Pulse Detonation Combustor Research Articles

Effect of Blockage Ratio on Detonation Flame Acceleration in Pulse Detonation Combustor Using CFD

Detonation is the supersonic mode of combustion process which is essential for energy release from combustion process. Detonation is the more energetic process compare to deflagration mode of combustion process. The turbulence combustion flame cannot transit itself into detonation combustion process. So objective of this paper is to investigate the effect of obstacles configuration landed in detonation tube channel to propagate the detonation wave and diffraction encounters in an obstacles site. Four different cases of obstacles blockage ratio (BR) 0.4, 0.5, 0.6 and 0.7 were studied for detonation flame acceleration in detonation tube. A three dimensional computational simulation was done using unsteady green-gauss cell based solver for adopting the combustion simulation. As a result detonation flame propagation, detonation flame velocity and detonation flame pressure were increase in reducing blockage ratio from 0.7 to 0.4 and eddy viscosity of combustible mixture was increase with increasing the blockage ratio. From the analyzed four blockage ratio BR=0.4 is suitable for detonation mode of combustion and flame acceleration.

Shockless Explosion Combustion: An Innovative Way of Efficient Constant Volume Combustion in Gas Turbines

Constant volume combustion (CVC) in gas turbines is a promising way to achieve a step change in the efficiency of such systems. The most widely investigated technique to implement CVC in gas turbine systems is pulsed detonation combustion (PDC). Unfortunately, the PDC is associated with several disadvantages, such as sharp pressure transitions, entropy generation due to shock waves, and exergy losses due to kinetic energy. This work proposes a new way to implement CVC in a gas turbine combustion system: shockless explosion combustion (SEC). This technique utilizes acoustic waves inside the combustor to fill and purge the combustion tube. The combustion itself is controlled via the ignition delay time of the fuel/air mixture. By adjusting the ignition delay in a way such that the entire fuel/air volume undergoes homogeneous auto-ignition, no shock waves occur. Accordingly, the losses associated with a detonation wave are not present in the proposed system. Instead, a smooth pressure rise is created due to the heat release of the homogeneous combustion. The current article explains the SEC process in detail, and presents the identified challenges. Solutions to these challenges and the numerical and experimental approach are presented subsequently alongside with first preliminary results of the numerical studies.

Time-Resolved Flow Properties in a Turbine Driven by Pulsed Detonations

Theoretically, pressure-gain combustion like that of a pulsed-detonation combustor allows heat addition with reduced entropy loss compared to a typical steady deflagration combustor. The pulsed-detonation combustor is inherently unsteady, and time-resolved measurements are required to identify unsteady flow processes at the turbine inlet and exit. In this study, the radial turbine of a Garrett automotive turbocharger was coupled directly to and driven, full admission, by a pulsed-detonation combustor fueled by hydrogen or ethylene. Full-cycle time-resolved turbine inlet and exit pressures were obtained from pressure transducers mounted to wall static pressure ports. Tunable diode laser absorption spectroscopy was used to obtain full-cycle time-resolved turbine inlet and exit temperatures and velocities. Optical access permitted the use of high-speed video for particle streak velocimetry to obtain time-resolved turbine inlet velocities during blowdown with a 20 Hz ethylene-fueled pulsed-detonation combustor with 0.9 fueled fraction and 0.5 purge fraction. A comparison of laser-based and video-based measurements was made for the blowdown portion of the pulsed-detonation combustor cycle. First-ever full-cycle time histories of turbine inlet and exit velocities, total temperature, total pressure, mass flow, and total enthalpy rate revealed reflected pressure waves and showed momentary reverse flow, and thus unsteady accumulation and expulsion of mass and enthalpy.

Infrared laser absorption sensors for multiple performance parameters in a detonation combustor

Laser absorption sensors for temperature and H2O near 1.4 and 2.5μm, CO2 near 2.7μm, and CO near 4.8μm were developed, validated, and deployed in an ethylene-fueled pulse-detonation combustor (PDC). Each sensor was fiber-coupled to enable remote light delivery to the PDC and photo-voltaic detectors were mounted directly to the PDC for improved light collection. Measurements were acquired simultaneously along two orthogonal lines-of-sight (LOS) in both the PDC combustion chamber and the throat of a converging–diverging nozzle located at the PDC exit. Measurements in the nozzle throat were combined with a choked-flow assumption to calculate the time-resolved enthalpy flow rate exiting the PDC. All sensors used first-harmonic-normalized wavelength-modulation spectroscopy with second-harmonic detection (WMS-2f/1f) to account for non-absorbing transmission losses and emission. Furthermore, strong mid-infrared absorption transitions were used to enable improved measurement accuracy and precision. These sensors were validated in non-reactive shock-tube experiments at temperatures and pressures up to 2700K and 50atm. There, these sensors exhibited a nominal accuracy of 3 to 5% with bandwidths from 2 to 20kHz. Measurements in the PDC indicated peak temperatures near 3500K with near-stoichiometric proportions of H2O and incomplete conversion of CO to CO2. The stagnation enthalpy flow rate was highly transient, but repeatable with peak flow rates near 25MW.

Quantum cascade laser absorption sensor for carbon monoxide in high-pressure gases using wavelength modulation spectroscopy

A tunable quantum cascade laser sensor, based on wavelength modulation absorption spectroscopy near 4.8 μm, was developed to measure CO concentration in harsh, high-pressure combustion gases. The sensor employs a normalized second harmonic detection technique (WMS-2f/1f) at a modulation frequency of 50 kHz. Wavelength selection at 2059.91 cm⁻¹ targets the P(20) transition within the fundamental vibrational band of CO, chosen for absorption strength and relative isolation from infrared water and carbon dioxide absorption. The CO spectral model is defined by the Voigt line-shape function, and key line-strength and line-broadening spectroscopic parameters were taken from the literature or measured. Sensitivity analysis identified the CO-N₂ collisional broadening coefficient as most critical for uncertainty mitigation in hydrocarbon/air combustion exhaust measurements, and this parameter was experimentally derived over a range of combustion temperatures (1100-2600 K) produced in a shock tube. Accuracy of the wavelength-modulation-spectroscopy-based sensor, using the refined spectral model, was validated at pressures greater than 40 atm in nonreactive shock-heated gas mixtures. The laser was then free-space coupled to an indium-fluoride single-mode fiber for remote light delivery. The fiber-coupled sensor was demonstrated on an ethylene/air pulse detonation combustor, providing time-resolved (~20 kHz), in situ measurements of CO concentration in a harsh flow field.

Wavelength-modulation spectroscopy near 1.4 µm for measurements of H2O and temperature in high-pressure and -temperature gases

The development, validation and demonstration of a two-color tunable diode laser (TDL) absorption sensor for measurements of temperature and H2O in high-pressure and high-temperature gases are presented. This sensor uses first-harmonic-normalized wavelength-modulation spectroscopy with second-harmonic detection (WMS-2f/1f) to account for non-absorbing transmission losses and emission encountered in harsh, high-pressure environments. Two telecommunications-grade TDLs were used to probe H2O absorption transitions near 1391.7 and 1469.3 nm. The lasers were frequency-multiplexed and modulated at 160 and 200 kHz to enable a measurement bandwidth up to 30 kHz along a single line-of-sight. In addition, accurate measurements are enabled at extreme conditions via an experimentally derived spectroscopic database. This sensor was validated under low-absorbance (<0.05) conditions in shock-heated H2O–N2 mixtures at temperatures and pressures from 700 to 2400 K and 2 to 25 atm. There, this sensor recovered the known temperature and H2O mole fraction with a nominal accuracy of 2.8% and 4.7% RMS, respectively. Lastly, this sensor resolved expected transients with high bandwidth and high precision in a reactive shock tube experiment and a pulse detonation combustor.

Natural-Gas-Fueled Pulse-Detonation Combustor

The feasibility of controlled repeatable (with a frequency of up to 2 Hz) deflagration-to-detonation transition within a length of and further detonation propagation at an average velocity above in a 150-mm-diam tube, with an open end and separate supply of natural gas and air as fuel components, has been demonstrated experimentally. Based on experimental studies, a model of a pulsed-detonation combustor, and a prototype of industrial burners of a new generation, which produces a combined shock wave (mechanical) and thermal effect on objects blown on with combustion products, has been designed, constructed, and tested. Reported are the results of experiments demonstrating the pulsed-detonation combustor operation process in terms of cyclic detonation initiation via deflagration-to-detonation transition, as well as pulsed-detonation combustor thermal regime during long-duration (300 s) testing without forced cooling. The results provide a basis for further work aimed at increasing the operation frequency and thermal power of pulsed-detonation combustors for different industrial (metallurgy, chemical engineering, waste incineration, etc.) and propulsion applications.

First and Second Law Analysis of Future Aircraft Engines

An optimal baseline turbofan cycle designed for a performance level expected to be available around year 2050 is established. Detailed performance data are given in take-off, top of climb, and cruise to support the analysis. The losses are analyzed, based on a combined use of the first and second law of thermodynamics, in order to establish a basis for a discussion on future radical engine concepts and to quantify loss levels of very high performance engines. In light of the performance of the future baseline engine, three radical cycles designed to reduce the observed major loss sources are introduced. The combined use of a first and second law analysis of an open rotor engine, an intercooled recuperated engine, and an engine working with a pulse detonation combustion core is presented. In the past, virtually no attention has been paid to the systematic quantification of the irreversibility rates of such radical concepts. Previous research on this topic has concentrated on the analysis of the turbojet and the turbofan engine. In the developed framework, the irreversibility rates are quantified through the calculation of the exergy destruction per unit time. A striking strength of the analysis is that it establishes a common currency for comparing losses originating from very different physical sources of irreversibility. This substantially reduces the complexity of analyzing and comparing losses in aero engines. In particular, the analysis sheds new light on how the intercooled recuperated engine establishes its performance benefits.

Experimental investigations on the power extraction of a turbine driven by a pulse detonation combustor

In order to grasp the interaction mechanism between the pulse detonation combustor (PDC) and the turbine, the experimental work in this paper investigates the key factors on the power extraction of a turbocharger turbine driven by a PDC. A PDC consisting of an unvalved tube is integrated with a turbocharger turbine which has a nominal mass flow rate of 0.6kg/s and 50000r/min. The PDC–turbine hybrid engine is operated on gasoline-air mixtures and runs for 6+min to achieve a thermal steady state, and then the engine performance is evaluated under different operating conditions. Results show that the momentum difference per unit area between the turbine inlet and outlet plays an important role in the power extraction, while the pressure peak of the detonation has little effect. The equivalence ratio of fuel and air mixture and the transition structure between PDC and turbine are also important to the power extraction of the turbine. The present work is promising as it suggests that the performance benefit of a PDC–turbine hybrid engine can be realized by increasing the momentum difference per unit area through the optimal design of transition section between the PDC and turbine.

Measurements of multiple gas parameters in a pulsed-detonation combustor using time-division-multiplexed Fourier-domain mode-locked lasers

Hyperspectral absorption spectroscopy is being used to monitor gas temperature, velocity, pressure, and H(2)O mole fraction in a research-grade pulsed-detonation combustor (PDC) at the Air Force Research Laboratory. The hyperspectral source employed is termed the TDM 3-FDML because it consists of three time-division-multiplexed (TDM) Fourier-domain mode-locked (FDML) lasers. This optical-fiber-based source monitors sufficient spectral information in the H(2)O absorption spectrum near 1350 nm to permit measurements over the wide range of conditions encountered throughout the PDC cycle. Doppler velocimetry based on absorption features is accomplished using a counterpropagating beam approach that is designed to minimize common-mode flow noise. The PDC in this study is operated in two configurations: one in which the combustion tube exhausts directly to the ambient environment and another in which it feeds an automotive-style turbocharger to assess the performance of a detonation-driven turbine. Because the enthalpy flow [kilojoule/second] is important in assessing the performance of the PDC in various configurations, it is calculated from the measured gas properties.

Fiber-coupled 2.7 µm laser absorption sensor for CO2 in harsh combustion environments

A tunable diode laser absorption sensor near 2.7 µm, based on 1f-normalized wavelength-modulation spectroscopy with second-harmonic detection (WMS-2f), was developed to measure CO2 concentration in harsh combustion flows. Wavelength selection at 3733.48 cm−1 exploited the overlap of two CO2 transitions in the ν1 + ν3 vibrational band at 3733.468 cm−1 and 3733.498 cm−1. Primary factors influencing wavelength selection were isolation and strength of the CO2 absorption lines relative to infrared water absorption at elevated pressures and temperatures. The HITEMP 2010 database was used to model the combined CO2 and H2O absorption spectra, and key line-strength and line-broadening spectroscopic parameters were verified by high-temperature static cell measurements. To validate the accuracy and precision of the WMS-based sensor, measurements of CO2 concentration were carried out in non-reactive shock-tube experiments (P ∼ 3–12 atm, T ∼ 1000–2600 K). The laser was then free-space fiber-coupled with a zirconium fluoride single-mode fiber for remote light delivery to harsh combustion environments, and demonstrated on an ethylene/air pulse detonation combustor at pressures up to 10 atm and temperatures up to 2500 K. To our knowledge, this work represents the first time-resolved in-stream measurements of CO2 concentration in a detonation-based engine.

Experimental investigation on wall temperature of an air-breathing kerosene/air pulse detonation combustor with impingement cooling

Abstract The design of a practical pulse detonation combustor requires the components to be capable of enduring the severe thermal environment created by repetitive detonations. In the present study, an air-breathing kerosene/air detonation combustor was designed and fabricated. And the temperature distributions on the combustor liner were measured at various operating frequencies (from 10 Hz to 50 Hz) for both natural cooling and jet impinging cooling cases. The results show that the temperature distribution on detonation combustor liner under natural cooling mode is seriously non-uniform and the hottest region appears corresponding to where transition from deflagration to detonation occurs. The temperature rise amplitude corresponding to 10 Hz increase at higher operational frequency is smaller than that at lower operational frequency. As expected, the maximum temperature on detonation combustor liner is decreased as the increase of coolant flow rate. The impinging distance between jet orifice tube and circular liner is of important influence on the reasonable temperature distribution of a cooled combustor liner.

Optical and thrust measurement of a pulse detonation combustor with a coaxial rotary valve

We developed a rotary valve for a pulse detonation engine (PDE), and confirmed its basic characteristics and performance. In a square cross-section combustor, we visualized a multi-shot of a pulse detonation rocket engine (PDRE) cycle at an operation frequency of 160 Hz by using a high-speed camera (time resolution: 3.33 μs, space resolution: 0.4 mm) and a Schlieren method. The propellant filling process and the purge process were confirmed, and each process was modeled. Moreover, we confirmed the processes of detonation wave generation and burned gas blowdown. In addition, we investigated the impact of shortening the passage width of a combustor and negative-time ignition (ignition time is earlier than the end-time of the propellant filling process) on the deflagration-to-detonation transition (DDT) distance and time. The DDT distance did not depend on the passage width of a combustor and decreased under the negative-time ignition condition. With a passage width of 20 mm, the DDT distance decreased by 22% under the negative-time ignition condition to a minimum value (76 ± 8 mm). The DDT time from spark time reached a minimum value (69 ± 14 μs) under the condition of a passage width of 10 mm and negative-time ignition. The detonation initiation time and the DDT distance were represented by the time until the flame expanded toward the tube-axis one-dimensionally from ignition (characteristic time). We also carried out thrust measurement using a PDRE system composed of a circular cross-section combustor and the newly developed valve. We obtained a stable time-averaged thrust in a wide range of operation frequency (40–160 Hz) and confirmed the increase of specific impulse due to a partial-fill effect. At a maximum operation frequency of 159 Hz, we achieved a maximum propellant-based specific impulse of 232 s and a maximum time-averaged thrust of 71 N.

Thermodynamic performance analysis of turbofan engine with a pulse detonation duct heater

The potential performance gain of utilizing pulse detonation combustion in the bypass duct of a turbofan engine was investigated in this study. Pulse detonation turbofan engine concepts were studied and their performances were assessed. There were six cases according to different bypass ratio, different exhaust methods and with Pulse Detonation Combustor (PDC) in bypass duct or not. Four study combinations were established to compare the performance of different cases. The specific thrust (Fs), specific fuel consumption (sfc) of conventional turbofan engines and the new pulse detonation turbofan engine concepts were calculated and compared on design point and with the change of flight Mach number or altitude. The nozzle performance was considered in the present work. The calculation methods of PDC in bypass duct with different exhaust methods were explored. And the calculation methods of PDC adopted in this research could be a reference for other hybrid propulsion system with PDC.

Experimental Investigations of the Performance of a Multitube Pulse Detonation Turbine System

A multitube pulse detonation combustor consisting of eight unvalved tubes arranged in a can-annular configuration was integrated with a single-stage axial turbine nominally rated for 10 lbm=s, 25,000 rpm and 1000 hp. The multitube pulse-detonation-combustor-turbine hybrid was operated using ethylene-air mixtures for runs of 5 min to achieve thermal steady state in order to quantify performance at two operating conditions using simultaneous and sequential firing patterns. Analysis of these data reveal that turbine efficiency under pulsedetonation-combustor-fired operation was indistinguishable from steady performance within the 8 points measurement uncertainty. Furthermore, the pulse-detonation-combustor-turbine hybrid system demonstrated a potential 25% improvement in efficiency once corrections were made for the suboptimal fueling design, which resulted in lower-than-anticipated combustion efficiency. The present work is promising as it suggests that a pulsedetonation-combustor-turbine hybrid engine performance benefit can be realized with further development.

Performance Evaluation of Hybrid Gas Turbine Engine Embedded with Pulse Detonation Combustor

The numerical investigations of performance evaluation of a hybrid gas turbine engine embedded with a pulse detonation combustor (PDC) were performed to examine the improvement of the performance of the hybrid propulsion system. The calculation model and method were described. The architecture, configuration and size of detonation tubes were investigated in the calculation. Two models of detonation tube exit temperature were utilized. Eight configuration choices for the PDC based on the calculation model were designed. Specific fuel consumption of a hybrid gas turbine engine was compared with that of the baseline engine at the condition of the same engine net thrust. The experimental research of a PDC interacted with a radial flow turbine of a turbocharger was conducted. The numerical results show that if the net thrust of hybrid PDC engine is matched to that of baseline engine, specific fuel consumption of hybrid PDC engine is 20–25% less than that of baseline engine. The total volume of the hybrid engine combustor is reduced. The incorporation of PDC into gas turbine engine can improve the performance of hybrid PDC engine, decrease the combustor weight, and increase the thrust-weight ratio. The experimental results show that the fully developed detonation waves are achieved in the experimental apparatus.

Effects of Pulse Detonation Wave on Film Dynamics

Thermal protection is very important and necessary for the design of a practical, long-life pulse detonation combustor which suffers repetitive detonation attack. Film cooling is an attractive technique which has been successfully used in aero-engine but seldom mentioned in pulsed detonation engine. In the present study, unsteady CFD (Computational Fluid Dynamics) simulation of the dynamic performance of cooling film flows subjected to forced cyclic detonation waves was performed. It is seen that the cooling film could be sustained on the tube wall for a substantial duration of the detonative cycle. The dynamic interaction of the detonation wave with the film flow is an important issue because it has obvious effect on the structure and stability of the cooling film. Although there is a brief period of reverse flow in the cooling hole, the positive film flow is quickly re-established and leads to an effective heat-flux reduction in the post-detonation period. The effect of film cooing on a tube was validated using the experimental test results. It indicated a decrease in heat flux with cooling flow. This is a clear indication that the cooling film is effective on reducing the heat-flux of the detonation tube.

Experimental Investigations of a Pulse Detonation Combustor Integrated with a Turbine

Experimental Investigations of a Pulse Detonation Combustor Integrated with a Turbine

Schlieren Visualization of Multicyclic Flame Acceleration Process in Valveless Pulsed Detonation Combustors

Turbulent flame acceleration process has been studied in a two-dimensional valveless pulsed detonation combustor (PDC) using Schlieren flow visualization, ionization probes, and companion high-frequency pressure instrumentation. Experiments were conducted using near stoichiometric propane-air mixtures in a 76.2 mm × 76.2 mm, 1.68 m long modular detonation tube with the windowed combustor under multicyclic conditions within the PDC operating in a valveless mode. The effect of the presence of the obstacle was investigated at various stages of the flame acceleration process in aerodynamically confined environment. Amplification of the leading shock waves through the interaction of the obstacle enhances the flame acceleration process owing to the reduction of post-shock induction time and the hydrodynamic instabilities at the baroclinic interface. The internal configuration of the combustor including the obstacle plays a major role of the flame acceleration due to the shock amplification pattern. Sequences of schlieren images also show the flow drifting and blowdown processes confirming that the gasdynamic valves are successfully operative in the valveless PDC.

Multicyclic-Detonation-Initiation Studies in Valveless Pulsed Detonation Combustors

A pulsed detonation combustor was developed with the total absence of mechanical valves. The device was operated over a range of conditions with air and gaseous fuels (ethylene and propane) with both fuel streams and airstreams operating in valveless mode. This configuration allows a simplified design that could be used for propulsion and power generation systems as well as combined-cycle propulsion concepts. High-frequency pressure instrumentation reveals that overdriven detonations and Chapman-Jouguet detonations are generated within the device using both fuels. A novel vortex-generator design is employed to promote the deflagration-to-detonation transition process in a cyclic fashion. Pressure instrumentation within inlet manifolds confirms the capabilities of the configuration to cease fuel and air flows over a portion of the cycle. The character of wave development in the device is discussed, and a performance map is included to denote the range of operation in valveless mode.