Wright-Patterson Afb

Flowfield Characterization of a Rotating Detonation Engine

A rotating detonation engine (RDE) at the Air Force Research Lab (AFRL) has been modified to allow optical access to the annulus while in operation. High speed video of chemiluminescence was taken for three operating conditions to characterize the RDE flowfield. Two-dimensional representations of the entire RDE are presented to show basic flow structure. Detonation height, detonation angle, oblique shock wave angle, shear layer angle, and contact surface angle were measured. Average value for each of these measurements did not change drastically over the range tested, but large deviations of the values were observed. These considerable deviations of the flowfield point toward device variation as a major factor to be understood.

Buildup and Operation of a Rotating Detonation Engine

A rotating detonation engine (RDE), originally designed and built by Pratt & Whitney – Seattle Aerosciences Center and on loan to the Air Force Research Laboratory (AFRL/RZTC) at Wright-Patterson Air Force Base (WPAFB), was rebuilt and tested. New parts were acquired, fabricated, or modified to replace those damaged or missing from initial experimentation. Originally designed for ethylene and oxygen, the RDE was modified to run on hydrogen and air while the detonation initiator uses hydrogen and oxygen. Analytical mass-flow and pressure loss equations were used to estimate combustion chamber conditions. Estimation of detonation cell size was performed via comparison to published experimental data in terms of static pressure and stoichiometry. The experimental rig was instrumented to measure feed and fill temperatures and pressures, combustion chamber temperatures and pressures, and combustion chamber wave speed. The reactant feed system was characterized in terms of mass flow and pressure. A detonation initiation device showed successful light-off of the main chamber reactants. However, cold-flow testing revealed low static pressures in the combustion chamber unfavorable to detonation propagation. Limited combustion tests showed successful and predictable light-off, but no substantial evidence of detonation combustion. Low static pressure in the combustion chamber is thought to have inhibited propagation of the detonation from the initiation device into the main chamber.

Unsteady Performance of a Turbine Driven by a Pulse Detonation Engine

American Institute of Aeronautics and Astronautics This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States. Approved for public release; distribution is unlimited. Disclaimer: The views expressed in this presentation are those of the authors and do not reflect the official policy or position of the United States Air Force, Department of Defense, or the U.S. Government. Unsteady Performance of a Turbine Driven by a Pulse Detonation Engine

Detonation Propagation Through Ducts in a Pulsed Detonation Engine

A study of configurations to allow a consistent and predictable transition of a detonation from one detonation tube to another is presented. Development of a continuously operating pulsed detonation engine (PDE) without a high energy ignition system or a deflagration-todetonation transition (DDT) device will increase engine efficiency, reduce cost, improve performance, and reduce weight. The intent of this study was to minimize energy losses of a detonation wave in order to directly initiate another detonation wave. Detonation tube fill fraction, purge fraction, equivalence ratio, cross-over length and cross-over geometry were varied to determine their effect on direct initiation via a cross-over tube. Velocities at or above the upper Chapman-Jouguet (C-J) velocity point are desired and considered successful detonations. It was found that a cross-over tube with a “U” shaped geometry and a width at least 75% of the initiated tube’s width provided the best conditions for direct initiation.

Experimental Performance Evaluation of a Turbine Driven by Pulsed Detonations

The inlet of a radial turbine driven by pulsed detonations is shown to experience large, rapid excursions in pressures, temperatures and velocities with momentary reverse flow and unsteady accumulation and expulsion of mass and energy. The radial turbine in this study is part of a Garrett automotive turbocharger, coupled to a hydrogen-fueled pulsed detonation combustor (PDC) with 0.75 fueled and purge fractions. Time-resolved turbine inlet and exit measurements included static pressures from wall-mounted transducers, temperatures from tunable diode laser absorption spectroscopy, and velocities from laser absorption Doppler shift. Shaft power was obtained from the coupled turbine-driven turbocharger compressor. Time-resolved turbine operating points for the pulsed-detonation driven arrangement were plotted on the manufacturer steady turbine map, indicating a large operating envelope that deviated from the steady turbine operating line. Turbine inlet and exit flowfield temperatures were out of phase such that moments during the purge phase included exit total enthalpy rates greater than that of the inlet. Furthermore, portions of the pulsed detonation cycle included moments when the turbine pressure ratio was near unity. Timeresolved turbine efficiency based on a steady formulation produces non-physical results greater than 100% at some instances and negative at others. An unsteady thermodynamic model is proposed for evaluating cycle-average, mean effective turbine efficiency, accounting for unsteady heat transfer effects and including an extensive weighting parameter for turbine pressure ratio. The pulsed detonation driven turbine with a cycle-average mean effective turbine efficiency of 40% produced equivalent power as a steady deflagration driven turbine with 60% efficiency. Cycle-average specific work and mean effective turbine efficiency rose as PDC frequency increased from 20 Hz to 30 Hz.

Parametric Study of Unsteady Turbine Performance Driven by a Pulse Detonation Combustor

American Institute of Aeronautics and Astronautics This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States. Approved for public release; distribution is unlimited. Disclaimer: The views expressed in this presentation are those of the authors and do not reflect the official policy or position of the United States Air Force, Department of Defense, or the U.S. Government. Parametric Study of Unsteady Turbine Performance Driven by a Pulse Detonation Combustor

Time-Accurate Flow Field and Rotor Speed Measurements of a Pulsed Detonation Driven Turbine

Time-accurate measurements of turbine inlet and exit flow fields and rotor speed are presented for a pulsed detonation driven radial turbine, using various instrumentation techniques: flush wall-mounted static pressure transducers, background oriented Schlieren, optical pyrometry, particle streak velocimetry, laser tachometers, and variable reluctance speed sensors. The primary motivation is to evaluate instrumentation methods with sampling frequencies greater than 10 kHz, acquiring data required for future unsteady turbine performance assessments. Time-resolved temperature, pressure, and velocity are required to calculate unsteady turbine efficiency, and time-resolved rotor speed is essential for describing turbine response to detonations. Previous experimental studies of pulsed detonations have not reported flow field temperatures, pressures, and velocities at high sampling frequencies. The operating environment in a pulsed detonation driven turbine is characterized by large, rapid excursions in temperature, pressure, and mass flow. Peak gas pressures, temperatures, and velocities are on the order of 60 atm, 3000 deg K, and 1000 m/s, respectively. Rotor speeds increase more than 15,000 RPM in less than 10 ms. The current work presents unsteady results for a Garrett T3-class automotive turbocharger driven by a pulsed detonation combustor. Evaluation of time-accurate flow field instrumentation techniques is made using measurements upstream and downstream of the pulsed detonation driven radial turbine. Additionally, a comparison of rotor speed instrumentation techniques is made with measurements of compressor blade passing frequencies.

Detonation Propagation across an Asymmetric Step Expansion

A study of confined detonation transmission across a step expansion was conducted using experimental and computational techniques. Transmission success of ethylene/air detonations at various equivalence ratios was compared to the transmission of hydrogen/air detonations. The highest rate of transmission success in the experimental results for ethylene was noted at equivalence ratios richer than the conditions of minimum cell size, indicating the presence of effects other than cell size and expansion ratio. Hydrogen proved to have better transmission success than ethylene, even when normalized to the cell size upstream of the expansion. The difference is theorized to result from the disparity in critical initiation energy of the two fuels. Computational results show the presence of a relationship between the number of transverse waves upstream of the expansion and the degree of expansion that correlates to success or failure of a confined detonation transmission. Detonation transmissions are also observed to fail when a single transverse wave in the upstream channel is partially reflected at the step expansion.

Design Optimization Methods for Improving HPT Vane Pressure Side Cooling Properties Using Genetic Algorithms and Efficient CFD

Typical modern-day high pressure turbine (HPT) durability design methods in industry utilize dated correlations and spreadsheet methods based on “rules of thumb”. Of the over 2,700 film cooling references in existence, no known efforts have been made towards an optimized overall film cooling design for a realistic HPT vane geometry in proper flow conditions. Nor has there been a major attempt in open literature to improve component cooling design methods in general. This work invests greater effort in the design and optimization of a HPT vane film cooling array by way of considering numerous configurations, variables, and variable value ranges within the design space. Cooling hole surface location, size, injection orientation, and row patterns are varied in the design space. Optimization occurs by way of Latin hypercube sampling (LHS) and multi-objective genetic algorithms (GAs) to maximize the cooling effectiveness and minimize area-averaged heat transfer over the pressure surface (PS) of a baseline nozzle guide vane currently being tested experimentally within a full-scale blowdown facility. Full-map PS heat transfer predictions from 3-D computational fluid dynamics (CFD) simulations that efficiently approximate the cooling hole physics are used with prescribed fitness functions to arrive at a much improved PS cooling array design. 1,300 cooling designs were evaluated within design-space exploration that allows an extremely high number (0.32 x 10 552 ) of cooling array possibilities.

Enhanced Mixing in Supersonic Flow Using a Pulse Detonation Combustor

Pulse detonation combustors applied to a supersonic flow were investigated experimentally and numerically for their ability to enhance mixing. The high-pressure and high-temperature plume of a pulse detonation combustor exhausting into a M=2 flow was characterized through high-speed shadowgraphy and NO planar laser induced fluorescence. The pulse detonation combustor plume showed significant penetration into the flow with blow-down times greater than 4 ms and downstream pressures elevated between 1.5 and 2 times the static tunnel pressure for several milliseconds. Planar laser induced fluorescence measurements of NO captured the spanwise structure of the plume and the large counterrotating vortex structure. Numerical simulations of the pulse detonation combustor exhausting into the supersonic flow showed good agreement with shadowgraph and NO PLIF images. Injection upstream of the pulse detonation combustor showed enhanced mixing and indicated that there was an optimal separation distance between the upstream jet and pulse detonation combustor for maximum penetration and mixing with the core supersonic flow.