Showing papers in "Journal of Propulsion and Power in 1994"

Design of optimum propellers

Abstract: Improvements have been made in the equations and computational procedures for design of propellers and wind turbines of maximum efficiency. These eliminate the small angle approximation and some of the light loading approximations prevalent in the classical design theory. An iterative scheme is introduced for accurate calculation of the vortex displacement velocity and the flow angle distribution. Momentum losses due to radial flow can be estimated by either the Prandtl or Goldstein momentum loss function. The methods presented here bring into exact agreement the procedure for design and analysis. Furthermore, the exactness of this agreement makes possible an empirical verification of the Betz condition that a constant-displacement velocity across the wake provides a design of maximum propeller efficiency. A comparison with experimental results is also presented.

Advanced Injection and Mixing Techniques for Scramjet Combustors

Abstract: Scramjet combustor fuel injection and mixing enhancement techniques are reviewed. The injection techniques include hole injection from the combustor wall, slot injection parallel to the flow, and injection from struts and the rear of ramps. Three new advanced mixing techniques are presented. The first is a combustor, curved so that buoyancy forces will aid in the penetration of the fuel across the combustor. The second is pulsation of the fuel injectors to increase penetration and mixing. A fluidic technique, a modified Hartmann-Sprenger tube, is identified as a strong candidate to generate the pulsations. The third is the injection behind pylons to allow deep penetration into the air stream. This technique is likely to produce high base pressures on the injector structure, particularly if base burning is encouraged. Curved or slanted pylons can be used to increase the recovery of fuel jet momentum. The potential of the new mixing techniques to increase scramjet engine performance is assessed.

Experimental investigation of a supersonic swept ramp injector using laser-induced iodine fluorescence

Abstract: Planar measurements of injectant mole fraction and temperature have been conducted in a nonreacting supersonic combustor configured with underexpanded injection in the base of a swept ramp. The temperature measurements were conducted with a Mach 2 test section inlet in streamwise planes perpendicular to the test section wall on which the ramp was mounted. Injection concentration measurements, conducted in cross flow planes with both Mach 2 and Mach 2.9 free stream conditions, dramatically illustrate the domination of the mixing process by streamwise vorticity generated by the ramp. These measurements, conducted using a nonintrusive optical technique (laser-induced iodine fluorescence), provide an accurate and extensive experimental data base for the validation of computation fluid dynamic codes for the calculation of highly three-dimensional supersonic combustor flow fields.

Turbulent combustion regimes for hypersonic propulsion employing hydrogen-air diffusion flames

Abstract: Astronautics and Aeronautics, AIAA, Washington, DC, 1980, pp. 8-10. Vaught, C., Witt, M., and Netzer, D., ''Investigation of SolidFuel, Dual-Mode Combustion Ramjets," Journal of Propulsion and Power, Vol. 8, No. 5, 1992, pp. 1004-1011. Zhongqin, Z., Zhenpeng, Z., Jinfu, T., and Wenlan, F., "Experimental Investigation of Combustion Efficiency of Air-Augmented Rockets," Journal of Propulsion and Power, Vol. 2, No. 4, 1986, pp. 305-310. Anil, K. N., and Damodaran, K. A., "Experimental Investigations on Flow Through a Petal Nozzle," Journal of the Institute of Engineers, India, Vol. 70, March, 1990, pp. 1-4. "Anil, K. N., and Damodaran, K. A., "Mixing of Two Axisymmetric, Coaxial, High-Speed Streams with Heat Transfer/' National Symposium on Experiments in Fluids, IIT Kanpur, India, Dec. 1992. Anil, K. N., "An Experimental Study on Mixing of Two Coaxial High Speed Streams," Ph.D. Dissertation, Aerospace Engineering, IIT Madras, India, Dec. 1992. Liepmann, H. W., and Roshko, A., Fundamentals of Gas Dynamics, Galcit Aeronautical Series, Wiley, New York, 1957.

Application of dynamical systems theory to nonlinear combustion instabilities

Abstract: Two important approximations have been incorporated in much of the work with approximate analysis of unsteady motions in combustion chambers: truncation of the series expansion to a finite number of modes, and time averaging. A major purpose of the analysis reported in this paper has been to investigate the limitations of those approximations. In particular two fundamental problems of nonlinear behavior are discussed: the conditions under which stable limit cycles of a linearly unstable system may exist; and conditions under which bifurcations of the limit cycle may occur. A continuation method is used to determine the limit cycle behavior of the equations representing the time dependent amplitudes of the longitudinal acoustic modes in a cylindrical combustion chamber. The system includes all linear processes and second-order nonlinear gas dynamics. The results presented show that time averaging works well only when the system is, in some sense, only slightly unstable. In addition, the stability boundaries predicted by the two-mode approximation are shown to be artifacts of the truncation of the system. Systems of two, four, and six modes are analyzed and show that more modes are needed to analyze more unstable systems. For the six-mode approximation with an unstable second mode two bifurcations are found to exist. A pitchfork bifurcation causes a new branch of limit cycles to exist in which the odd acoustic modes are excited. This new branch of limit cycles then undergoes a torus bifurcation that causes the system to exhibit stable quasi-periodic motions.

Axial compressor performance during surge

Abstract: There is limited information available in the literature about flow conditions in axial compressors during surge. This article presents detailed measurements from a low-speed test rig instrumented to pick up details of axial and circumferential flow disturbances. The results show that surge is initiated by rotating stall, and that the ensuing surge cycle is a sequence of well ordered cause-and-effect events. The differences in cycle behavior between "classic surge" and "deep surge" are investigated, and it is shown that the shape of the compressor characteristic determines which of these will occur. From the results it is also concluded that some important factors, such as overall pressure rise and size of hysteresis loop, have not received sufficient attention in existing techniques for predicting the rotating stall/surge boundary. In line with these findings an Appendix by E. M. Greitzer presents a more general version of the "B Parameter," which takes into account the influence of compressor design variables on the stalling behavior of the compressor.

Study of the compressible flow in a diffusing S-duct

Abstract: Benchmark aerodynamic data are presented for compressible flow through a representative S-duct configuration. Measurements of the three-dimensional velocity field, total pressures, and static pressures were obtained in five cross-sectional planes. Surface static pressure and surface flow visualization data were also acquired. All reported tests were conducted with an inlet centerline Mach number of 0.6. The Reynolds number, based on the inlet centerline velocity and duct inlet diameter, was 2.6 x 10 6. Thin inlet turbulent boundary layers existed. The collected data should be beneficial to aircraft inlet designers and are suitable for the validation of computational codes. The results show that a region of streamwise flow separation occurred within the duct. Measurements indicate that the duct curvature induced strong pressure-driven secondary flows. The crossflows evolved into counter-rotating vortices. These vortices convected low momentum fluid of the boundary layer toward the center of the duct, degrading both the uniformity and magnitude of the total pressure profile.

Effect of multidimensional flamelets in composite propellant combustion

Abstract: Studies of edge burning of oxidizer-fuel laminae were extended by inclusion of different sizes of fine ammonium perchlorate (AP) into the binder lamina to study the combustion zone microstructure in composite-propellant combustion. Different modes of burning in heterogeneous systems resulting from pressure effects, AP particle-size effects, lamina thickness, and AP/binder mix ratios were determined. Results are interpreted in terms of the effects on flame structure and multidimensional processes in the combustion zone. 13 refs.

Assumed joint probability density function approach for supersonic turbulent combustion

Abstract: In a recent experiment, Cheng et al. used UV spontaneous vibrational Raman scattering and laser-induced predissociative fluorescence techniques for simultaneous measurements of temperature and concentrations of O2, H2, H2O, OH, and N2 (and the rms of their fluctuations) in supersonic turbulent reacting shear layers. Because present computational techniques are not suited for the prediction of all of the above measurements, a new approach has been developed and is being used to predict all relevant flow properties and the rms of their fluctuations (where appropriate). The approach explores the use of a multivariate Beta PDF for concentrations. In particular, a version developed by Girimaji to model scalar mixing in turbulent flows is employed. Predictions using this model were, in general, satisfactory in regions preceding ignition, but not in regions downstream of ignition. Part of the discrepancy is a result of our current inability to relate Favre and time averages. 10 refs.

Effect of film cooling/regenerative cooling on scramjet engine performances

Abstract: Film cooling was modeled to allow performance prediction of scramjet engine design. The model was based on experimental results of the compressible mixing layer for the vicinity of the injection slot, and on analytical results of the turbulent boundary layer in the region far from the slot. The film cooling model was integrated with a quasi-one-dimensional scramjet performance prediction model. In an engine employing a combination of film cooling and regenerative cooling, coolant flow rate of the engine slightly exceeded the stoichiometric flow rate, even at high flight Mach numbers, and had the best specific impulse and system pressure performances. These advantages were achieved by an increase in the volume flow rate and a decrease in the velocity difference between the main flow and the coolant, both due to an increase in the film coolant temperature. The effective cooling system with a combination of film cooling and regenerative cooling was also advantageous with regard to avoidance of excess cooling of the engine wall.

Review of shock-induced supersonic combustion research and hypersonic applications

Abstract: This article focuses on research in supersonic combustion and combustion kinetics in high-speed flow between 1959-1968, and the application of the experimental results to hypersonic propulsion. The analysis discusses both advantages and problems for premixing the fuel and employing shock-induced combustion as an ignition method for a scramjet flying at a high Mach number. The experimental tests are discussed, including implications to the chemical kinetics of the high-velocity combustion process. The conditions were confined to relatively low pressure, less than 2 atm (200 kPa). The results were considered to be mainly applicable for high-altitude scramjet flight, at low static pressure, where chemical reaction distances will be long. At these lower pressures, "shock-induced combustion" may be the predominant effect in a scramjet application, and it has some advantages that are discussed. The relation between shock-induced combustion and "detonation" is also discussed. In addition, an attempt is made to resolve the conflicting experimental data published in the 1960s relating to "standing detonation waves" and shock-induced combustion.

Vane-blade interaction in a transonic turbine. I - Aerodynamics

Abstract: Part I of this article presents results of a computational investigation of the effects of blade row interaction on the aerodynamics of a transonic turbine stage. The predictions are obtained using a two-dimensional unsteady Navier-Stokes code based on an explicit Runge-Kutta algorithm and an overlapping O-H grid system. This code simulates the flow in time-accurate fashion using nonreflective stage inflow and outflow boundary conditions and phase-lagging procedures for modeling arbitrary airfoil counts in the vane and blade rows. The O-H grid provides high spatial resolution of the high gradient regions near the airfoil surfaces and allows for arbitrary placement of stage inflow and outflow boundaries. Unsteady and time-averaged airfoil surface pressure predictions are compared with those from an older version of the code based on the explicit hopscotch algorithm and an O-grid system, and experimental data obtained in a short-duration shock tunnel facility.

Theoretical study of sensor-actuator schemes for rotating stall control

Abstract: A theoretical study has been conducted to determine the influence of actuator and sensor choice on active control of rotating stall in axial-flow compressors. The sensors are used to detect small amplitude traveling waves that have been observed at the inception of rotating stall in several different compressors. Control is achieved by feeding the sensed quantity back to the actuator with a suitable gain and spatial phase shift relative to the measured wave. Actuators using circumferential arrays of jets, intake ports, and movable inlet guide vanes upstream of the compressor, and valves downstream of the compressor were considered. The effect of axial velocity, static pressure, or total pressure measurement on control effectiveness was investigated. In addition, the influence of the actuator bandwidth on the performance of the controlled system was determined. The results of the study indicate that the potential for active control of rotating stall is greater than that achieved thus far with movable inlet guide vanes. Furthermore, axial velocity sensing was most effective. Actuator bandwidth affected the performance of the controlled compressors significantly, but certain actuators were affected less severely than others.

Direct simulation monte carlo model of low reynolds number nozzle flows

Abstract: A numerical analysis of low Reynolds number nozzle flows is performed to investigate the loss mechanisms involved and to determine the nozzle wall contour that minimizes these losses. The direct simulation Monte Carlo method is used to simulate nitrogen flows through conical, trumpet-shaped, and bell-shaped nozzles at inlet stagnation temperatures of 300 and 1000 K. The Reynolds number of the flows based on throat diameter range from 90 to 125. The trumpet-shaped nozzle has the highest efficiency with the unheated flow. With the heated flow both the trumpet and bell-shaped nozzles have a 6.5% higher efficiency than the conical nozzle. The conical nozzle has the highest discharge coefficient, which is unaffected by the change in stagnation temperature; however, the increase in stagnation temperature increases the heat-transfer and viscous losses in the boundary layer. These results suggest that the trumpet-shaped wall contour performs most efficiently except near the throat region, where it incurs large viscous losses. However, the bell-shaped nozzle may increase its overall performance with an increase in stagnation temperature.

Motor and plume particle size measurements in solid propellant micromotors

Abstract: Particle size distributions were measured in the chamber, nozzle, and plume of a subscale solid propellant rocket motor. A significant reduction in the mean size of the aluminum/aluminum oxide particles occurred within the motor chamber. The mass fraction of small particles (<2 ft) at the nozzle entrance was less than 10%. Also, most particles were smaller than 50 ft, although a few as large as 85 ft were present. In the converging and throat portions of the nozzle it appeared that particle breakup dominated over collision coalescence. Collision coalescence was observed to occur more dominantly in the supersonic nozzle flow. At the nozzle exit the particle mean size was usually less than 2.5 ft. The particle size distributions were bimodal or trimodal, with the larger particles concentrated near the plume centerline. Because of the short nozzle residence times it is not known whether or not these results are also applicable to full-scale motors. The mean size of the small A12O3 particles in the plume edges was less than 0.5 ft in diam, with an index of refraction of 1.64 ± 0.04 (apparently y-Al2O3), independent of propellant composition, motor operating conditions and nozzle geometry.

Damage-Mitigating Control of a Reusable Rocket Engine

Abstract: This article presents a control concept for damage prediction and damage mitigation in reusable rocket engines for enhancement of structural durability The key idea here is to achieve high performance without overstraining the mechanical structures so that 1) the functional lives of critical components are increased, resulting in enhanced safety, operational reliability, and availability; and 2) the plant (ie, the rocket engine) can be inexpensively maintained, and safely and efficiently driven under different operating conditions To this effect, dynamics of fatigue damage have been modeled in the continuous-time setting instead of the conventional cycle-based approach, and an optimal control policy is formulated by constraining the accumulated damage and its time derivative Efficacy of the proposed damage mitigation concept is evaluated for life extension of the turbine blades of a bipropellant rocket engine via simulation experiments The simulation results demonstrate the potential of increasing the structural durability of reusable rocket engines with no significant loss of performance

Facility opportunities and associated stream chemistry considerations for hypersonic air-breathing propulsion

Abstract: Wallace Chinitz is a Principal Scientist at General Applied Science Laboratories (GASL) as well as a Professor of Mechanical Engineering at Cooper Union's Nerken School of Engineering. He received his Ph.D. from the Polytechnic Institute of Brooklyn in 1962, and taught at that institution before leaving to conduct research at GASL until 1972. Dr. Chinitz began his career at the Fairchild Engine Division in 1957, then worked at Republic Aviation Corporation until 1960. He has published over 100 articles, papers, and reports. Dr. Chinitz is an Associate Fellow of the AIAA and is a member of the ASME, the ASEE, and the Combustion Institute.

Adaptive modeling of jet engine performance with application to condition monitoring

Abstract: A method of simulation of the performance of jet engines, with the possibility of adapting to engine particularities, is presented. It employs an adaptation procedure coupled to a performance model solving the component matching problem. The proposed method can provide accurate simulation for engines of the same type, with differences that are due to manufacturing or assembly tolerances. It does not require accurate component maps, because they are derived during the adaptation procedure. It can also be used for health monitoring purposes, for component fault identification, and condition assessment. The effectiveness of the proposed method is demonstrated by application to two commercial jet engines.

Vane-blade interaction in a transonic turbine. II - Heat transfer

Abstract: Part II of this article presents results of a combined computational/experimental investigation into the effects of stator-rotor interaction on the heat transfer distributions on the vane and blade of a transonic turbine stage The predictions were obtained using a two-dimensional unsteady Navier-Stokes code described in Part I of this article, and the measurements were acquired in a short-duration shock tunnel facility Twenty miniature thinfilm heat flux button gauges were mounted at the midspan of the vane and blade, and contoured inserts containing many thin-film gauges were used on the blade leading edge to spatially resolve the heat transfer rates in that high-gradient region A grid refinement study was performed with steady noninteractive solutions to ascertain the minimum grid size needed to obtain grid-independent solutions Predicted time-averaged and phase-resolved heat transfer rates are compared with measurements on the vane and blade

Computation of shock-induced combustion using a detailed methane-air mechanism

Abstract: The shock-induced combustion of methane-air mixtures in hypersonic flows is investigated using a new reaction mechanism consisting of 19 reacting species and 52 elementary reactions. This reduced model is derived from a full kinetic mechanism via the detailed reduction technique. Zero-dimensional computations of several shocktube experiments are presented first. The computed values for ignition delay and flame speed are in close agreement with experimental data and with results obtained using a full mechanism. The new reaction mechanism is then combined with a fully implicit Navier-Stokes CFD code to simulate two-dimensional and axisymmetric shock-induced combustion experiments of stoichiometric methane-air mixtures at a Mach number of M = 6.61. Good agreement with the experiments is obtained. Furthermore, it is shown that previous calculations were unable to accurately predict this type of flow due to their use of severely limited reduced chemical mechanisms. Finally, applications to the ram accelerator concept are also presented, based on a novel double-ramp configuration.

Velocity characteristics of reacting and nonreacting flows in a dump combustor

Abstract: A two-component laser Doppler anemometer (LDA) was used to measure mean and turbulent velocities in an axisymmetric dump combustor. Data were acquired for the axial and tangential components and provide a comparison between combusting and isothermal flows. Inlet profiles were carefully documented to provide a suitable boundary condition for future comparison to predictive models. The results show significant differences between the reacting and nonreacting flows. One large-scale observable effect was the difference between the recirculation regions for each case. Combustion decreased the length of the region by approximately 50%, while increasing the maximum velocities. This made for a more compact, but stronger, recirculation region. The reduction in recirculation zone size is recommended as a benchmark criterion in modeling studies. The presence of reaction caused higher turbulent fluctuations near the expansion, and lower fluctuations in downstream regions. The turbulence was observed to be nonisotropic in both the reacting and nonreacting cases.

Hysteresis and bristle stiffening effects in brush seals

Abstract: Extensive testing of conventional brush seals has identified the phenomena of bristle `hysteresis` and `stiffening` with pressure as their two major drawbacks. Subsequent to any differential movement of the runner into the bristle pack due to its radial excursions or centrifugal/thermal growths, the displaced bristles do not recover against the frictional forces between them and the backing plate. As a result, a significant leakage increase is observed following any runner movement. Furthermore, the bristle pack exhibits a considerable stiffening effect with the application of pressure. This phenomenon may adversely affect the life of the seal and the runner due to a highly increased mechanical contact pressure at the sliding interface. In comparison with these conventional design seals, the characteristics of an improved design, known as the `low hysteresis` design, are presented here. This design shows a substantially lower degree of the detrimental effects mentioned above. This type of seal can maintain its reduced leakage characteristics throughout the running cycle with runner excursions and growths. The bristles also do not show any stiffening, up to a certain pressure threshold. Therefore, this seal also has a potential for a longer life than a brush seal of conventional design. 5 refs.

Low-noise, high-strength, spiral-bevel gears for helicopter transmissions

Abstract: Improvements in spiral-bevel gear design were investigated to support the Army/NASA Advanced Rotorcraft Transmission program. Program objectives were to reduce weight by 25 percent, reduce noise by 10 dB, and increase life to 5000 hr mean-time-between-removal. To help meet these goals, advanced-design spiral-bevel gears were tested in an OH-58D helicopter transmission using the NASA 500-hp Helicopter Transmission Test Stand. Three different gear designs tested included: (1) the current design of the OH-58D transmission except gear material X-53 instead of AISI 9310; (2) a higher-strength design the same as the current but with a full fillet radius to reduce gear tooth bending stress (and thus, weight); and (3) a lower-noise design the same as the high-strength but with modified tooth geometry to reduce transmission error and noise. Noise, vibration, and tooth strain tests were performed and significant gear stress and noise reductions were achieved.

Experimental Study of a Model Gas Turbine Combustor Swirl Cup, Part 11: Droplet Dynamics

Abstract: As the second part of a study to characterize the performance of a 3 x scale gas turbine combustor swirl cup, the focus of the present article addresses the droplet dynamics. Droplet axial, radial, and tangential velocities as well as size were measured using phase Doppler interferometry. Bimodal droplet axial and tangential velocities distributions were observed in the shear layer close to the exit plane of the swirl cup. Bimodal droplet radial velocity distributions around the closing point of the on-axis recirculation zone and at the periphery of the spray were observed as well. Droplet velocity histograms, based on droplet size classes, reveal how these velocity bimodal distributions are formed and explain why velocity fluctuations for larger droplets are greater than those of the smaller droplets in some regions. Correlations between size and velocity, individual velocity components, and size and flow angle provide evidence of intermittent flowfield structures superimposed on the local and global turbulence. Specific flowfield structures and flowfield intermittencies are found to give rise to these bimodal distributions. Overall, these data add measurably to the understanding of the two-phase transport that can occur in practical systems. 2 refs.

Simultaneous velocity and temperature measurements in a premixed dump combustor

Abstract: Experimental measurements of velocity and temperature were made in a low-speed turbulent flowfield following an axisymmetric sudden expansion with and without combustion. The combustion case considered here used a lean, completely premixed propane-air mixture whose flame was stabilized by the recirculation zone present in this flow. Simultaneous two-component laser velocimeter measurements were made giving the mean axial and radial velocities and the Reynolds stresses throughout the flowfield. In addition, simultaneous timeresolved temperature measurements were made in the reacting flow using a fast response thermocouple. Velocitytemperature correlations were formed from the velocity and temperature measurements. The reacting flow case was found to have a shorter and stronger recirculation zone, and the turbulence intensity level was found to be suppressed over most of the flowfield when compared with the cold flow case.

Coupling between vorticity and pressure oscillations in combustion instability

Abstract: The operation of premixed dump combustors is hindered by large-amplitude, low-frequency oscillations leading to flame flashback. The genesis of this instability is the subject of this article. A physical model that accounts for the relevant components of flow-combustion interactions in this system is formulated assuming that the combustor is an acoustically compact trough, the flow is two dimensional, the flame is a thin front, the inlet section is charged by a constant-pressure reservoir, and the exit pressure is forced. Numerical solutions at high Reynolds number are obtained using the vortex method. The nonreacting flow exhibits coherent oscillations that scale with the trough depth and inlet velocity to a Strouhal number of O(0.1). Simulations of the reacting flow show that vorticity-flame-acoustic coupling is the driving mechanism for the observed instability. In this work, the exit pressure is modulated near the frequency of the natural mode, and the amplitude of oscillation is observed to grow with increasing heat release due to phase-matched coupling between flow and heat release oscillations, in accordance with the Rayleigh criterion. This amplification ultimately leads to flow reversal and the propagation of the flame into the inlet channel. 35 refs.

Experimental Study of a Model Gas Turbine Combustor Swirl Cup, Part I: Two-Phase Characterization

Abstract: The behavior of droplets and the continuous phase (i.e., gas in the presence of the droplets) downstream of a 3 x model gas turbine combustor dome swirl cup is characterized via phase Doppler interferometry in the absence of reaction. The goal is to improve the understanding of droplet-gas interaction in a complex flow typical of that produced by engine hardware. Three components of continuous phase and droplet velocities were measured along with droplet size. The measurements reveal that (1) at the exit plane of the swirl cup, more uniform and finer droplets are produced relative to the atomizer alone, (2) both the continuous phase and droplets recirculate, (3) the region downstream of the swirl cup into which droplets join the recirculation is correlated with droplet size, and (4) significant slip velocities exist between the continuous phase and the droplets which are also correlated with droplet size and reflect a strong momentum exchange between the phases. 11 refs.