Roger Woodward

Deflagration to detonation transition processes by turbulence-generating obstacles in pulse detonation engines

The results from a series of detonation experiments conducted to characterize the deflagration-to-detonation transition (DDT) process for ethylene-air mixtures in a 44-mm-square, 1.65-m-long tube are described. Experiments were conducted for both single-shot detonations involving quiescent mixtures as well as multicycle detonations involving dynamic fill. For the experiments, high-frequency pressure and flame emission measurements were made to obtain the compression wave and flame speeds, respectively. In addition, schlieren and hydroxyl-radical/planar-laser-induced-fluorescence (OH-PLIF) imaging were applied to investigate the interactions between the shock-wave and combustion phenomena during both deflagration and detonation. For ethylene-air mixtures, strategically placed obstacles were necessary to achieve DDT. The effect of the presence of obstacles on flame acceleration was systematically investigated by changing the obstacle configuration. The parametric study of obstacle blockage ratio, spacing between obstacles, and length of the obstacle configuration indicated that for successful detonations the obstacle needs to accelerate the flame to a minimum flame speed of roughly half the Chapman-Jouguet detonation velocity. Differences in the flame and compression wave velocities demonstrated the development of a coupled feedback mechanism as the wave propagated along the tube. A series of simultaneous schlieren and OH-PLIF images showed that the obstacle plays a major role in generating small/large-scale turbulence that enhances flame acceleration. Localized explosions of pockets of unburned mixture further enhanced the shock-wave strength to continuously increase the flame speed. The results of this experimental study support the importance of obstacles as a means to enhance DDT and provide a potential solution for practical pulse-detonation-engine applications.

Benchmark Wall Heat Flux Data for a GO2/GH2 Single Element Combustor

Wall heat flux measurements in a 1.5 in. diameter circular cross-section rocket chamber for a uni-element shear coaxial injector element operating on gaseous oxygen (GOz)/gaseous hydrogen (GH,) propellants are presented. The wall heat flux measurements were made using arrays of Gardon type heat flux gauges and coaxial thermocouple instrumentation. Wall heat flux measurements were made for two cases. For the first case, GOZ/GHz oxidizer-rich (O/F=l65) and fuel-rich preburners (O/F=1.09) integrated with the main chamber were utilized to provide vitiated hot fuel and oxidizer to the study shear coaxial injector element. For the second case, the preburners were removed and ambient temperature gaseous oxygen/gaseous hydrogen propellants were supplied to the study injector. Experiments were conducted at four chamber pressures of 750, 600, 450 and 300psia for each case. The overall mixture ratio for the preburner case was 6.6, whereas for the ambient propellant case, the mixture ratio was 6.0. Total propellant flow was nominally 0.27-0.29 Ibm/s for the 750 psia case with flowrates scaled down linearly for lower chamber pressures. The axial heat flux profile results for both the preburner and ambient propellant cases show peak heat flux levels a t axial locations between 2.0 and 3.0 in. from the injector face. The maximum heat flux level was about two times greater for the preburner case. This is attributed to the higher injector fuel-to-oxidizer momentum flux ratio that promotes mixing and higher initial propellant temperature for the preburner case which results in a shorter reaction zone. The axial heat flux profiles were also scaled with respect to the chamber pressure to the power 0.8. The results at the four chamber pressures for both cases collapsed to a single profile indicating that at least to first approximation, the basic fluid dynamic structures in the flow field are pressure independent as long as the chamber/njector/nozzle geometry and injection velocities remain the same.

LOX/GH2 Shear Coaxial Injector Atomization Studies at Large Momentum Flux Ratios

The general view of the atomization process for gas/liquid shear-coaxial rocket engine injectors envisions a relatively short intact liquid core from which ligaments or drops are continually shed due to surface instabilities. At some point in the process, the presence of the intact liquid core ceases and only the drop field remains. This classical phenomenological breakup model indicates that the progress of liquid atomization depends primarily on the momentum flux and/or velocity ratios between the gas and liquid streams. Recent rocket combustion research experiments at gas-to-liquid momentum flux ratios (J) of about one to five had cast doubt on this breakup/atomization model for some typical injection conditions. Some rocket engines, such as the Ariane 5 Vulcain, run injectors at higher momentum fluxes (J ≈ 10-11). To evaluate the applicability of the core-stripping spray model at even extreme conditions for rocket applications, liquid oxygen (LOX) / gaseous hydrogen (GH2) combustion experiments were carried out at much higher momentum flux values (J ~ 22 and 50). A shadowgraph imaging technique was applied to record the LOX jet structure at various axial locations from which LOX dense-core length measurements were made. LOX core lengths were found to scale with the inverse of momentum flux; however, the flowfield even at J > 20, exhibits a long sinuous LOX core region, eventually breaking up into large LOX structures that gasify in the core wake. This LOX core fragmentation process seems to dominate the primary atomization process even at very high momentum flux ratios. Several existing core length correlations, none of which were developed under combustion conditions, were examined for applicability to realistic rocket engine conditions.

Experimental and Computational Investiagtion of Combustor Acoustics and Instabilities, Part II: Transverse Modes

A combined experimental-computational study of transverse acoustic modes and combustion instabilities in a rectangular liquid rocket chamber is presented. Experimental results show that transverse modes can be spontaneously excited in the rectangular chamber. The amplitudes of the acoustic response are governed by the number and location of the injector elements. In general, stronger response of the 1W mode is observed when the injector element is positioned near a pressure anti-nodal location. Companion CFD solutions of the Euler and Navier-Stokes solutions are also performed and compared with the experimental measurements. Good qualitative agreement of the acoustic chamber response is obtained. Further, the computational studies are utilized to perform parametric studies of the eects of non-linear forcing and viscous eects due to the presence of side-wall boundary layers.

Injector research for Shuttle OMS upgrade using LOX/ethanol propellants

The combustion performance, heat transfer, and stability characteristics of bi-propellant swirl coaxial, and pintle injectors using LOX and ethanol propellants were investigated at representative thrust chamber conditions. The injectors were designed for sub-scale uni-element operation based on the target geometry and operating characteristics of an upgraded non-toxic orbital maneuvering system (OMS) engine. Several different versions of the Russian heritage swirl injector were tested. Combustion efficiency, injector face temperature, high frequency pressure, and shadowgraph spray imaging results are reported for hot-fire conditions. The results exhibit a wide range of C* efficiency values. Pulsed shadowgraph images reveal the effect of several flow phenomena, such as LOX flashing and spray self-pulsation, on combustion efficiency.

DEFLAGRATION TO DETONATION TRANSITION STUDY USING SIMULTANEOUS SCHLIEREN AND OH PLIF IMAGES

A series of single-shot detonation experiments was conducted to characterize the deflagration to detonation transition (DDT) process for ethylene-air mixtures in a 45 mm square tube that was 1.65 m in length. Pressure and flame emission measurements were made to obtain the compression wave and flame speeds, while OH PLIF and Schlieren imaging were applied to investigate the combustion phenomena during both deflagration and detonation. The effect of the presence of obstacles on flame acceleration was systematically investigated by changing the obstacle configuration. A parametric study of obstacle blockage ratio, spacing between obstacles, and length of the obstacle section demonstrated that the obstacle must accelerate the flame to a minimum flame speed of roughly half the Chapman-Jouguet (C-J) detonation velocity in order for the wave to transition to a detonation. Differences in the flame and compression wave velocities demonstrated the development of a coupling mechanism as the wave propagated along the tube. A series of simultaneous Schlieren and OH PLEF images showed that the obstacle plays a major role in generating large-scale turbulence to enhance flame acceleration. Localized explosions of pockets of unburned mixture further enhanced the shock wave strength to continuously increase the flame speed. The results of this study support the importance of obstacles as a means to enhance DDT and provide a potential solution for practical pulse detonation engine applications.

High Speed Visualization of LOX/GH2 Rocket Injector Flowfield: Hot-Fire and Cold-Flow Experiments

The primary atomization and combustion characteristics of a liquid oxygen (LOX) / gaseous hydrogen (GH2) shear coaxial injector element were experimentally investigated. High speed movies using a shadowgraph imaging technique to visualize the LOX core were recorded for both hot-fire (LOX/GH2) and cold-flow (LOX/gaseous oxygen (GO2)) conditions with the same injector and chamber. Flow conditions were set to approximate realistic rocket conditions. For the hot-fire tests (LOX/GH2), chamber pressures were 600, 730, and 920 psia, with momentum flux ratios (annulus flow/post flow) of 2.7, 2.0 and 1.6 respectively. The rocket assembly utilized a preburner to provide a background flow (M≈0.1) of hot gaseous nitrogen (GN2)/GH2/water (H2O) gas with 25% volumetric concentration of hydrogen. For the cold-flow tests (LOX/GO2 with GO2 background flow), chamber pressures were 650 and 830 psia, thus above and below the critical pressure of oxygen (731.6 psia), with momentum flux ratios (annulus flow/post flow) of 2.2 and 1.8 respectively. The high speed visualizations under hot-fire conditions show a long sinuous LOX core region that breaks into large dense-oxygen structures, which are then quickly consumed. These results do not agree with the classical phenomenological breakup model that suggests a liquid core that is rapidly sheared into a drop cloud. Rather, a large-scale fragmentation model may be better suited to describe the primary atomization behavior in combusting flow from a LOX/GH2 shear coaxial injector element at realistic rocket conditions. Unlike the hot-fire case, cold-flow LOX visualization movies show a clear difference between the two chamber pressures, with the higher pressure (supercritical) case resembling behavior indicative of gaseous mixing compared to the typically two phase mixing appearance of the lower pressure (subcritical) case. Time-resolved measurements of the intact-core length are presented, along with size and frequency distributions of separating large dense-oxygen structures under hot-fire conditions.

Chamber Wall Heat Flux Measurements for a LOX/CH4 Propellant Uni-Element Rocket

Wall heat flux measurements in a 1.0 in. diameter circular cross-section rocket chamber for three uni-element injector elements operating on liquid oxygen (LOX) / gaseous methane (GCH4) propellants are presented. The wall heat flux measurements were made using a rocket chamber instrumented with coaxial thermocouples. Wall heat flux measurements were made for three uni-element injectors, viz., two versions of a shear coaxial element and a swirl coaxial element. The three injectors were designed for a chamber pressure of 1200 psia, LOX flowrate of 0.7 lbm/s and mixture ratio of 3.0. Experiments were conducted at the design pressure of 1200 psia and also at reduced pressures of 1000, 800, 600 and 300 psia for each injector at the design mixture ratio of 3.0 and also at 2.5 and 3.25. For experiments at the lower pressures, the propellant mass flowrates were scaled down linearly. The local wall heat flux measurements show higher heat flux levels for the swirl coaxial injector than the two versions of the shear coaxial injector at near injector face locations. This is attributed to enhanced LOX atomization, mixing and combustion provided by the conical swirling spray in the near injector face region for the swirl coaxial injector. The two tested shear coaxial injectors differ in the design fuel-to-oxidizer momentum flux ratios. The shear coaxial injector with the higher fuel-to-oxidizer momentum flux ratio showed higher heat flux levels in the near injector face region. One of the two shear coaxial injectors was designed such that the LOX post could be configured flush or recessed with respect to the injector face. The configuration with the LOX post recessed showed higher heat flux levels in the near injector face region than its LOX post flush counterpart, indicating that the mixing cup provided by recessing the LOX post has a positive effect on the mixing and combustion characteristics of the injector. Finally, the axial wall heat flux profiles for different chamber pressures were scaled with respect to chamber pressure to the power 0.8. The scaling brought the profiles closer together but not to the extent of collapsing the data, indicating that for liquid/gas injectors where the fuel-to-oxidizer momentum flux ratio decreases with chamber pressure, the resulting coupled atomization/mixing/combustion phenomena does not scale simply with pressure.

Combustion Instability Studies Using Gaseous Methane and Liquid Oxygen

Experimental studies of combustion instabilities in a gaseous methane and liquid oxygen multi-element rectangular chamber are presented. Various methods of driving combustion instabilities in the chamber are examined including mass flowrate effects, propellant flow modulation and introduction of a non reacting nitrogen flow at an inter-element location. Combustion instabilities observed were random in time and exhibited burst behavior, decaying out after 5 to 20 cycles. Only the introduction of the nitrogen flow at an inter- element location showed a consistent increase in the magnitude of the observed pressure oscillation. This effect is argued to be due to simultaneous enhanced atomization between the two adjacent injector elements.

Diode Laser Absorption Spectroscopy Measurements in a Gaseous Hydrogen / Oxygen Rocket

Tunable diode laser absorption spectroscopy (TDLAS) has been used to make pathintegrated temperature measurements in a gaseous oxygen / gaseous hydrogen uni-element rocket chamber with hot background flow. Four mixture conditions were studied at a nominal chamber pressure of 100 psia. Near infrared diode lasers were utilized to target rovibrational transitions of water vapor combustion product. Both direct absorption spectroscopy and wavelength modulation spectroscopy with second harmonic normalized by first harmonic (1f-normalized WMS-2f) techniques were applied, with the harmonic detection technique found to yield the best results. Illustrative examples of time-resolved measurements of temperature in the rocket chamber are presented.