Herbert Olivier

Ignition of shock-heated H2-air-steam mixtures

The self-ignition characteristics of H 2 -air-steam mixtures have been investigated in a heated shock tube for various temperatures, pressures and gas compositions behind the reflected shock wave. The measured ignition delay times show a strong dependence on the steam concentration and the temperature, while there is only a weak influence of the pressure for the conditions studied. The measured ignition delay times are consistent with the theoretical prediction only in the high temperature region. In the low temperature region the measured delay times are clearly shorter than those determined by the theoretical prediction. The pressure increase caused by ignition processes depends on the temperature behind the reflected shock wave. With increasing temperature the peak pressure first increases, then decreases. In terms of pressure signals the temperature can be divided into three regions: strong, transitional, and mild ignition regions. In the transitional region, a secondary explosion generates a high overpressure somewhere between the reflected shock wave and the end wall.

On the Chemical Kinetics of Ethanol Oxidation: Shock Tube, Rapid Compression Machine and Detailed Modeling Study

Abstract Auto-ignition characteristics of ethanol were experimentally investigated using two Shock Tube (ST) facilities and a Rapid Compression Machine (RCM). Ignition delay times for stoichiometric ethanol-air mixtures were measured for temperatures between 775–1300 K in a High Pressure Shock Tube (HPST) at pressures close to 80 bar by probing pressure time histories and CH* emission. In some experiments the HPST was additionally employed for schlieren imaging to visualize ignition behavior by probing density gradients during ignition for ethanol-air mixtures. The ignition delay experiments in HPST were complemented by RCM measurements for extending the temperature regime to the Low Temperature Combustion (LTC) regime, down to 705 K, providing kinetic model validation data over a very wide temperature and pressure range. The current results also extend the earlier shock tube measurements performed in the same laboratory for pressures around 40 bar for temperatures down to 800 K [Heufer et al., Shock Waves 20 (2010) 307]. Furthermore, a Rectangular Shock Tube (RST) was solely used for additional schlieren imaging experiments to acquire information on ignition modes in stoichiometric ethanol-air mixtures around 10 bar. An improved chemical kinetic model was developed based on the Li et al. mechanism [Li et al., “Ethanol Model v1.0”, Princeton University, 2009] which was updated with evaluated rate parameters from the literature and validated through results obtained from the aforementioned facilities. The model predictions were compared to previously published low-pressure, premixed flat flame molecular beam mass spectrometry speciation data [Kasper et al., Combust. Flame 150 (2007) 220; Wang et al., J. Phys. Chem. A 112 (2008) 9255] where reasonable agreement is obtained considering the uncertainties in experiments and model. However, the model provides excellent agreement for the auto-ignition results obtained in the RCM and the high temperature shock induced ignition delays. Significant disparities with the model predictions are obtained for the shock tube results at temperatures below 1000 K as it transitions from the intermediate to the low temperature regime. The reasons for these deviations are assigned to strong fuel specific “pre-ignition” effects observed in ethanol auto-ignition, in contrast to other investigated fuels, which was satisfactorily explained through schlieren experimental results. To our knowledge this work is first of its kind that combines results from complementary experimental methods from three different facilities providing a holistic description on the auto-ignition behavior of ethanol. Furthermore, this paper reports ignition delay measurements for ethanol in air, at the highest pressures applicable to practical combustors.

Influence of Cooling-Gas Properties on Film-Cooling Effectiveness in Supersonic Flow

Film cooling is considered a prerequisite for the safe operation of future high-performance rocket engines. Wall-normal or inclined cooling-gas injection into a laminar boundary-layer airflow at Ma...

Tube expansion by gas detonation

The expansion of tubes by direct application of gas detonation waves is an alternative forming method for hollow section workpieces. In particular the process can be used for typical hydroforming parts, for example car body or exhaust elements in automotive industry. The gas charge of oxygen and hydrogen is both pressure medium and energy source and has the potential to cause high forming velocities. The introduced process belongs to the category of high speed forming methods and provides typical advantages such as higher achievable strains compared to quasi-static methods using high water pressure. Another advantage of this process is the avoidance of high press forces by application of an “inertia-locked tool” system due to the extremely short process time. To develop a controllable process, good knowledge of the interdependencies in the system “medium, workpiece and tool” is essential. This can be achieved using simulations in combination with experimental investigations. The results are topic of this paper, also including special investigations on the material behavior at high strain rates and temperature gradients.

Film Cooling of an Inclined Flat Plane in Hypersonic Flow

*† In this work a numerical and experimental investigation of the basic behavior of film cooling at hypersonic flow conditions is presented. For the experiments a ramp model with a transverse blowing slot in respect to the incoming flow is employed. Thereby the thermal loads are measured with thermocouples and an infrared camera. Additionally Schlieren pictures and pressure measurements were made. The experimental results are compared to numerical calculations and both show a high effectiveness of the film cooling technique at hypersonic flow conditions. Furthermore velocity and temperature profiles show the effect of the cooling gas injection to the boundary layer. Finally an estimation is made for cooling a body under real flight conditions.

Initiation and suppression of explosive processes in hydrogen-containing mixtures by means of permeable barriers

Experimental data on the interaction of detonation and shock waves with permeable barriers in cylindrical channels in hydrogen—air mixtures are presented. The method of soot tracks on semi-cylindrical smoked inserts made it possible to elucidate the mechanisms of initiation of the detonation downstream of the barriers. The possibility of initiating explosive processes, including detonation wave decay, is demonstrated. The numerical simulation of the detonation initiation behind the permeable barrier represented by a perforated screen is performed. Analysis of the experimental and calculation results demonstrated that in the near zone behind the barrier, detonation is initiated upon collision (focusing) of hemispherical shock waves emerging from separate holes. In the far zone, the initiation mechanism is similar to that observed on deflagration-to-detonation transition.

Aero-structural Dynamics Experiments at High Reynolds Numbers

The elastic wing model, its excitation and comprehensive high frequency measuring equipment for the High Reynolds Number Aero-Structural Dynamics (HIRENASD) tests in the European Transonic Windtunnel (ETW) are shortly described. Some of the stationary polars are presented in terms of wing deformation, as well as aerodynamic coefficients and pressure distributions. Then unsteady processes observed in the measurements of static aerodynamic coefficients, are regarded with focus on small amplitude pressure waves travelling upstream from the trailing edge and triggering periodically break-down and redeployment of the local supersonic domains with transonic shock waves to run upstream and to disappear. Another focus is on stochastic vibrations excitation while moving forward during nominally static experiments. Emphasis is put on measured variations of pressure distribution on the wing surface caused by defined vibration excitation applying internal force couples at the wing root, whereby the exciter frequencies were chosen close to natural frequencies of the wing model. Phase and magnitude of measured local lift fluctuations as well as real and imaginary parts of pressure distributions are presented.

Investigation of Unsteady Transonic Airfoil Flow

Numerical and experimental investigations of the flow over a supercritical airfoil are performed at the Shock Wave Laboratory. The transonic airfoil flow contains complex structures such as local supersonic regions and local separation regions, shock-boundary layer interaction, vortex-wake interaction and boundary layer transition. Furthermore, a phenomenon of upstream moving waves is observed in the experiments as well as in the numerical simulations. This paper focus on the numerical investigation of the upstream moving wave phenomenon in the two-dimensional unsteady transonic airfoil flow. The influence of the Reynolds number, Mach number and the incidence on the mean and unsteady flow properties is analyzed. A direct eect of the Mach number and incidence is observed on the wave phenomenon, whereas the Reynolds number influences the time averaged flow field and hereby indirect the upstream moving waves.

Use of shock tunnels for hypersonic propulsion testing

Shock tunnels represent a powerful tool for simulating hypersonic flow conditions as they occur during hypersonic flight and reentry. For propulsion testing it has been demonstrated that they are capable not only to allow the outer and intake flow field of complex propulsion systems but also important phenomena inside of the engines like fuel mixing and supersonic combustion. In the Aachen shock tunnel TH2 experiments have been performed with simple and complex ramp geometries to study typical intake flow phenomena. For this a high Mach number, high Reynolds number flow condition is used. But for hypersonic flight vehicles a simulation in wind tunnels as accurate and realistic as possible requires to simulate the high wall temperatures occuring in flight as close as possible. Therefore, in a first attempt a 24 degree ramp model was heated from the inner side by electrical resistance heating elements. The maximum temperature achieved so far amounts to 690 K. In order to increase the total enthalpy of the flow, a detonation driver has been developed which in near future will replace the conventional helium driver of the shock tunnel TH2. Up to now experiments have been performed with the detonation driver in the shock tube mode with a short 6 m long driven section. Results achieved with this facility (THD) show the feasibility of the detonation driver concept for shock tunnel applications over a wide range of flow conditions. THE AACHEN SHOCK TUNNEL TH2 The shock tube of the Aachen shock tunnel has an inner diameter of 140 mm with a wall thickness of 80 mm. The lengths of the driver and driven section are 6 m and 15.4 m, respectively. The building which *Professor, Member AIAA tResearch Engineer Coypright 01999 by the American Institute of Aeronautics and Astronautics Inc. All rights reserved. houses the shock tunnel was especially built for the use of such tunnels. A 800 mm steel-enforced concrete wall which separates the rooms for driver and driven section serves as a protecting wall but is also used for supporting the recoil absorbing system of the tunnel. Driver and driven section are separated by a double-diaphragm chamber which at maximum pressure utilizes two 10 mm thick stainless steel plates as diaphragms scored in the form of a cross by a milling cutter. Another diaphragm of brass or copper sheet is located between the driven section and the nozzle entrance. The maximum operating (steady) pressure of the complete tube is 1500 bar. The driver can electrically be heated to a maximum T4 of 600 K. There are two conical nozzles available, one with a half opening angle of 5.8 degree with an exit diameter of 586 mm and one with 10.5 degree and an exit diameter of 572 mm. For the last one two other truncated cones allow nozzle exit diameters of 1 m and 2 m. The nozzle throat diameter and therefore, the test section Mach number can also be changed by inserting different throat pieces. A contoured nozzle is available for a nominal exit Mach number of 7. In Fig. 1 a side view to scale is shown of the shock tunnel with the 5.8 degree conical nozzle. During the last years the tunnel has mainly been operated under test conditions which are listed in Table 1. Helium is used as driver gas and synthetic air as test gas. The main purpose of these conditions is to study the influence of real gas effects on hypersonic aerodynamics. Therefore, the stagnation temperature is varied between 1500 K and 4700 K which covers perfect gas behaviour at the lower and real gas behaviour at the upper limit with significant oxygen dissociation behind the bow shock. Of course, not only these test conditions can be generated but also a variety of other test section flows which are within the operating characteristics of the shock tunnel. Especially for studying the intake flow in a hypersonic propulsion engine, test condition X with an unit Reynolds number of 16.5 million per meter was calibrated.

Influence of trailing-edge brushes on upstream-moving pressure waves in transonic flow

Trailing-edge brushes on a supercritical airfoil are investigated in order to analyze their influence on upstream-moving pressure waves evolving from the trailing edge. The upstream-moving pressure waves show a highly-instationary behavior which significantly depends on the local flow velocity on the airfoil. The pressure waves are suspected to have an influence on the flow properties of the airfoil. The current results show their influence on the shock formation. Trailing-edge brushes are installed in order to reduce the amplitude of the upstream-moving pressure waves. By damping the pressure waves a reduction of dynamical loads on the airfoil becomes feasible. Experiments are preformed with a modified shock tube on the supercritical BAC 3-11 airfoil model in a Mach number range from 0.6 up to 0.8 and at a chord Reynolds number of 10. High-speed schlieren photography and pressure measurement are used to observe and quantify the strength of the waves.