Daniel E. Paxson

Unsteady Ejector Performance:An Experimental Investigation Using a Pulsejet Driver

An experimental investigation is described in which thrust augmentation and mass entrainment were measured for a variety of simple cylindrical ejectors driven by a gasoline-fueled pulsejet. The ejectors were of varying length, diameter, and inlet radius. Measurements were also taken to determine the effect on performance of the distance between pulsejet exit and ejector inlet. Limited tests were also conducted to determine the effect of driver cross-sectional shape. Optimal values were found for all three ejector parameters with respect to thrust augmentation. This was not the case with mass entrainment, which increased monotonically with ejector diameter. Thus, it was found that thrust augmentation is not necessarily directly related to mass entrainment, as is often supposed for ejectors. Peak thrust augmentation values of 1.8 were obtained. Peak mass entrainment values of 30 times the driver mass flow were also observed. Details of the experimental setup and results are presented. Preliminary analysis of the results indicates that the enhanced performance obtained with an unsteady jet (primary source) over comparably sized ejectors driven with steady jets is due primarily to the structure of the starting vortex-type flow associated with the former.

Parametric investigation of thrust augmentation by ejectors on a pulsed detonation tube

A parametric investigation has been made of thrust augmentation of a 1 in. diameter pulsed detonation tube by ejectors. A set of ejectors was used which permitted variation of the ejector length, diameter, and nose radius, according to a statistical design of experiment scheme. The maximum augmentation ratios for each ejector were fitted using a polynomial response surface, from which the optimum ratios of ejector diameter to detonation tube diameter, and ejector length and nose radius to ejector diameter, were found. Thrust augmentation ratios above a factor of 2 were measured. In these tests, the pulsed detonation device was run on approximately stoichiometric air-hydrogen mixtures, at a frequency of 20 Hz. Later measurements at a frequency of 40 Hz gave lower values of thrust augmentation. Measurements of thrust augmentation as a function of ejector entrance to detonation tube exit distance showed two maxima, one with the ejector entrance upstream, and one downstream, of the detonation tube exit. A thrust augmentation of 2.5 was observed using a tapered ejector.

Active Combustion Control for Aircraft Gas Turbine Engines

John C. DeLaat, Kevin J. Breisacher, Joseph R. Saus, and Daniel E. PaxsonGlenn Research Center, Cleveland, OhioPrepared for the36th Joint Propulsion Conference and Exhibitionsponsored by the American Institute of Aeronautics and AstronauticsHuntsville, Alabama, July 17-19, 2000National Aeronautics andSpace AdministrationGlenn Research Center

An improved numerical model for wave rotor design and analysis

A numerical model has been developed which can predict both the unsteady flows within a wave rotor and the steady averaged flows in the ports. The model is based on the assumptions of one-dimensional, unsteady, and perfect gas flow. Besides the dominant wave behavior, it is also capable of predicting the effects of finite tube opening time, leakage from the tube ends, and viscosity. The relative simplicity of the model makes it useful for design, optimization, and analysis of wave rotor cycles for any application. This paper discusses some details of the model and presents comparisons between the model and two laboratory wave rotor experiments.

Unsteady Ejector Performance: An Experimental Investigation Using a Resonance Tube Driver

AbstractA statistically designed experiment to characterizethrust augmentation for unsteady ejectors has beenconducted at the NASA Glenn Research Center. Thevariable parameters included ejector diameter, length,and nose radius. The pulsed jet driving the ejectors wasproduced by a shrouded resonance (or Hartmann-Sprenger) tube. In contrast to steady ejectors, anoptimum ejector diameter was found, which coincidedwith the diameter of the vortex ring created at thepulsed jet exit. Measurements of ejector exit velocityusing a hot-wire permitted evaluation of the massaugmentation ratio, which was found to correlate tothrust augmentation following a formula derived forsteady ejectors.IntroductionCurrently, efforts are underway to explore the use ofpulsed detonation engines (PDE) for aerospacepropulsion. Technical issues involved includeintegration, noise, and thrust to weight ratio. Adding anejector to a PDE may enhance thrust, and reduce noise.The ejector will then be driven by a pulsating flow. Paststudies of unsteady ejectors

Ejector Enhanced Pulsejet Based Pressure Gain Combustors: An Old Idea With a New Twist

Abstract An experimental investigation of pressure-gain combustion for gas turbine application is described. The test article consists of an off-the-shelf valved pulsejet, and an optimized ejector, both housed within a shroud. The combination forms an effective ‘can’ combustor across which there is a modest total pressure rise rather than the usual loss found in conventional combustors. Although the concept of using a pulsejet to affect semi-constant volume (i.e., pressure-gain) combustion is not new, that of combining it with a well designed ejector to efficiently mix the bypass flow is. The result is a device which to date has demonstrated an overall pressure rise of approximately 3.5 percent at an overall temperature ratio commensurate with modern gas turbines. This pressure ratio is substantially higher than what has been previously reported in pulsejet-based combustion experiments. Flow non-uniformities in the downstream portion of the device are also shown to be substantially reduced compared to those within the pulsejet itself. The standard deviation of total pressure fluctuations, measured just downstream of the ejector was only 5.0 percent of the mean. This smoothing aspect of the device is critical to turbomachinery applications since turbine performance is, in general, negatively affected by flow non-uniformities and unsteadiness. The experimental rig will be described and details of the performance measurements will be presented. Analyses showing the thermodynamic benefits from this level of pressure-gain performance in a gas turbine will also be assessed for several engine types. Issues regarding practical development of such a device are discussed, as are potential emissions reductions resulting from the rich burning nature of the pulsejet and the rapid mixing (quenching) associated with unsteady ejectors.

Experimental Investigation of Unsteady Thrust Augmentation Using a Speaker-Driven Jet

An experimental investigation is described in which a simple speaker-driven jet was used as a pulsed thrust source (driver) for an ejector configuration. The objectives of the investigation were twofold. The first was to expand the experimental body of evidence showing that an unsteady thrust source, combined with a properly sized ejector generally yields higher thrust augmentation values than a similarly sized, steady driver of equivalent thrust. The second objective was to identify characteristics of the unsteady driver that may be useful for sizing ejectors, and for predicting the thrust augmentation levels that may be achieved. The speaker-driven jet provided a convenient source for the investigation because it is entirely unsteady (i.e., it has no mean velocity component) and because relevant parameters such as frequency, time-averaged thrust, and diameter are easily variable. The experimental setup will be described, as will the two main measurements techniques employed. These are thrust and digital particle imaging velocimetry of the driver

Experimental and Numerical Evaluation of Pressure Gain Combustion in a Rotating Detonation Engine

The detonation structure, pressure gain, and thrust production in a rotating detonation engine (RDE) are studied using a combination of experimental and numerical approaches. High frequency time-dependent and low frequency time-averaged static pressure and thrust measurements are acquired for a range of operating conditions and geometry configurations. Acoustic coupling between the detonation channel and air plenum is important for low air mass flow rates and large air injection slots based on analyses of the pressure measurements in the time and frequency domains. The static pressure increases across the air inlet by up to approximately 15% when utilizing a large air injection slot. The pressure increase across the air inlet demonstrates encouraging progress towards realizing pressure gain combustion in RDEs with corresponding challenges associated with isolating the inlet plenums. The time-dependent pressure measurements acquired using a semi-infinite tube arrangement and time-averaged pressure measurements acquired using a capillary tube attenuated arrangement agree to within 30% depending upon location. Quantification of the similarities and differences between the two techniques represents important progress towards acquiring quantitative time-dependent pressure measurements in the challenging environment presented by RDEs. Twodimensional simulations of the RDE capture the essential features of the flow field such as the detonation wave height and angle, trailing edge oblique shock wave, shear layer between the freshly and previously detonated products, and deflagration between the fuel fill region and expansion region containing detonated products. The presence of air purging from the plenum to the channel behind the detonation wave is suggested by the comparison of measured and simulated channel pressure distributions. The pressure, thrust, and wave speed measurements provide benchmark data that are useful for evaluating low and high fidelity simulations of RDEs and improving fundamental understanding of the critical design parameters that influence RDE operation and performance.

Experimental Results of Performance Tests on a Four-Port Wave Rotor

A series of tests has been performed on a four-port wave rotor suitable for use as a topping stage on a gas turbine engine, to measure the overall pressure ratio obtainable as a function of temperature ratio, inlet mass flow, loop flow ratio, and rotor speed. The wave rotor employed an open high pressure loop that is the high pressure inlet flow was not the air exhausted from the high pressure outlet, but was obtained from a separate heated source, although the mass flow rates of the two flows were balanced. This permitted the choice of a range of loop-flow ratios (i.e., ratio of high pressure flow to low pressure flow), as well as the possibility of examining the effect of mass flow imbalance. Imbalance could occur as a result of leakage or deliberate bleeding for cooling air. Measurements of the pressure drop in the high pressure loop were also obtained. A pressure ratio of 1.17 was obtained at a temperature ratio of 2.0, with an inlet mass flow of 0.6 lb/s. Earlier tests had given a pressure ratio of less than 1.12. The improvement was due to improved sealing between the high pressure and low pressure loops, and a modification to the movable end-wall which is provided to allow for rotor expansion.

Assessment of the Performance of a Pulsejet and Comparison with a Pulsed-Detonation Engine

The performance of a Solar PJ32 pulsejet engine, which is a 1/5-scale model of the Argus V-1 pulsejet engine developed for the Navy in 1951, is evaluated under static conditions and compared with that of a pulsed-detonation engine (PDE) firing at similar inlet and operating conditions. The pulsejet has a fuel-flow operating range of 2.5-4.5 lbm/min, which corresponds to a thrust range of 40 lbf (at lean out) to 102 lbf (at flood out). Thrust is calculated from combustion-chamber pressure histories and agrees with measured thrust within 5-10%. Peak combustion-chamber head pressures range from 8 to 20 psig, while significantly higher pressures (80-120 psig) are attained in PDEs. Airflow at the inlet of the pulsejet is measured and used to calculate specific thrust and equivalence ratio. Specific thrust ranges from 40-100 lbf-s/lbm over the range of fuel flows from lean to rich conditions. A similarly operating PDE has a specific thrust around 120 lbf-s/lbm, making the PDE more efficient in terms of air flow. The pulsejet equivalence ratio ranges from 0.6-1.0, with rated/peak thrust occurring at rich conditions. Typical fuel-specific impulse (Isp) for the pulsejet is 1400-1500 s for rated thrust conditions, whereas PDE performance (with a fill fraction of 1) is around 1800 s. For the PDE operating in the same fill fraction range as the pulsejet (~0.1), PDE Isp is estimated to be 6000-8000 s making the PDE cycle far more efficient and desirable at comparable conditions.