define a ranking of importance for the uncertainties on the probe behavior; the uncertainty analysis allowed to estimate failures of the system and/or of the code. Cold flow tests carried out at the vertical facility of DLR-Cologne enabled to gain a proof of concept for both fluid dynamic behavior and collection methodology.

Development of a Probe for Particle Collection in High-Temperature, Supersonic Flow: Conceptual and Detailed Design

An intrusive technique for the collection of the condensed combustion products in the proximity of the rocket nozzle is proposed. In particular, a supersonic probe able to withstand the harsh environment of a plume was sized to handle a progressive deceleration and cool down of the exhaust gas, preventing from liquid particle breakup. The task was achieved by diluting the swallowed flow with a cold inert gas and quenching the suspended particles using a liquid spray in a specific chamber. Preliminary tests performed in the supersonic wind tunnel at DLR confirmed the quality of the collection device.

Inlet Aerodynamics and Ram Drag of Laser-Propelled Lightcraft Vehicles

Numerical simulations are used to study the aerodynamic inlet properties of three axisymmetric configurations of laser-propelled Lightcraft vehicles operating at subsonic, transonic, and supersonic speeds up to Mach 5. The 60 cm vehicles were sized for launching 0.1-1.0 kg nanosatellites with combined-cycle airbreathing/rocket engines, transitioning between propulsion modes at roughly Mach 5-6. The selected external-compression inlet forebodies included the Mercury nose, a body with a power-law nose, and a body with a parabolic nose—all equipped with cylindrical shrouds. The simulations utilize the Fluent Reynolds-Averaged-Navier-Stokes (RANS) flow solver coupled with the two-equation k - e model for near-wall turbulence. The numerical mesh is generated by the commercial grid generator Centaur 6.0. This package is able to create structured, unstructured, and hybrid meshes around complex geometries. Results provide the pressure, temperature, density, and velocity fields around the three representative configurations as well as the resulting ram drag and total drag coefficients—all as a function of flight Mach number. Given the three alternatives, it is demonstrated that the configurations with the power-law nose and with the parabola nose provide near-optimal geometries for minimum drag during the whole flight trajectory.