Journal of Propulsion and Power

Active control is one way to cope with surge in compression systems. For this, a suitable combination of actuators and sensors must be selected. In this paper, the goal is to find a combination fulfilling the condition that a controller exists, which stabilizes the nominal model in the face of disturbances and actuator limitations. For a collection of linearizations of the nonlinear system, this condition is quantified as an ℋ_∞ norm bound on the closed-loops. Among the proposed actuators and sensors, the close-coupled valve and mass flow sensor are the most promising for the considered system, but a movable wall should be added to meet the specifications

The Mars atmosphere contains a significant load of suspended dust. Settling of atmospheric dust onto the surface of the solar array is potentially a lifetime limiting factor for a power system on any Mars mission. For long-term operation of arrays on Mars, it may be necessary to develop techniques to remove deposited dust. Dust is expected to adhere to the array by Van der Waals adhesive forces. These forces are quite strong at the dust particle sizes expected. If the array surface is insulating, it is possible that they may also be subject to electrostatic adhesion, which may be extremely strong. Dust removal methods must overcome this force. Dust removal methods can be categorized briefly into four categories: natural, mechanical, electromechanical, and electrostatic. The environment of Mars is expected to be an ideal one for use of electrostatic dust removal techniques

A steady incompressible three-dimensional viscous flow analysis has been conducted for the Space Shuttle external tank/orbiter propellant feed line disconnect flapper valves with upstream elbows. The Navier-Stokes code, INS3D, is modified to handle interior obstacles and a simple turbulence model. The flow solver is tested for stability and convergence in the presence of interior flappers. An under-relaxation scheme has been incorporated to improve the solution stability. Important flow characteristics such as secondary flows, recirculation, vortex and wake regions, and separated flows are observed. Computed values for forces, moments, and pressure drop are in satisfactory agreement with water flow test data covering a maximum tube Reynolds number of 3.5 million. The predicted hydrodynamical stability of the flappers correlates well with the measurements.

A two-dimensional (2-D) computer code was developed for modeling enclosed volumes of gas with oscillating boundaries, such as Stirling machine components. An existing 2-D incompressible flow computer code, CAST, was used as the starting point for the project. CAST was modified to use the compressible non-acoustic Navier-Stokes equations to model an enclosed volume including an oscillating piston. The devices modeled have low Mach numbers and are sufficiently small that the time required for acoustics to propagate across them is negligible. Therefore, acoustics were excluded to enable more time efficient computation. Background information about the project is presented. The compressible non-acoustic flow assumptions are discussed. The governing equations used in the model are presented in transport equation format. A brief description is given of the numerical methods used. Comparisons of code predictions with experimental data are then discussed.

The transition behavior of a free-shear layer above a cavity with high and low levels of freestream acoustic disturbances has been investigated at Mach 3.5. Optical, mean pitot pressure, and hot-wire techniques were used to detect transition locations. The transition Reynolds numbers obtained were between ReTr = 363,000 and 530,000, in good agreement with previous (conventional, high noise facility) results. The study indicated that the effects of lowering freestream noise on ReTr were minimal but, as expected, favorable (higher ReTr). However, hot-wire anemometry indicated that the shear layer disturbance growth differed considerably for high and low noise levels. Also, the growth frequencies are considerably below expectations from linear stability theory. Correlations between hot-film sensors mounted on the cavity floor indicated the presence of disturbances, believed to be vorticity and/or entropy waves, convecting upstream at about 7.2-11.5% of the freestream velocity. Upstream convected disturbances (acoustic as well as the vorticity/entropy wave indicated herein) are believed to be at least partially responsible for the insensitivity of ReTr to the freestream acoustic disturbance field.

A unique high temperature modal test and model correlation/update program has been performed on the composite nozzle of the FASTRAC engine for the NASA X34 Reusable Launch Vehicle. The program was required to provide an accurate high temperature model of nozzle for incorporation into the engine system structural dynamics model for loads calculation; this model is significantly different from the ambient case due to the large decrease in composite material properties due to heating. The modal test was enabled by piggybacking onto a hot-fire test of the nozzle at NASA/Marshall and tracking the natural frequencies as they decreased. A series of high fidelity modal tests and finite element model correlation of the nozzle in a free-free configuration was initially performed. This model was then attached to a modal-test verified model of the engine hot-fire test stand and the' ambient system mode shapes identified. A reduced set of accelerometers was then attached to the nozzle, the engine fired full-duration, and the frequency peaks corresponding to the ambient nozzle modes individually isolated and tracked. To update the finite element model of the nozzle to these frequency curves, a multiplicative factor was. applied to the rate of decline of the composite material property versus temperature table. This new property table was used to create high-temperature nozzle models corresponding to 10 second engine operation increments and tied into the engine system model for loads determination.

Spur gear endurance tests and rolling-element surface tests were conducted to investigate vacuum-induction-melted, vacuum-arc-melted (VIM-VAR) M50NiL steel for use as a gear steel in advanced aircraft applications, to determine its endurance characteristics, and to compare the results with those for standard VAR and VIM-VAR AISI 9310 gear material. Tests were conducted with spur gears and rolling-contact bars manufactured from VIM-VAR M50NiL and VAR and VIM-VAR AISI 9310. The gear pitch diameter was 8.9 cm (3.5 in.). Gear test conditions were an inlet oil temperature of 320 K (116 F), and outlet oil temperature of 350 K (170 F), a maximum Hertz stress of 1.71 GPa (248 ksi), and a speed of 10,000 rpm. Bench rolling-element fatigue tests were conducted at ambient temperatures with a bar speed of 12,500 rpm and a maximum Hertz stress of 4.83 GPA (700 ksi). The VIM-VAR M50NiL gears had a surface fatigue life that was 4.5 and 11.5 times that for VIM-VAR and VAR AISI 9310 gears, respectively. The surface fatigue life of the VIM-VAR M50NiL rolling-contact bars was 13.2 and 21.6 times that for the VIM-VAR and VAR AISI 9310, respectively. The VIM-VAR M50NiL material was shown to have good resistance to fracture through a fatigue spall and to have fatigue life far superior to that of both VIM-VAR and VAR AISI 9310 gears and rolling-contact bars.

LASER ablation propulsion (LAP) is a major new electric propulsion concept with a 35-year history. In LAP, an intense laser beam [pulsed or continuous wave (CW)] strikes a condensedmatter surface (solid or liquid) and produces a jet of vapor or plasma. Just as in a chemical rocket, thrust is produced by the resulting reaction force on the surface. Spacecraft and other objects can be propelled in this way. In some circumstances, there are advantages for this technique compared with other chemical and electric propulsion schemes. It is difficult to make a performance metric for LAP, because only a few of its applications are beyond the research phase and because it can be applied in widely different circumstances that would require entirely different metrics. These applications range from milliwatt-average-power satellite attitude-correction thrusters through kilowatt-average-power systems for reentering near-Earth space debris and megawatt-to-gigawatt systems for direct launch to lowEarth orbit (LEO). We assume an electric laser rather than a gas-dynamic or chemical laser driving the ablation, to emphasize the performance as an electric thruster. How is it possible for moderate laser electrical efficiency to givevery high electrical efficiency? Because laser energy can be used to drive an exothermic reaction in the target material controlled by the laser input, and electrical efficiency only measures the ratio of exhaust power to electrical power. This distinction may seem artificial, but electrical efficiency is a key parameter for space applications, in which electrical power is at a premium. The laser system involved in LAP may be remote from the propelled object (on another spacecraft or planet-based), for example, in laser-induced space-debris reentry or payload launch to low planetary orbit. In other applications (e.g., the laser–plasma microthruster that we will describe), a lightweight laser is part of the propulsion engine onboard the spacecraft.

Aspects are examined of an investigation by the Marshall Space Flight Center into the conceptual feasibility of increasing the thrust performance of the Space Shuttle Main Engine (SSME) by using a conical nozzle extension fitted with an ablative insert in order to achieve a low-cost, near-term gain in payload. The ablating insert would provide a controlled increase in nozzle expansion ratio during launch and early climbout (first 30-60 seconds) so as to reduce thrust loss from nozzle over-expansion in the lower atmosphere. Summaries are given of JPL studies in the area of: defining the near-wall flow environment in the extended nozzle insert region; selecting potential insert materials; conceptualizing an extension/insert geometrical configuration; and identifying future experimental efforts necessary to verify the feasibility of the concepts.

Finite-element theory is used to calculate the acoustic field of a propeller in a soft-walled circular wind tunnel and to compare the radiation patterns to the same propeller in free space. Parametric solutions are presented for a Gutin propeller for a variety of flow Mach numbers, admittance values at the wall, microphone position locations, and propeller-to-duct radius ratios. Wind-tunnel boundary layer is not included in this analysis. For waU admittance nearly equal to the characteristic value of free space, the free-field and ducted propeller models agree in pressure level and directionality. In addition, the need for experimentally mapping the acoustic field is discussed. © 1987 American Institute of Aeronautics and Astronautics, Inc.

Experimental investigations have been carried out to examine film cooling effectiveness of an accelerated hot gas in a subscale rocket combustion chamber. In support of future first-stage high-performance rocket combustion chambers, a Vulcain2-like test case has been examined with combustion pressure levels up to 12 MPa. The effectiveness of an almost tagentially injected film of hydrogen with an initial temperature of approximately 280 K has been determined. Axial distributions of temperature were measured inside the copper liner as well as on the chamber surface in the convergent and divergent parts of the nozzle segment. An existing film cooling model has been modified for application in a combined convective and filmcooled combustion chamber with an accelerated hot gas. The new model predicts film cooling effectiveness at different combustion-chamber pressures and film blowing rates at sub-, trans-, and supersonic conditions.

The ram accelerator is a ramjet-in-tube device which has been demonstrated at velocities up to 2.5 km/sec, has potential for operation up to approximately 10 km/sec and could be used for direct space launch or large ballistic ranges. The ends of the main ram accelerator tube must have end closures which support substantial pressure differences. There are potentially serious difficulties using solid end closures such as diaphragms pierced by the projectile or explosively removed end closures or fast acting valves. These include risks of significant damage to the projectile and launch tube and the wasting of tube length. A new end closure system which uses the momentum of a annular axial air jet to support the required pressure differences is described. This system avoids the difficulties of the solid end closure system at the cost of some increase in overall launch system complexity. A preliminary design of such an air jet end closure is presented and it is concluded that the requirements for air flow rates and storage are reasonable and would likely add only a modest increase to the overall cost of the launch system. If the difficulties with solid end closures prove to be significant, the air jet end closure system may offer a solution.

The theoretical performance of diagonal conducting wall crossed-field accelerators is examined on the basis of an infinite segmentation assumption using a cross-plane averaged generalized Ohm s law for a partially ionized gas, including ion slip. The desired accelerator performance relationships are derived from the cross-plane averaged Ohm s law by imposing appropriate configuration and loading constraints. A current-dependent effective voltage drop model is also incorporated to account for cold-wall boundary layer effects, including gasdynamic variations, discharge constriction, and electrode falls. Definition of dimensionless electric fields and current densities leads to the construction of graphical performance diagrams, which further illuminate the rudimentary behavior of crossed-field accelerator operation.

Pressure fluctuations measured in turbine engine combustors at low engine speed show good agreement with theory. Above idle speed, the turbine chokes and a significant change in the shape of the measured combustor pressure spectrum is observed. A theoretical model of the acoustic pressure generated in the combustor due to the turbulence/flame front interaction did not account for acoustic waves reflected from the turbine. By retaining this combustion noise source model and adding a reflecting plane at the turbine and combustor inlet, a theoretical model has been developed that reproduces the undulations in the combustor fluctuating pressure spectra. Plots of the theoretical combustor fluctuating pressure spectra are compared to the measured pressure spectra obtained from the CF6-50 turbofan engine over a range of engine operating speeds. Reasonable agreement exists. It is thus concluded that the simplified combustion noise theory, when modified by a simple turbine reflecting plane, adequately accounts for the changes in measured combustor pressure spectra.

The paper reports a detailed study on the acoustic control of free jet mixing at realistic Reynolds numbers. The experimental results were obtained at Mach numbers of 0.3 and 0.8 and flow total temperatures up to 800 K. The Reynolds numbers ranged from 350,000 to 1,300,000. Excitation Strouhal numbers were in the range of 0.2 to 0.6. The experimental results are compared with predictions, based on an extension of the analysis by Tam and Morris. The results showed that proper upstream tone excitation enhances mixing of unheated jets for both high-speed, high Reynolds number conditions and for low-speed conditions. The heated jet, however, shows a response to upstream excitation that depends on jet Mach number. The agreement between the predictions and the experiments is very good for unheated jet conditions. However, for jets heated to temperatures above 600 K, the theoretical predictions differ from the experimental results.

The analytical solution for the eigenmodes of a cylindrical combustor allows for several realizations of the first tangential mode. The limiting cases of standing and spinning first tangential modes are well known. The general solution for a first tangential mode, however, allows for a large variety of realizations. In this study the analytical solution of the acoustic problem and a numerical time-resolved simulation of the acoustic pressure field are compared with experimental data from hot-fire tests. The experimentally observed pressure data are in excellent agreement with the simulation for all possible theoretical solutions. The approach enables detailed insight into the dynamics of the pressure field and thus constitutes a tool for the characterization of acoustic pressure waves in a cylindrical combustor even when the amplitude of the investigated eigenmode is small as compared with the mean pressure fluctuation.

Efficient technique are studied for accomplishing propellant resettling through the minimization of propellant usage through impulsive thrust. A comparison between the use of constant-thrust and impulsive-thrust accelerations for the activation of propellant resettlement shows that impulsive thrust is superior to constant thrust for liquid reorientation in a reduced-gravity environment. This study shows that when impulsive thrust with 0.1-1.0-, and 10-Hz frequencies for liquid-fill levels in the range between 30-80 percent is considered, the selection of 1.0-Hz-frequency impulsive thrust over the other frequency ranges of impulsive thrust is the optimum. Characteristics of the slosh waves excited during the course of 1.0-Hz-frequency impulsive-thrust liquid reorientation were also analyzed.

Computer models of rocket engines and single-stage-to-orbit vehicles that were developed previously and recently combined have now been extended to include variable-mixture-ratio engines and two-stage vehicles. The variable-mixture-ratio engine is of interest because it offers the opportunity for reduced vehicle dry mass relative to single-mixture-ratio engines without the operational complexity of two fuels or two engine designs. Two-stage vehicles are of interest because the next generation of advanced launch vehicles probably will not use a sufficiently high level of technology to allow a single-stage vehicle to be a reasonable choice from the standpoint of technical risk and economics. Results are presented for vertical-takeoff, horizontal-landing, winged, manned, fully reusable vehicles with a payload of 13.6 Mg (30,000 lb) to low Earth orbit. Both single-stage and two-stage vehicles are included. Hydrogen, methane, and propane engines were studied with a staged-combustion cycle. Optimizations included the nozzle exit pressure, thrust split between the booster and orbiter, chamber pressure, and mixture ratio.

An efficient 3-D hybrid scheme is applied for solving Euler equations to analyze advanced propellers. The scheme treats the spanwise direction semi-explicitly and the other two directions implicitly, without affecting the accuracy, as compared to a fully implicit scheme. This leads to a reduction in computer time and memory requirement. The calculated power coefficients for two advanced propellers, SR3 and SR7L, and various advanced ratios showed good correlation with experiment. Spanwise distribution of elemental power coefficient and steady pressure coefficient differences also showed good agreement with experiment. A study of the effect of structural flexibility on the performance of the advanced propellers showed that structural deformation due to centrifugal and aero loading should be included for better correlation.

Several nozzle concepts that promise a gain in performance over existing conventional nozzles are discussed in this paper. It is shown that significant performance gains result from the adaptation of the exhaust flow to the ambient pressure. Special attention is then given to altitude-adaptive nozzle concepts, which have recently received new interest in the space industry. Current research results are presented for dual-bell nozzles and other nozzles with devices for forced flow separation and for plug nozzles with external freestream expansion. In addition, results of former research on nozzles of dual-mode engines such as dual-throat and dual-expander engines and on expansion-deflection nozzles are shown. In general, flow adaptation induces shocks and expansion waves, which result in exit profiles that are quite different from idealized one-dimensional assumptions. Flow phenomena observed in experiments and numerical simulations during different nozzle operations are highlighted, critical design aspects and operation conditions are discussed, and performance characteristics of selected nozzles are presented. The consideration of derived performance characteristics in launcher and trajectory optimization calculations reveal significant payload gains at least for some of these advanced nozzle concepts.

New control concepts for the next generation of advanced air-breathing and rocket engines and hypersonic combined-cycle propulsion systems are analyzed. The analysis provides a database on the instrumentation technologies for advanced control systems and cross matches the available technologies for each type of engine to the control needs and applications of the other two types of engines. Measurement technologies that are considered to be ready for implementation include optical surface temperature sensors, an isotope wear detector, a brushless torquemeter, a fiberoptic deflectometer, an optical absorption leak detector, the nonintrusive speed sensor, and an ultrasonic triducer. It is concluded that all 30 advanced instrumentation technologies considered can be recommended for further development to meet need of the next generation of jet-, rocket-, and hypersonic-engine control systems.

Several nozzle concepts that promise a gain in performance over existing conventional nozzles are discussed in this paper, It is shown that significant performance gains result from the adaptation of the exhaust now to the ambient pressure. Special attention is then given to attitude-adaptive nozzle concepts, which have recently received new interest in the space industry, Current research results are presented for dual-bell nozzles and other nozzles with devices for forced flow separation and for plug nozzles with external freestream expansion. In addition, results of former research on nozzles of dual-mode engines such as dual-throat and dual-expander engines and on expansion-deflection nozzles are shown. In general, now adaptation induces shocks and expansion waves, which result in exit profiles that are quite different from idealized one-dimensional assumptions. Flow phenomena observed in experiments and numerical simulations during different nozzle operations are highlighted, critical design aspects and operation conditions are discussed, and performance characteristics of selected nozzles are presented. The consideration of derived performance characteristics in launcher and trajectory optimization calculations reveal significant payload gains at least for some of these advanced nozzle concepts.

An advanced rocket thrust chamber for future space applications is described along with an improved method of fabrication. Included are fabrication demonstrator and test chambers produced by this method. This concept offers the promise of improved cyclic life, reusability, reliability, and performance. The performance is improved because of the enhanced enthalpy extraction. The life, reusability, and reliability is improved because of the enhanced structural compliance inherent in the construction. The method of construction involves the forming of the combustion chamber by a tube-bundle of high-conductivity copper or copper alloy tubes, and the bonding of these tubes by a unique electroforming operation. Further, the method of fabrication reduces chamber complexity by incorporating manifolds, jackets, and structural stiffeners while having the potential for thrust chamber cost and weight reduction.

Current interest in advanced turboprop blade designs and new propeller installation schemes motivated an effort to include unsteady loading effects in an existing propeller prediction computer program. The present work validates the prediction capability, while studying the effects of shaft inclination on the radiated sound field. Classical methods of propeller performance analysis supply the time-dependent blade loading needed to calculate noise. Polar plots of the sound pressure level (SPL) of the first four harmonics and to overall SPL are indicative of the change in the directivity pattern as a function of the propeller angle of attack. Noise predictions are compared with newly available wind tunnel data and the accuracy and applicability of the prediction method are discussed. It is concluded that unsteady blade loading caused by inclining the propeller with respect to the flow changes the directionality and intensity of the radiated noise. These changes are well modeled by the present quasisteady prediction method.

A time-marching Navier-Stokes code called PARC3D was used to study the three-dimensional viscous flow associated with an advanced ducted propeller (ADP) subsonic inlet at takeoff operating conditions. At a freestream Mach number of 0.2, experimental data of the ADP inlet model indicated that the flow on the cowl windward lip remained attached at an angle of attack of 25 deg became unstable at 29 deg and separated at 30 deg. An experimental through-flow test study (without propeller) using another, but similar inlet indicated that the separation occurred at an angle of attack a few degrees below the value observed when the inlet was tested with the propeller. This evidence showed that the propeller exerted a favorable effect on the inlet low speed and high angle-of-attack performance. A stationary blockage device was designed and tested with this latter inlet to simulate the propeller effect. The test data showed that the blockage device helped prevent the inlet flow from being separated at angle of attack a few degrees higher than the inlet through-flow but was 1 deg lower than the inlet with the propeller. In the present numerical study this flow blockage was modeled via a PARC3D computational boundary condition (BC) called the screen BC. The principle formulation of this BC was based on the 'one-and-half dimensions' actuator disk theory. This screen BC was prescribed at the inlet propeller face station. Numerical results were obtained for inlet flow calculations with and without the screen BC. The results with the screen BC compared better with the ADP experimental test data than those obtained without the screen BC, particularly when the inlet flow separated.

The objectives of the NASA/Army-sponsored Small Engine Component Technology (SECT) study were to identify high-payoff technologies for year-2000 small gas turbine applications and to formulate the required technology programs. The projected technologies were evaluated in terms of their influence on operating cost for the four candidate applications: rotorcraft, commuter, cruise missile, and auxiliary power unit (APU). This paper reviews the reference missions, engines, and aircraft, and year-2000 technology projections, cycle studies, advanced engine selections, and technology evaluations. Conventional simple cycles and heat-recovery cycles were both evaluated.

Two Euler analyses techniques, finite difference and finite volume, were employed to predict the blade surface pressure distributions of a large scale advanced propeller. The predicted pressure distributions were compared with wind tunnel data. Both techniques produced blade pressure distributions that are in fairly good agreement with the data over the range of test Mach numbers of 0.2 to 0.78. However, the numerical simulations fail to predict correctly the measured pressure distributions for the low Mach number, high power case. The data indicate the presence of a leading edge vortex for this case. A discussion of the compressibility effects is also presented.

Cycle studies have shown the benefits of increasing engine pressure ratios and cycle temperatures to decrease engine weight and improve performance of commercial turbine engines. NASA is working with industry to define technology requirements of advanced engines and engine technology to meet the goals of NASA's Advanced Subsonic Technology Initiative. As engine operating conditions become more severe and customers demand lower operating costs, NASA and engine manufacturers are investigating methods of improving engine efficiency and reducing operating costs. A number of new technologies are being examined that will allow next generation engines to operate at higher pressures and temperatures. Improving seal performance - reducing leakage and increasing service life while operating under more demanding conditions - will play an important role in meeting overall program goals of reducing specific fuel consumption and ultimately reducing direct operating costs. This paper provides an overview of the Advanced Subsonic Technology program goals, discusses the motivation for advanced seal development, and highlights seal technology requirements to meet future engine performance goals.

An unsteady, three-dimensional Euler analysis technique was used to compute the flowfield of an advanced propeller operating at an angle of attack. The predicted blade-pressure waveforms for an angular inflow of 3 deg were compared with wind-tunnel data at two Mach numbers, 0.5 and 0.2. For an inflow Mach number of 0.5, the predicted pressure response is in fair agreement with data: the predicted phases of the waveforms are in close agreement with data, whereas the magnitudes are underpredicted. At the low Mach number of 0.2 (takeoff), the numerical solution shows the formation of a leading-edge vortex that is in qualitative agreement with measurements. However, the highly nonlinear pressure response measured on the blade suction surface is not captured in the present inviscid analysis.

Studies are currently under way to evaluate and develop test techniques for safely handling hydrogen exhaust products from hypersonic propulsion systems, as well as future liquid propellant rocket engines, during simulated altitude tests. Results of a study that considered several different handling concepts for air-breathing propulsion testing are presented. The study concluded that exhaust inerting to the oxidizer lean combustion limit is practical for direct-connect testing during normal engine operation, provided that bleed and plant surge-control airflows are negligible; however, diluent costs and exhaust pumping requirements can be high. The study also considered the so-called "hydrogen afterburning" concept. This technique uses injection of oxygen or air into the exhaust system to afterburn the exhaust products in real time in a controlled process. Hydrogen afterburning is attractive because of its potentially wide range of applicability and low secondary flow requirements. An analytical study was begun to investigate the technical details of hydrogen afterburning, which included the mixing of the exhaust products with the injected oxidizer and the potential for ignition and burning of the mixed gases. Conclusions from the analytical study were that adequate mixing can be obtained, and that autoignition and afterburning are feasible.

A two-dimensional mixed compression inlet model with a subsequent isolator section is tested under Mach 4 and 5 flight conditions. Different configurations of the isolator are assessed with respect to their compression efficiency. The experimental investigations yield schlieren pictures of the isolator flow and static surface pressure measurements. The numerical simulations are performed with a Reynolds-averaged Navier-Stokes solver using a k-omega turbulence model, especially,extended for modeling high-speed wall-bounded flows and separation regions. The close collaboration of experiment and simulation is beneficial: Validation of the simulation is achieved by the test data and the flowfield information available in the computational fluid dynamics results is employed to interpret the experimental findings and to compute the performance parameters. The computed static pressure ratios are in excellent agreement with empirical predictions. Furthermore, the investigations show that increasing the isolator length reduces the pressure sensitivity of the inlet. However, the experimental tests show that above a certain isolator length, no further increase of the sustainable backpressure is possible.

In this paper, we advocate a strategy for controlling a class of turbomachine instabilities, whose primitive phases can be molded by linear theory, but that eventually grow into a performance-limiting modification of the basic flow. The pehenomena of rotating stall and surge are two very different practical examples in which small disturbances grow to magnitudes such that they limit machine performance. We develop a theory that shows how an additional disturbance, driven from real-time data measured within the turbomachine, can be generated so as to realize a device with characteristics fundamentally different than those of the machine without control. For the particular compressor analyzed, the control increases the stable operating range by 20% of the mean flow. We show that active control can also be used to destabilize a compressor in an undesirable state such as nonrecoverable stall.

The aerodynamics of small amplitude pitching motions near stall have been studied experimentally in order to improve understanding of the torsional stall flutter problem for propeller blades. A model wing was oscillated in pitch at several small amplitudes over a wide and representative range of conditions. Unsteady surface pressures were measured and integrated to determine the aerodynamic damping at five spanwise stations. Strong negative damping was found for motions centered near static stall for all studied reduced frequencies, Mach numbers, and sweep angles. The 30-deg sweptback configuration was found to become negatively damped over the entire span nearly simultaneously, while the unswept model exhibited local regions of negative damping that moved toward the wing tip as the mean angle of attack was increased.

A mathematical model is developed to predict the unstalled to torsion mode flutter of an aerodynamically and structurally detuned rotor operating in a supersonic inlet flowfield with a subsonic leading-edge locus. Alternate blade structural detuning is considered. The aerodynamic detuning is accomplished by alternating the circumferential spacing of adjacent rotor blades.

A three-dimensional unsteady, viscous aerodynamic analysis has been developed for the flow inside a transonic, high-through-flow, single stage compressor. The compressor stage is comprised of a low-aspect-ratio rotor and a closely coupled stator. The analysis is based on a numerical method for solving the three-dimensional Navier-Stokes equation for unsteady viscous flow through multiple turbomachinery blade rows. The method solves the fully three-dimensional Navier-Stokes equation with an implicit scheme. A two-equation turbulence model with a low-Reynolds-number modification is applied for the turbulence closure. A third-order accurate upwinding scheme is used to approximate convection terms while a second-order accurate central difference scheme is used for the discretization of the viscous terms. A second-order accurate scheme is employed for the temporal discretization. The numerical method is applied to study the unsteady flow field inside a transonic, high-through-flow, axial compressor stage. The numerical results are compared with available experimental data.

A mathematical model is developed to demonstrate the potential benefit of splitter blades as a passive control mechanism for the flow-induced forced response of supersonic turbomachine rotors. The splitters introduce both aerodynamic and structural detuning and also provide a large measure of control of the location of the Mach wave full-chord airfoil and splitter intersections. The level of aerodynamic detuning is associated with the relative locations of the splitters in the full-chord airfoil passages, with the level of structural detuning a function of the ratio of the natural frequencies of the splitters to that of the full-chord airfoils.

The efficiency of nanosecond discharges as an active-particle generator for plasma-assisted combustion and ignition has been shown. The kinetics of alkane oxidation have been investigated from methane to decane in stoichiometric and lean mixtures with oxygen and air at room temperature under the action of high-voltage nanosecond unform discharge. The study of nanosecond barrier discharge influence on a flame propagation and flame blowoff velocity has been carried out. A significant increase of the flame blowoff velocity has been demonstrated. A decrease of 2-3 orders of magnitude of the plasma-assisted ignition delay time in comparison with the autoignition has been registered. Detonation initiating by high-voltage gas discharge has been demonstrated. The energy deposition in the discharge ranging from 70 mJ to 12 J for propane-oxygen-nitrogen mixtures leads to the transition to detonation at a distance of less than one diameter of the detonation tube. The influence of pulsed surface dielectric discharge on the flow separation for airfoils at a high angle of attack has been investigated within the velocity range from 20 to 110 m/s for the power consumption less than 1 W/cm of the wing span. The conclusion has been made that the main mechanism of plasma impact is the boundary-layer turbulization rather than acceleration.

Measured vibratory strain amplitudes resulting from off-axis flow are compared for the blades of two eight-bladed, 0.62-m- (2-ft-) diam propfan model rotors with mistuning. One rotor had inherent mistiming. The other was intentionally mistuned by replacing every other blade of the first rotor with a blade of same geometry but different frequencies and mode shapes. The data show that the intentional mistuning had a beneficial effect on the aeroelastic response of the propfan blades for a wide range of off-axis flow angles, blade pitch angles, and rotational speeds. The data also illustrate that large and intuitively unpredictable variations in the aeroelastic response of propfan blades can occur because of inherent mistuning. Statistical trends of blade strain amplitudes are compared for both the rotors in terms of the ratio of the maximum to the mean. © 1991 American Institute of Aeronautics and Astronautics, Inc., All rights reserved.

The flutter problem associated with the blades of a class of vertical-axis wind turbines called Darrieus is studied in detail. The spinning blade is supposed to be initially curved in a particular shape characterized by a state of pure tension at the blade cross section. From this equilibrium position a 3D linear perturbation pattern is superimposed to determine the dynamic aeroelastic stability of the blade in the presence of free wind speed by means of the Floquet-Liapunov theory for periodic systems.

A comprehensive study of Coriolis forces and shaft flexibility effects on the structural dynamics and aeroelastic stability of a rotating bladed-disk assembly attached to a cantilever, massless, flexible shaft is presented. Analyses were performed for an actual bladed-disk assembly, used as the first stage in the fan of the 'E3' engine. In the structural model, both in-plane and out-of-plane elastic deformation of the bladed-disk assembly were considered relative to their hub, in addition to rigid disk translations and rotations introduced by shaft flexibility. Besides structural coupling between blades (through the flexible disk), additional coupling is introduced through quasisteady aerodynamic loads. Rotational effects are accounted for throughout the work, and some mode shapes for the whole structure are presented at a selected rpm.

This paper is an overview of a wide range of recent aerospace technologies under development at the NASA Glenn Research Center, in collaboration with other NASA centers, government agencies, industry and academia. The focused areas are space solar power, advanced power management and distribution systems, Stirling cycle conversion systems, fuel cells, advanced thin film photovoltaics and batteries, and combustion technologies. The aerospace-related objectives of the technologies are generation of space power, development of cost-effective and reliable, high performance power systems, cryogenic applications, energy storage, and reduction in gas-turbine emissions, with attendant clean jet engines. The terrestrial energy applications of the technologies include augmentation of bulk power in ground power distribution systems, and generation of residential, commercial and remote power, as well as promotion of pollution-free environment via reduction in combustion emissions.

The two-phase flowfield in the aft-dome region of a solid rocket motor (SRM) with submerged nozzle has been simulated using a combined Eulerian-Lagrangian analysis. This analysis uses the numerical solution of ensemble averaged Navier-Stokes equations for the continuous (gas) phase coupled with a Lagrangian analysis for the discrete (particulate) phase to simulate the two-phase internal flow. A linearized block-implicit (LBI) scheme is used to solve the governing equations for the continuous phase, which allows the use of a highly stretched grid with sublayer resolution. The motion of the particles is tracked in computational coordinate space resulting in computational efficiency, and the interphase coupling terms for the Eulerian analysis are computed from the instantaneous distribution of the particles. A low Reynolds number form of the k-ε turbulence model is used with modifications for injection driven flows. Calculations have been performed for a particular grain configuration of the Space Shuttle SRM. The flowfield in the vicinity of the submerged nozzle, the particle trajectories, and the sensitivity of two-phase effects (such as slag accumulation) to the particle injection parameters are presented in this paper.

This study represents an evaluation of the extent to which conventional and recently introduced modern diagnostics can assess the symmetry of sprays formed by three atomizers of identical design. The conventional diagnostics include sheet-lit photography, patternation, and laser diffraction. The modern diagnostic is laser interferometry (phase Doppler). Symmetry is assessed in ambient conditions for four atomizer orientations, and comparisons are made between the diagnostic techniques. The results demonstrate that conventional and modern diagnostics are consistent in the assessment of symmetry, patternation and phase Doppler are most effective in establishing symmetry of mass flux, and phase Doppler, although more tedious to employ, provides the additional information necessary to establish the sources of detected asymmetries in terms of nonuniformities in droplet velocities, size distributions, volume flux, and concentration.

A numerical simulation of the temporally developing now through a generic scramjet combustor duct is presented for stagnation conditions typical of night at Mach 13. The particular focus is to examine the startup transients and to determine the time required for certain flow parameters to become established. The calculations were made with a Navier-Stokes solver SPARK with temporally relaxing inflow conditions derived from the operation of the T4 shock tunnel at the University of Queensland in Australia. The generic combustor geometry includes the injection of hydrogen fuel from the base of a centrally located strut. The flow was assumed laminar and fuel combustion was not included. The establishment process is presented for viscous parameters in the boundary layer and for parameters related to the fuel-air mixing.

The time-dependent Navier-Stokes equations, including the effects of finite rate chemistry, are numerically integrated to predict the steady-state behavior of several model supersonic flameholders. The conservation equations governing chemically reacting flows are solved using a technique based on the idea of rescaling the equations in time such that all convective and chemical processes evolve on similar 'pseudo' time scales. To accomplish this, the conservation equations are preconditioned to remove the stiffness associated with these equations. The method can be used to compute the steady-state solution very efficiently, regardless of whether the flow is frozen, finite rate, or in equilibrium. Two candidate supersonic flameholders are analyzed to assess their operating characteristics. The geometries include a ramp and a rearward-facing step. All flows consider a mixed H//2-airstream and use the global chemistry model of Rogers and Chinitz. Several different kinds of flowfields are generated, depending on the level of heat release.

A detailed study of the two-phase flow produced by a gas-turbine air-blast atomizer is performed with the goal of identifying the interaction between the two phases for both nonreacting and reacting conditions. A two-component phase Doppler interferometry is utilized to characterize three flowfields produced by the atomizer: (1) the single-phase flow, (2) the two-phase nonreacting spray, and (3) the two-phase reacting spray. Measurements of the mean and fluctuating axial and azimuthal velocities for each phase are obtained. In addition, the droplet size distribution, volume flux, and concentration are measured. The results reveal the strong influence of the dispersed phase on the gas, and the influence of reaction on both the gas and the droplet field. The presence of the spray significantly alters the inlet condition of the atomizer. With this alteration quantified, it is possible to deduce that the inertia associated with the dispersed phase damps the fluctuating velocities of the gas. Reaction reduces the volume flux of the droplets, broadens the local volume distribution of the droplets in the region of the reaction zone, increases the axial velocities and radial spread of the gas, and increases the anisotropy in the region of the reaction zone.

Initially uniform clouds of fuel particulates in air sustain processes which may lead to particle-cloud nonuniformities. In low gravity, flame-induced Kundt's tube phenomena are observed to form regular patterns of nonuniform particle concentrations. Irregular patterns of particle concentrations also are observed to result from selected nonuniform mixing processes. Low-gravity flame propagation for each of these classes of particle-cloud flames has been found to depend importantly on the flame-generated infrared radiative fields. The spatial structures of these radiative fields are described. Application is made for the observed cases of lycopodium-air flames.

A simplified kinetic mechanism is presented for 1006 reactions involving 134 chemical species of that 23 include nitrogen atoms that describes the combustion of n-heptane, iso-octane, and their mixtures in air. The submechanism for the C-5-C-8 species is directly derived from the earlier presented mechanisms for the combustion of methane, ethane, propane, and butane in air. The following experimental data taken from the literature and obtained by means of shock-tube experiments have been chosen to validate the mechanism: the pyrolysis Of i-C8H18 and the ignition delay times of mixtures of n-C7H16 and i-C8H18 with air, respectively. The comparisons have yielded good agreement for the initial temperature T-0 varying between 650 and 1200 K, the initial pressure P-0 ranging from 0.65 to 4.5 MPa, and the fuel-to-oxygen ratio phi 0 ranging from 0.5 to 2.