ALBERT

Description: Seal technology development is an important part of the Air Force's participation in the Integrated High Performance Turbine Engine Technology (IHPTET) initiative, the joint DOD, NASA, ARPA, and industry endeavor to double turbine engine capabilities by the turn of the century. Significant performance and efficiency improvements can be obtained through reducing internal flow system leakage, but seal environment requirements continue to become more extreme as the engine thermodynamic cycles advance towards these IHPTET goals. Seal technology continues to be pursued by the Air Force to control leakage at the required conditions. This presentation briefly describes current seal research and development programs and gives a summary of seal applications in demonstrator and developmental engines.

Description: Designers and customers are demanding higher performance turbomachine systems that have long life between overhauls and satisfy the more restrictive environmental constraints. This overview provides sources of design data, numerical, and experimental results along with selected new seal configurations and static sealing challenges such as in the combustors. The following categories are presented: (1) Seal Rotordynamic Data Base (experimental analytical program at Texas A&M); (2) Secondary Flow Interactions (validation studies at CFDRC, Huntsville AL); (3) Contact Sealing (selected types with finger seal model); and (4) Environmental Constraints (emphasis on combustors).

Description: This effort is to develop large diameter (22 - 36 inch) Aspirating Seals for application in aircraft engines. Stein Seal Co. will be fabricating the 36-inch seal(s) for testing. GE's task is to establish a thorough understanding of the operation of Aspirating Seals through analytical modeling and full-scale testing. The two primary objectives of this project are to develop the analytical models of the aspirating seal system, to upgrade using GE's funds, GE's 50-inch seal test rig for testing the Aspirating Seal (back-to-back with a corresponding brush seal), test the aspirating seal(s) for seal closure, tracking and maneuver transients (tilt) at operating pressures and temperatures, and validate the analytical model. The objective of the analytical model development is to evaluate the transient and steady-state dynamic performance characteristics of the seal designed by Stein. The transient dynamic model uses a multi-body system approach: the Stator, Seal face and the rotor are treated as individual bodies with relative degrees of freedom. Initially, the thirty-six springs are represented as a single one trying to keep open the aspirating face. Stops (Contact elements) are provided between the stator and the seal (to compensate the preload in the fully-open position) and between the rotor face and Seal face (to detect rub). The secondary seal is considered as part of the stator. The film's load, damping and stiffness characteristics as functions of pressure and clearance are evaluated using a separate (NASA) code GFACE. Initially, a laminar flow theory is used. Special two-dimensional interpolation routines are written to establish exact film load and damping values at each integration time step. Additionally, other user-routines are written to read-in actual pressure, rpm, stator-growth and rotor growth data and, later, to transfer these as appropriate loads/motions in the system-dynamic model. The transient dynamic model evaluates the various motions, clearances and forces as the seals are subjected to different aircraft maneuvers: Windmilling restart; start-ground idle; ground idle-takeoff; takeoff-burst chop, etc. Results of this model show that the seal closes appropriately and does not ram into the rotor for all of the conditions analyzed. The rig upgrade design for testing Aspirating Seals has been completed. Long lead-time items (forgings, etc.) have been ordered.

Description: Digital Particle Imaging Velocimetry (DPIV) is a powerful measurement technique, which can be used as an alternative or complementary approach to Laser Doppler Velocimetry (LDV) in a wide range of research applications. The instantaneous planar velocity measurements obtained with PIV make it an attractive technique for use in the study of the complex flow fields encountered in turbomachinery. Many of the same issues encountered in the application of LDV to rotating machinery apply in the application of PIV. Techniques for optical access, light sheet delivery, CCD camera technology and particulate seeding are discussed. Results from the successful application of the PIV technique to both the blade passage region of a transonic axial compressor and the diffuser region of a high speed centrifugal compressor are presented. Both instantaneous and time-averaged flow fields were obtained. The 95% confidence intervals for the time-averaged velocity estimates were also determined. Results from the use of PIV to study surge in a centrifugal compressor are discussed. In addition, combined correlation/particle tracking results yielding super-resolution velocity measurements are presented.

Description: The next generation of subsonic engines can be expected to continue the historical trend towards increased thrust to weight (T/W) and decreased specific fuel consumption (SFC). Development programs currently underway throughout the gas turbine industry such as DOD's Integrated High Performance Turbine Engine Technology (IHPTET), and more recently NASA's Advanced Subsonic Transport (AST) programs, have altered these trends in both pace and magnitude. Advanced seals and sealing technologies have become a prominent part of these efforts due to the large potential performance gains which can be realized. Allison has recently completed a study for NASA the goal of which was to quantize the potential performance benefits which might accrue through the use of advanced seals in future subsonic gas turbine engines. For the study, two engines where analyzed, a small turboshaft and a larger turbofan engine to help assess the effect of engine size on the results. Engines were analyzed stage by stage with the most sensitive areas highlighted. Leakage characteristics for advanced seals were then substituted into secondary airflow models, and the leakage reductions documented. These leakage reductions were then converted to changes in performance, i.e. increased range, decreased takeoff gross weight, etc. and presented. It was found that the development and use of a realtively few advanced seals, less than 5, could for example reduce SFC by 10% or more.

Description: Thermal analysis in both simple and complex models can require calculating the propagation of radiant energy to and from multiple surfaces. This can be accomplished through simple estimation techniques or complex computationally intense computer modeling simulations. Currently there are a variety of computer analysis techniques used to simulate the propagation of radiant energy, each having advantages and disadvantages. The major objective of this effort was to compare two ray tracing radiation propagation analysis programs (NEVADA and TSS) Net Energy Verification and Determination Analyzer and Thermal Synthesizer System with experimental data. Results from a non-flowing, electrically heated test rig was used to verify the calculated radiant energy propagation from a nozzle geometry that represents an aircraft propulsion nozzle system. In general the programs produced comparable overall results, and results slightly higher then the experimental data. Upon inspection of individual radiation interchange factors, differences were evident and would have been magnified if a more radical model temperature profile was analyzed. Bidirectional reflectivity data (BRDF) was not used do to modeling limitations in TSS. For code comparison purposes, this nozzle geometry represents only one case for one set of analysis conditions. Since each computer code has advantages and disadvantages bases on scope, requirements, and desired accuracy, the usefulness of this single case study may be limiting.

Description: Through the NASA/Industry Cooperative Effort (NICE) agreement, NASA Lewis and industry partners are developing a new engine simulation, called the National Cycle Program (NCP), which is the initial framework of NPSS. NCP is the first phase toward achieving the goal of NPSS. This new software supports the aerothermodynamic system simulation process for the full life cycle of an engine. The National Cycle Program (NCP) was written following the Object Oriented Paradigm (C++, CORBA). The software development process used was also based on the Object Oriented paradigm. Software reviews, configuration management, test plans, requirements, design were all apart of the process used in developing NCP. Due to the many contributors to NCP, the stated software process was mandatory for building a common tool intended for use by so many organizations. The U.S. aircraft and airframe companies recognize NCP as the future industry standard for propulsion system modeling.

Description: 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 generations 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.

Description: The crack propagation life of tested specimens has been repeatedly shown to strongly depend on the loading history. Overloads and extended stress holds at temperature can either retard or accelerate the crack growth rate. Therefore, to accurately predict the crack propagation life of an actual component, it is essential to approximate the true loading history. In military rotorcraft engine applications, the loading profile (stress amplitudes, temperature, and number of excursions) can vary significantly depending on the type of mission flown. To accurately assess the durability of a fleet of engines, the crack propagation life distribution of a specific component should account for the variability in the missions performed (proportion of missions flown and sequence). In this report, analytical and experimental studies are described that calibrate/validate the crack propagation prediction capability for a disk alloy under variable amplitude loading. A crack closure based model was adopted to analytically predict the load interaction effects. Furthermore, a methodology has been developed to realistically simulate the actual mission mix loading on a fleet of engines over their lifetime. A sequence of missions is randomly selected and the number of repeats of each mission in the sequence is determined assuming a Poisson distributed random variable with a given mean occurrence rate. Multiple realizations of random mission histories are generated in this manner and are used to produce stress, temperature, and time points for fracture mechanics calculations. The result is a cumulative distribution of crack propagation lives for a given, life limiting, component location. This information can be used to determine a safe retirement life or inspection interval for the given location.

Description: Unsteady blade row interactions in turbomachines generate discrete-frequency tones at blade pass frequency and its harmonics. Specific circumferential acoustic modes are generated. However, only certain of these modes propagate upstream and downstream to the far field, with these the discrete frequency noise received by an observer. This paper is directed at experimentally demonstrating the viability of active noise control utilizing active airfoils to generate propagating spatial modes that interact with and simultaneously cancel the upstream and downstream propagating acoustic modes. This is accomplished by means of fundamental experiments performed in the Purdue Annular Cascade Research Facility configured with 16 rotor blades and 18 stator vanes. At blade pass frequency, only the k(sub 0) = -2 spatial mode propagates. Significant simultaneous noise reductions are achieved for these upstream and downstream propagating spatial modes over a wide range of operating conditions.

Description: A comprehensive assessment is made of the predictive capability of the average passage flow model as applied to multi-stage axial flow compressors. The average passage flow model describes the time average flow field within a typical passage of a blade row embedded in a multi-stage configuration. In this work data taken within a four and one-half stage large low speed compressor will be used to assess the weakness and strengths of the predictive capabilities of the average passage flow model. The low speed compressor blading is of modern design and employs stators with end-bends. Measurements were made with slow and high response instrumentation. The high response measurements revealed the velocity components of both the rotor and stator wakes. Based on the measured wake profiles it will be argued that blade boundary layer transition is playing an important role in setting compressor performance. A model which mimics the effects of blade boundary layer transition within the frame work of the average passage model will be presented. Simulations which incorporated this model showed a dramatic improvement in agreement with data.

Description: Cycle studies have shown the benefits of increasing engine pressure ratios and cycle temperatures to decrease engine weight and improve performance in next generation turbine engines. Advanced seals have been identified as critical in meeting engine goals for specific fuel consumption, thrust-to-weight, emissions, durability and operating costs. NASA and the industry are identifying and developing engine and sealing technologies that will result in dramatic improvements and address the goals for engines entering service in the 2005-2007 time frame. This paper provides an overview of advanced seal technology requirements and highlights the results of a preliminary design effort to implement advanced seals into a regional aircraft turbine engine. This study examines in great detail the benefits of applying advanced seals in the high pressure turbine region of the engine. Low leakage film-riding seals can cut in half the estimated 4% cycle air currently used to purge the high pressure turbine cavities. These savings can be applied in one of several ways. Holding rotor inlet temperature (RIT) constant the engine specific fuel consumption can be reduced 0.9%, or thrust could be increased 2.5%, or mission fuel burn could be reduced 1.3%. Alternatively, RIT could be lowered 20 'F resulting in a 50% increase in turbine blade life reducing overall regional aircraft maintenance and fuel bum direct operating costs by nearly 1%. Thermal, structural, secondary-air systems, safety (seal failure and effect), and emissions analyses have shown the proposed design is feasible.

Description: This paper presents a coupled analysis of the interaction between mainpath and secondary flowpaths in gas turbines using transient simulations. Some of the topics include: 1) Need for Coupled Analysis; 2) Primary-Secondary Coupling Schematic; 3) Secondary Flow Requirement; 4) Objectives of Present Methodology; 5) Current Methodologies Recap; 6) Proposed Coupled Code Methodology; 7) Description of SCISEAL Code; 8) Description of Turbo Code; 9) Code Coupling/Interface Issues; and 10) Current Interface Strategy. This paper is presented in viewgraph form.

Description: NASA's General Aviation Propulsion (GAP) program is a cooperative program between government and industry. NASA's strategic direction is described by the "Three Pillars" and their Objectives as set forth by NASA Administrator Daniel S. Goldin. NASA's Three Pillars are: 1) Global Civil Aviation, 2) Revolutionary Technology Leaps, and 3) Access To Space.

Description: A formal method is described to quantify structural damage tolerance and reliability in the presence of multitude of uncertainties in turbine engine components. The method is based at the materials behaviour level where primitive variables with their respective scatters are used to describe the behavior. Computational simulation is then used to propagate those uncertainties to the structural scale where damage tolerance and reliability are usually specified. Several sample cases are described to illustrate the effectiveness, versatility, and maturity of the method. Typical results from these methods demonstrate that the methods are mature and that they can be used for future strategic projections and planning to assure better, cheaper, faster, products for competitive advantages in world markets. These results also indicate that the methods are suitable for predicting remaining life in aging or deteriorating structures.

Description: This paper describes a micrometeroid protection system for the main engines of the Cassini spacecraft. The engine Cover Assembly is a deployable/restowable half sphere of multilayer insulation mounted to an articulatable frame over 2 meters (7 feet) in diameter. The Cover folds into a compact wedge only 25 cm (10 inches) at its maximum thickness. The micrometeroid environment and typical protection methods are described as well as the design details and development problems of the Cover Mechanism Assembly.

Description: A Rocket Based Combined Cycle (RBCC) engine system is designed to combine the high thrust to weight ratio of a rocket along with the high specific impulse of a ramjet in a single, integrated propulsion system. This integrated, combined cycle propulsion system is designed to provide higher vehicle performance than that achievable with a separate rocket and ramjet. The RBCC engine system studied in the current program is the Aerojet strutjet engine concept, which is being developed jointly by a government-industry team as part of the Air Force HyTech program pre-PRDA activity. The strutjet is an ejector-ramjet engine in which small rocket chambers are embedded into the trailing edges of the inlet compression struts. The engine operates as an ejector-ramjet from takeoff to slightly above Mach 3. Above Mach 3 the engine operates as a ramjet and transitions to a scramjet at high Mach numbers. For space launch applications the rockets would be re-ignited at a Mach number or altitude beyond which air-breathing propulsion alone becomes impractical. The focus of the present study is to develop and demonstrate a strutjet flowpath using hydrocarbon fuel at up to Mach 7 conditions.

Description: The material to be presented in these two lectures begins with cycle considerations of the turbojet engine combined with a ramjet engine to provide thrust over the range of Mach 0 to 5. We will then examine in some detail the aerodynamic behavior that occurs in the inlet operating near the peak speed. Following that, we shall view a numerical simulation through a baseline scramjet engine, starting at the entrance to the inlet, proceeding into the combustor and through the nozzle. In the next segment, we examine a combined rocket and ramjet propulsion system. Analysis and test results will be examined with a view toward evaluation of the concept as a practical device. Two other inlets will then be reviewed: a Mach 12 inlet and a Mach 18 configuration. Finally, we close our lectures with a discussion of the Detonation Wave engine, and inspect the physical and chemical behavior obtained from numerical simulation. A few final remarks will be made regarding the application of CFD for hypersonic propulsion components.

Description: The research vision of the NASA Lewis Research Center in the area of integrated flight and propulsion controls technologies is described. In particular, the integrated method for propulsion and airframe controls developed at the Lewis Research Center is described including its application to an advanced aircraft configuration. Additionally, future research directions in integrated controls are described.

Description: Extensive testing done on a T55-L-712 turboshaft engine compressor in a compressor test rig is being followed by engine tests in progress as part of the Army Non-Recoverable Stall Program. Goals include a greater understanding of the gas turbine engine start cycle and compressor/engine operation in the regions 'beyond' the normal compressor stall line (rotating stall/surge). Rig steady state instrumentation consisted of 497 steady state pressure sensors and 153 temperature sensors. Engine instrumentation was placed in similar radial/axial locations and consists of 122 steady state pressure sensors and 65 temperature sensors. High response rig instrumentation consisted of 34 wall static pressure transducers. Rig and engine high response pressure transducers were located in the same axial/radial/circumferential locations in front of the first three stages. Additional engine high response instrumentation was placed in mach probes in front of the engine and on the compressor hub. This instrumentation allows for the generation of detailed stage characteristics, overall compressor mapping, and detailed analysis of dynamic compressor events.

Description: This presentation highlights the activities that researchers at the NASA Lewis Research Center (LeRC) have been and will be involved in to assess integrated nozzle performance. Three different test activities are discussed. First, the results of the Propulsion Airframe Integration for High Speed Research 1 (PAIHSR1) study are presented. The PAIHSR1 experiment was conducted in the LeRC 9 ft x l5 ft wind tunnel from December 1991 to January 1992. Second, an overview of the proposed Mixer/ejector Inlet Distortion Study (MIDIS-E) is presented. The objective of MIDIS-E is to assess the effects of applying discrete disturbances to the ejector inlet flow on the acoustic and aero-performance of a mixer/ejector nozzle. Finally, an overview of the High-Lift Engine Aero-acoustic Technology (HEAT) test is presented. The HEAT test is a cooperative effort between the propulsion system and high-lift device research communities to assess wing/nozzle integration effects. The experiment is scheduled for FY94 in the NASA Ames Research Center (ARC) 40 ft x 80 ft Low Speed Wind Tunnel (LSWT).

Description: Interest in developing a new generation supersonic transport has increased in the past several years. Current projections indicate this aircraft would cruise at approximately Mach 2.4, have a range of 5000 nautical miles and carry at least 250 passengers. A large market for such an aircraft will exist in the next century due to a predicted doubling of the demand for long range air transportation by the end of the century and the growing influence of the Pacific Rim nations. Such a proposed aircraft could more than halve the flying time from Los Angeles to Tokyo. However, before a new economically feasible supersonic transport can be built, many key technologies must be developed. Among these technologies is noise suppression. Propulsion systems for a supersonic transport using current technology would exceed acceptable noise levels. All new aircraft must satisfy FAR 36 Stage III noise regulations. The largest area of concern is the noise generated during takeoff. A concerted effort under NASA's High Speed Research (HSR) program has begun to address the problem of noise suppression. One of the most promising concepts being studied in the area of noise suppression is the mixer/ejector nozzle. This study analyzes a typical noise suppressing mixer ejector nozzle at take off conditions, using a Full Navier-Stokes (FNS) computational fluid dynamics (CFD) code.

Description: Outline of presentation are: (1) Review of experimental apparatus. (2) Effect of natural screech of jet mixing; converging nozzle, underexpanded jet and converging-diverging nozzle, design pressure.(3) Effect of induced screech on jet mixing: produced by paddles in shear layers, similar to edge tones, and converging-diverging nozzle, design pressure. (4) Effect of paddles on near-field jet noise. and (5) Concluding remarks.

Description: The motivation of the testing was to reduce noise generated by eddy Mach wave emission via enhanced mixing in the jet plume. This was to be accomplished through the use of an ejector shroud, which would bring in cooler ambient fluid to mix with the hotter jet flow. In addition, the contour of the mixer, with its chutes and lobes, would accentuate the merging of the outer and inner flows. The objective of the focused schlieren work was to characterize the mixing performance inside of the ejector. Using flow visualization allowed this to be accomplished in a non-intrusive manner.

Description: Only recently has computational fluid dynamics (CFD) been relied upon to predict the flow details of advanced nozzle concepts. Computer hardware technology and flow solving techniques are advancing rapidly and CFD is now being used to analyze such complex flows. Validation studies are needed to assess the accuracy, reliability, and cost of such CFD analyses. At NASA Lewis, the PARC2D/3D full Navier-Stokes (FNS) codes are being applied to HSR-type nozzles. This report presents the results of two such PARC FNS analyses. The first is an analysis of the Pratt and Whitney 2D mixer-ejector nozzle, conducted by Dr. Yunho Choi (formerly of Sverdrup Technology-NASA Lewis Group). The second is an analysis of NASA-Langley's axisymmetric single flow plug nozzle, conducted by the author.

Description: The investigation includes carry out fundamental experiments studying mechanisms of effect: (1) experiments on subsonic and supersonic jets to assess influence of compressibility, (2) parametric study on tab geometry to optimize effect for given flow blockage (this effort led to "delta-tab"), (3) quantify mixing enhancement in the jet, and (4) analyze mechanism of streamwise vorticity generation.

Description: The theory of mixer-ejectors for noise suppression is illustrated in this cartoon. Since jet noise SPL scales as velocity to the eighth power and diameter squared, increasing the jet diameter while lowering its velocity and keeping thrust constant decreases the noise. However, in supersonic craft, the drag penalty for increasing diameter at supersonic cruise makes this option very expensive. One would like to have a large engine during takeoff which could be shrunk during cruise. The retractable ejector is such an expandable engine. If the mixer flow can be expanded to the size of the ejector exit, the noise generated downstream of the ejector will be much less than the small diameter mixer nozzle alone. Of course, this also requires that the noise created in expanding the flow to fill the ejector be absorbed by a liner in the ejector walls so that none of this noise is heard. Since this mixing of internal hot gas and external cold air must take place in as short a distance as possible, the mixer must be very effective and therefore probably much noisier than a simple nozzle.

Description: The work conducted during the Summer of 1995 for the Langley Aerospace Research Summer Scholars, or LARSS, Program was a continuation of Master's Degree work being conducted for the Mechanical Engineering Department at Old Dominion University. Since this work is not yet complete, an update of progress is provided here along with a generalized background. The main emphasis of this research is to find predicted correlations in the database generated by the SHIP3D code, which modeled different scramjet combustor configurations.

Description: Established analyses of conventional ramjet/scramjet performance characteristics indicate that a considerable decrease in efficiency can be expected at off-design flight conditions. This can be explained, in large part, by the deterioration of intake mass flow and limited inlet compression at low flight speeds and by the onset of thrust degradation effects associated with increased burner entry temperature at high flight speeds. In combination, these effects tend to impose lower and upper Mach number limits for practical flight. It has been noted, however, that Magnetohydrodynamic (MHD) energy management techniques represent a possible means for extending the flight Mach number envelope of conventional engines. By transferring enthalpy between different stages of the engine cycle, it appears that the onset of thrust degradation may be delayed to higher flight speeds. Obviously, the introduction of additional process inefficiencies is inevitable with this approach, but it is believed that these losses are more than compensated through optimization of the combustion process. The fundamental idea is to use MHD energy conversion processes to extract and bypass a portion of the intake kinetic energy around the burner. We refer to this general class of propulsion system as an MHD-bypass engine. In this paper, we quantitatively assess the performance potential and scientific feasibility of MHD-bypass airbreathing hypersonic engines using ideal gasdynamics and fundamental thermodynamic principles.

Description: Experimental data have shown that combustor temperature non-uniformities can lead to the excessive heating of first-stage rotor blades in turbines. This heating of the rotor blades can lead to thermal fatigue and degrade turbine performance. The results of recent studies have shown that variations in the circumferential location (clocking) of the hot streak relative to the first-stage vane airfoils can be used to minimize the adverse effects of the hot streak. The effects of the hot streak/airfoil count ratio on the heating patterns of turbine airfoils have also been evaluated. In the present investigation, three-dimensional unsteady Navier-Stokes simulations have been performed for a single-stage high-pressure turbine operating in high subsonic flow. In addition to a simulation of the baseline turbine, simulations have been performed for circular and elliptical hot streaks of varying sizes in an effort to represent different combustor designs. The predicted results for the baseline simulation show good agreement with the available experimental data. The results of the hot streak simulations indicate: that a) elliptical hot streaks mix more rapidly than circular hot streaks, b) for small hot streak surface area the average rotor temperature is not a strong function of hot streak temperature ratio or shape, and c) hot streaks with larger surface area interact with the secondary flows at the rotor hub endwall, generating an additional high temperature region.

Description: An investigation has been conducted to develop appropriate technologies for a low-NO(x), liquid-fueled combustor. The combustor incorporates an effervescent atomizer used to inject fuel into a premixing duct. Only a fraction of the combustion air is used in the premixing process. This fuel-rich mixture is introduced into the remaining combustion air by a rapid jet-shear-layer mixing process involving radial fuel-air jets impinging on axial air jets in the primary combustion zone. Computational modeling was used as a tool to facilitate a parametric analysis appropriate to the design of an optimum low-NO(x) combustor. A number of combustor configurations were studied to assess the key combustor technologies and to validate the three-dimensional modeling code. The results from the experimental testing and computational analysis indicate a low-NO(x) potential for the jet-shear-layer combustor. Key features found to affect NOx emissions are the primary combustion zone fuel-air ratio, the number of axial and radial jets, the aspect ratio and radial location of the axial air jets, and the radial jet inlet hole diameter. Each of these key parameters exhibits a low-NO(x) point from which an optimized combustor was developed Also demonstrated was the feasibility of utilizing an effervescent atomizer for combustor application. Further developments in the jet-shear-layer mixing scheme and effervescent atomizer design promise even lower NO(x) with high combustion efficiency.

Description: The flow through the tip clearance region of a transonic compressor rotor (NASA rotor 37) was computed and compared to aerodynamic probe and laser anemometer data. Tip clearance effects were modeled both by gridding the clearance gap and by using a simple periodicity model across the ungridded gap. The simple model was run with both the full gap height, and with half the gap height to simulate a vena-contracta effect. Comparisons between computed and measured performance maps and downstream profiles were used to validate the models and to assess the effects of gap height on the simple clearance model. Recommendations were made concern- ing the use of the simple clearance model Detailed comparisons were made between the gridded clearance gap solution and the laser anemometer data near the tip at two operating points. The computed results agreed fairly well with the data but overpredicted the extent of the casing separation and underpredicted the wake decay rate. The computations were then used to describe the interaction of the tip vortex, the passage shock, and the casing boundary layer.

Description: The need for efficient access to space has created interest in airbreathing propulsion as a means of achieving that goal. The NASP program explored a single-stage-to-orbit approach which could require scramjet airbreathing propulsion out to Mach 16 to 20. Recent interest in global access could require hypersonic cruise engines operating efficiently in the Mach 10 to 12 speed range. A common requirement of both these types of propulsion systems is that they would have to be fully integrated with the aero configuration so that the forebody becomes a part of the external compression inlet and the nozzle expansion is completed on the vehicle aftbody.

Description: The United States has embarked on a national effort to develop the technology necessary to produce a Mach 2.4 High Speed Civil Transport (HSCT) for entry into service by the year 2005. The viability of this aircraft is contingent upon its meeting both economic and environmental requirements. Two engine components have been identified as critical to the environmental acceptability of the HSCT. These include a combustor with significantly lower emissions than are feasible with current technology, and a lightweight exhaust nozzle that meets community noise standards. The Enabling Propulsion Materials (EPM) program will develop the advanced structural materials, materials fabrication processes, structural analysis and life prediction tools for the HSCT combustor and low noise exhaust nozzle. This is being accomplished through the coordinated efforts of the NASA Lewis Research Center, General Electric Aircraft Engines and Pratt & Whitney. The mission of the EPM Exhaust Nozzle Team is to develop and demonstrate this technology by the year 1999 to enable its timely incorporation into HSCT propulsion systems.

Description: This paper describes work currently in progress at Langley on liner concepts that employ structures that may be suitable for broadband exhaust noise attenuation in high speed flow environments and at elevated temperatures characteristic of HSCT applications. Because such liners will need to provide about 10 dB suppression over a 2 to 3 octave frequency range, conventional single-degree-of-freedom resonant structures will not suffice. Bulk absorbers have the needed broadband absorption characteristic; however, at lower frequencies they tend to be inefficient.

Description: Based on extensive work performed by Dr. Thomas H. Sobota (Advanced Projects Research Incorporated (APRI)) on swirling flows in circular-to-rectangular transition sections, a model assembly was designed and fabricated in support of a Phase 1 Small Business Innovation Research Contract between the NASA-Langley Research Center and APRI. This assembly was acoustically tested as part of this Phase 1 effort, the goal being to determine whether the controlled introduction of axial vorticity could affect the various noise generation mechanisms present in an underexpanded supersonic rectangular jet.

Description: This paper discusses a test that was conducted jointly by Pratt & Whitney Aircraft Engines and NASA Lewis Research Center. The test was conducted in NASA's 9- by 15-Foot Low Speed Wind Tunnel (9x15 LSWT). The test setup, methods, and aerodynamic results of this test are discussed. Acoustical results are discussed in a separate paper by J. Bridges and J. Marino.

Description: GE Aircraft's GE90 is a high bypass turbofan jetliner engine capable of well over 84,700 pounds thrust. The turbofan is a propulsion system that compresses some of the air taken in, burns it in a combuster and expells it to generate power for driving the fan and compressor. A greater amount of air bypasses the combustion process. The GE90 pushes the cooler bypass air rearward with a fan to mix it with the hot exhaust gas; the result is a gain in thrust with minimal fuel expenditure. Over a billion dollars and several years went into its development, which included incorporating technologies developed by Lewis Research Center work done in the 1970s and from projects with SNECMA of France. The engine will power the 777 and other subsonic commercial widebodies.

Description: In a effort to address current needs for efficient, air propulsion systems, we have developed some new analytical predictive tools for understanding and alleviating aircraft engine instabilities which have led to accelerated high cycle fatigue and catastrophic failures of these machines during flight. A frequent cause of failure in Jets engines is excessive resonant vibrations and stall flutter instabilities. The likelihood of these phenomena is reduced when designers employ the analytical models we have developed. These prediction models will ultimately increase the nation's competitiveness in producing high performance Jets engines with enhanced operability, energy economy, and safety. The objectives of our current threads of research in the final year are directed along two lines. First, we want to improve the current state of blade stress and aeromechanical reduced-ordered modeling of high bypass engine fans, Specifically, a new reduced-order iterative redesign tool for passively controlling the mechanical authority of shroudless, wide chord, laminated composite transonic bypass engine fans has been developed. Second, we aim to advance current understanding of aeromechanical feedback control of dynamic flow instabilities in axial flow compressors. A systematic theoretical evaluation of several approaches to aeromechanical feedback control of rotating stall in axial compressors has been conducted. Attached are abstracts of two .papers under preparation for the 1998 ASME Turbo Expo in Stockholm, Sweden sponsored under Grant No. NAG3-1571. Our goals during the final year under Grant No. NAG3-1571 is to enhance NASA's capabilities of forced response of turbomachines (such as NASA FREPS). We with continue our development of the reduced-ordered, three-dimensional component synthesis models for aeromechanical evaluation of integrated bladeddisk assemblies (i.e., the disk, non-identical bladeing etc.). We will complete our development of component systems design optimization strategies for specified vibratory stresses and increased fatigue life prediction of assembly components, and for specified frequency margins on the Campbell diagrams of turbomachines. Finally, we will integrate the developed codes with NASA's turbomachinery aeromechanics prediction capability (such as NASA FREPS).

Description: A series of analyses have been developed which permit the calculation of the performance of common inlet designs. The methods presented are useful for determining the inlet weight flows, total pressure recovery, and aerodynamic drag coefficients for given inlet geometric designs. Limited geometric input data is required to use this inlet performance prediction methodology. The analyses presented here may also be used to perform inlet preliminary design studies. The calculated inlet performance parameters may be used in subsequent engine cycle analyses or installed engine performance calculations for existing uninstalled engine data.

Description: The focus of this study was on the autoxidation kinetics of deposit precursor formation in jet fuels. The objectives were: (1) to demonstrate that laser-induced fluorescence is a viable kinetic tool for measuring rates of deposit precursor formation in jet fuels; (2) to determine global rate expressions for the formation of thermal deposit precursors in jet fuels; and (3) to better understand the chemical mechanism of thermal stability. The fuels were isothermally stressed in small glass ampules in the 120 to 180 C range. Concentrations of deposit precursor, hydroperoxide and oxygen consumption were measured over time in the thermally stressed fuels. Deposit precursors were measured using laser-induced fluorescence (LIF), hydroperoxides using a spectrophotometric technique, and oxygen consumption by the pressure loss in the ampule. The expressions, I.P. = 1.278 x 10(exp -11)exp(28,517.9/RT) and R(sub dp) = 2.382 x 10(exp 17)exp(-34,369.2/RT) for the induction period, I.P. and rate of deposit precursor formation R(sub dp), were determined for Jet A fuel. The results of the study support a new theory of deposit formation in jet fuels, which suggest that acid catalyzed ionic reactions compete with free radical reactions to form deposit precursors. The results indicate that deposit precursors form only when aromatics are present in the fuel. Traces of sulfur reduce the rate of autoxidation but increase the yield of deposit precursor. Free radical chemistry is responsible for hydroperoxide formation and the oxidation of sulfur compounds to sulfonic acids. Phenols are then formed by the acid catalyzed decomposition of benzylic hydroperoxides, and deposit precursors are produced by the reaction of phenols with aldehydes, which forms a polymer similar to Bakelite. Deposit precursors appear to have a phenolic resin-like structure because the LIF spectra of the deposit precursors were similar to that of phenolic resin dissolved in TAM.

Description: The objective of this program was to define the aerodynamic design and manufacturing coordinates for an advanced 4:1 pressure ratio, single stage centrifugal compressor at a 10 lbm/sec flow size. The approach taken was to perform an exact scale of an existing DDA compressor originally designed at a flow size of 3.655 lbm/sec.

Description: A simple and accurate nozzle performance analysis methodology has been developed. The geometry modeling requirements are minimal and very flexible, thus allowing rapid design evaluations. The solution techniques accurately couple: continuity, momentum, energy, state, and other relations which permit fast and accurate calculations of nozzle gross thrust. The control volume and internal flow analyses are capable of accounting for the effects of: over/under expansion, flow divergence, wall friction, heat transfer, and mass addition/loss across surfaces. The results from the nozzle performance methodology are shown to be in excellent agreement with experimental data for a variety of nozzle designs over a range of operating conditions.

Description: This thesis covers the design and setup of a laser doppler velocimeter (LDV) system used to take velocity measurements in an annular combustor model. The annular combustor model is of contemporary design using 60 degree flat vane swirlers, producing a strong recirculation zone. Detailed measurements are taken of the swirler inlet air flow and of the downstream enclosed swirling flow. The laser system used is a two color, two component system set up in forward scatter. Detailed are some of the special considerations needed for LDV use in the confined turbulent flow of the combustor model. LDV measurements in a single swirler rig indicated that the flow changes radically in the first duct height. After this, a flow profile is set up and remains constant in shape. The magnitude of the velocities gradually decays due to viscous damping.

Description: Experimental data from jet-engine tests have indicated that unsteady blade-row interaction effects can have a significant impact on the efficiency of low-pressure turbine stages. Measured turbine efficiencies at takeoff can be as much as two points higher than those at cruise conditions. Preliminary studies indicate that Reynolds number effects may contribute to the lower efficiencies at cruise conditions. In the current study, numerical experiments have been performed to quantify the Reynolds number dependence of unsteady wake/separation bubble interaction on the performance of a low-pressure turbine.

Description: Advanced airbreathing propulsion systems used in Mach 4-6 mission scenarios, usually involve turbo-ramjet configurations. As the engines transition from turbojet to ramjet, there is an operational envelope where both engines operate simultaneously. In the first phase of our study, an over/under nozzle configuration was analyzed. The two plumes from the turbojet and ramjet interact at the end of a common 2-D cowl, where they both reach an approximate Mach 3.0 condition and then jointly expand to Mach 3.6 at the common nozzle exit plane. For the problem analyzed, the turbojet engine operates at a higher nozzle pressure ratio than the ramjet, causes the turbojet plume overpowers the ramjet plume, deflecting it approximately 12 degrees downward and in turn the turbojet plume is deflected 6 degrees upward. In the process, shocks were formed at the deflections and a shear layer formed at the confluence of the two jets. This particular case was experimentally tested and the data were used to compare with a computational fluid dynamics (CFD) study using the PARC2D code. The CFD results were in good agreement with both static pressure distributions on the cowl separator and on nozzle walls. The thrust coefficients were also in reasonable agreement. In addition, inviscid relationships were developed around the confluence point, where the two exhaust jets meet, and these results compared favorably with the CFD results. In the second phase of our study, a 3-D CFD solution was generated to compare with the 2-D solution. The major difference between the 2-D and 3-D solutions was the interaction of the shock waves, generated by the plume interactions, on the sidewall. When a shock wave interacts with a sidewall and sidewall boundary layer, it is called a glancing shock sidewall interaction. These interactions entrain boundary layer flow down the shockline into a vortical flow pattern. The 3-D plots show the streamlines being entrained down the shockline. The pressure of the flow also decreases slightly as the sidewall is approached. Other difference between the 2-D and 3-D solutions were a lowering of the nozzle thrust coefficient value from 0.9850 (2-D) to 0.9807 (3-D), where the experimental value was 0.9790. In the third phase of our study, a different turbo-ramjet configuration was analyzed. The confluence of a supersonic turbojet and a subsonic ramjet in the turbine based combined-cycle (TBCC) propulsion system was studied by a 2-D CFD code. In the analysis, Mach 1.4 primary turbojet was mixed with the subsonic ramjet secondary flow in an ejector mode operation. Reasonable agreements were obtained with the supplied I-D TBCC solutions. For low downstream backpressure, the Fabri choke condition (Break-Point condition) was observed in the secondary flow within mixing zone. For sufficient high downstream backpressure, the Fabri choke no longer exist, the ramjet flow was reduced and the ejector flow became backpressure dependent. Highly non-uniform flow at ejector exit were observed, indicated that for smooth downstream combustion, the mixing of the two streams probably required some physical devices.

Description: Recent experience using ANOPP to predict turbofan engine flyover noise suggests that it over-predicts overall EPNL by a significant amount. An improvement in this prediction method is desired for system optimization and assessment studies of advanced UHB engines. An assessment of the ANOPP fan inlet, fan exhaust, jet, combustor, and turbine noise prediction methods is made using static engine component noise data from the CF6-8OC2, E(3), and QCSEE turbofan engines. It is shown that the ANOPP prediction results are generally higher than the measured GE data, and that the inlet noise prediction method (Heidmann method) is the most significant source of this overprediction. Fan noise spectral comparisons show that improvements to the fan tone, broadband, and combination tone noise models are required to yield results that more closely simulate the GE data. Suggested changes that yield improved fan noise predictions but preserve the Heidmann model structure are identified and described. These changes are based on the sets of engine data mentioned, as well as some CFM56 engine data that was used to expand the combination tone noise database. It should be noted that the recommended changes are based on an analysis of engines that are limited to single stage fans with design tip relative Mach numbers greater than one.

Description: A gas chromatograph (GC)/mass spectrometer (MS) system that allows the speciation of unburnt hydrocarbons in the combustor exhaust has been developed at the NASA Lewis Research Center. Combustion gas samples are withdrawn through a water-cooled sampling probe which, when not in use, is protected from contamination by a high-pressure nitrogen purge. The sample line and its connecting lines, filters, and valves are all ultraclean and are heated to avoid condensation. The system has resolution to the parts-per-billion (ppb) level.

Description: The NASA Lewis Research Center is capitalizing on breakthroughs in foil air bearing performance, tribological coatings, and computer analyses to formulate the Oil-free Turbomachinery Program. The program s long-term goal is to develop an innovative, yet practical, oil-free aeropropulsion gas turbine engine that floats on advanced air bearings. This type of engine would operate at higher speeds and temperatures with lower weight and friction than conventional oil-lubricated engines. During startup and shutdown, solid lubricant coatings are required to prevent wear in such engines before the self-generating air-lubrication film develops. NASA s Tribology Branch has created PS304, a chrome-oxide-based plasma spray coating specifically tailored for shafts run against foil bearings. PS304 contains silver and barium fluoride/calcium fluoride eutectic (BaF2/CaF2) lubricant additives that, together, provide lubrication from cold start temperatures to over 650 C, the maximum use temperature for foil bearings. Recent lab tests show that bearings lubricated with PS304 survive over 100 000 start-stop cycles without experiencing any degradation in performance due to wear. The accompanying photograph shows a test bearing after it was run at 650 C. The rubbing process created a "polished" surface that enhances bearing load capacity.

Description: One of the objectives of NASA's High Performance Computing and Communications Program's (HPCCP) Numerical Propulsion System Simulation (NPSS) is to reduce the time and cost of generating aerothermal numerical representations of engines, called customer decks. These customer decks, which are delivered to airframe companies by various U.S. engine companies, numerically characterize an engine's performance as defined by the particular U.S. airframe manufacturer. Until recently, all numerical models were provided with a Fortran-compatible interface in compliance with the Society of Automotive Engineers (SAE) document AS681F, and data communication was performed via a standard, labeled common structure in compliance with AS681F. Recently, the SAE committee began to develop a new standard: AS681G. AS681G addresses multiple language requirements for customer decks along with alternative data communication techniques. Along with the SAE committee, the NPSS Steady-State Cycle Deck project team developed a standard Application Program Interface (API) supported by a graphical user interface. This work will result in Aerospace Recommended Practice 4868 (ARP4868). The Steady-State Cycle Deck work was validated against the Energy Efficient Engine customer deck, which is publicly available. The Energy Efficient Engine wrapper was used not only to validate ARP4868 but also to demonstrate how to wrap an existing customer deck. The graphical user interface for the Steady-State Cycle Deck facilitates the use of the new standard and makes it easier to design and analyze a customer deck. This software was developed following I. Jacobson's Object-Oriented Design methodology and is implemented in C++. The AS681G standard will establish a common generic interface for U.S. engine companies and airframe manufacturers. This will lead to more accurate cycle models, quicker model generation, and faster validation leading to specifications. The standard will facilitate cooperative work between industry and NASA. The NPSS Steady-State Cycle Deck team released a batch version of the Steady-State Cycle Deck in March 1996. Version 1.1 was released in June 1996. During fiscal 1997, NPSS accepted enhancements and modifications to the Steady-State Cycle Deck launcher. Consistent with NPSS' commercialization plan, these modifications will be done by a third party that can provide long-term software support.

Description: Field measurement of noise radiated from flight vehicles is an important element of aircraft noise research programs. At NASA Langley, a dedicated effort that spans over two decades was devoted to the development of acoustic measurement systems to support the NASA noise research programs. The new challenge for vehicle operational noise reduction through varying glide slope and flight path require noise measurement to be made over a very large area under the vehicle flight path. Such a challenge can be met through the digital remote system currently under final development at NASA Langley.

Description: The effect of varying the cross-sectional flow area of the passages of a wave rotor is examined by means of the method of characteristics. An idealized expansion wave, an idealized inlet port, and an idealized compression stage are considered. It is found that area variation does not have a very significant effect on the expansion wave, nor on the compression stage. For the expansion wave, increasing the passage area in the flow direction has the same effect as a diffuser, so that the flow emerges at a lower velocity than it would for the constant area case. This could be advantageous. The inlet is strongly affected by the area variation, as it changes the strength of the hammer shock wave, thereby changing the pressure behind it. In this case, reduction in the passage area in the flow direction leads to increased pressure. However, this result is dependent on the assumption that the inlet conditions remain constant with area variation. This may not be the case.

Description: This thesis presents an experimental study of swirling flows with direct applications to gas turbine combustors. Two separate flowfields were investigated: a round, swirling jet and a non-combusting annular combustor model. These studies were intended to allow both a further understanding of the behavior of general swirling flow characteristics, such as the recirculation zone, as well as to provide a base for the development of computational models. In order to determine the characteristics of swirling flows the concentration fields of a round, swirling jet were analyzed for varying amount of swirl. The experimental method used was a light scattering concentration measurement technique known as marker nephelometry. Results indicated the formation of a zone of recirculating fluid for swirl ratios (rotational speed x jet radius over mass average axial velocity) above a certain critical value. The size of this recirculation zone, as well as the spread angle of the jet, was found to increase with increase in the amount of applied swirl. The annular combustor model flowfield simulated the cold-flow characteristics of typical current annular combustors: swirl, recirculation, primary air cross jets and high levels of turbulence. The measurements in the combustor model made by the Laser Doppler Velocimetry technique, allowed the evaluation of the mean and rms velocities in the three coordinate directions, one Reynold's shear stress component and the turbulence kinetic energy: The primary cross jets were found to have a very strong effect on both the mean and turbulence flowfields. These cross jets, along with a large step change in area and wall jet inlet flow pattern, reduced the overall swirl in the test section to negligible levels. The formation of the strong recirculation zone is due mainly to the cross jets and the large step change in area. The cross jets were also found to drive a four-celled vortex-type motion (parallel to the combustor longitudinal axis) near the cross jet injection plane.

Description: A three-dimensional, linearized, Euler analysis is being developed to provide an efficient unsteady aerodynamic analysis that can be used to predict the aeroelastic and aeroacoustic responses of axial-flow turbo-machinery blading.The field equations and boundary conditions needed to describe nonlinear and linearized inviscid unsteady flows through a blade row operating within a cylindrical annular duct are presented. A numerical model for linearized inviscid unsteady flows, which couples a near-field, implicit, wave-split, finite volume analysis to a far-field eigenanalysis, is also described. The linearized aerodynamic and numerical models have been implemented into a three-dimensional linearized unsteady flow code, called LINFLUX. This code has been applied to selected, benchmark, unsteady, subsonic flows to establish its accuracy and to demonstrate its current capabilities. The unsteady flows considered, have been chosen to allow convenient comparisons between the LINFLUX results and those of well-known, two-dimensional, unsteady flow codes. Detailed numerical results for a helical fan and a three-dimensional version of the 10th Standard Cascade indicate that important progress has been made towards the development of a reliable and useful, three-dimensional, prediction capability that can be used in aeroelastic and aeroacoustic design studies.

Description: The concept of using a system, consisting of a tow aircraft, glider and tow line, which would enable subsonic flight at altitudes above 24 km (78 kft) has previously been investigated. The preliminary results from these studies seem encouraging. Under certain conditions these studies indicate the concept is feasible. However, the previous studies did not accurately take into account the forces acting on the tow line. Therefore in order to investigate the concept further a more detailed analysis was needed. The code that was selected was the SEADYN cable dynamics computer program which was developed at the Naval Facilities Engineering Service Center. The program is a finite element based structural analysis code that was developed over a period of 10 years. The results have been validated by the Navy in both laboratory and at actual sea conditions. This code was used to simulate arbitrarily-configured cable structures subjected to excitations encountered in real-world operations. The Navy's interest was mainly for modeling underwater tow lines, however the code is also usable for tow lines in air when the change in fluid properties is taken into account. For underwater applications the fluid properties are basically constant over the length of the tow line. For the tow aircraft/glider application the change in fluid properties is considerable along the length of the tow line. Therefore the code had to be modified in order to take into account the variation in atmospheric properties that would be encountered in this application. This modification consisted of adding a variable density to the fluid based on the altitude of the node being calculated. This change in the way the code handled the fluid density had no effect on the method of calculation or any other factor related to the codes validation.

Description: The effect of upstream blade row wake passing on the showerhead film cooling performance of a downstream turbine blade has been investigated through a combination of experimental and computational studies. The experiments were performed in a steady-flow annular turbine cascade facility equipped with an upstream rotating row of cylindrical rods to produce a periodic wake field similar to that found in an actual turbine. Spanwise, chordwise, and temporal resolution of the blade surface temperature were achieved through the use of an array of nickel thin-film surface gauges covering one unit cell of showerhead film hole pattern. Film effectiveness and Nusselt number values were determined for a test matrix of various injectants, injectant blowing ratios, and wake Strouhal numbers. Results indicated a demonstratable reduction in film effectiveness with increasing Strouhal number, as well as the expected increase in film effectiveness with blowing ratio. An equation was developed to correlate the span-average film effectiveness data. The primary effect of wake unsteadiness was found to be correlated well by a chordwise-constant decrement of 0.094-St. Measurable spanwise film effectiveness variations were found near the showerhead region, but meaningful unsteady variations and downstream spanwise variations were not found. Nusselt numbers were less sensitive to wake and injection changes. Computations were performed using a three-dimensional turbulent Navier-Stokes code which was modified to model wake passing and film cooling. Unsteady computations were found to agree well with steady computations provided the proper time-average blowing ratio and pressure/suction surface flow split are matched. The remaining differences were isolated to be due to the enhanced mixing in the unsteady solution caused by the wake sweeping normally on the pressure surface. Steady computations were found to be in excellent agreement with experimental Nusselt numbers, but to overpredict experimental film effectiveness values. This is likely due to the inability to match actual hole exit velocity profiles and the absence of a credible turbulence model for film cooling.

Description: In recent years, environmental regulations have become more stringent, requiring lower emissions of mainly nitrogen oxides (NOx), as well as carbon monoxide (CO) and unburned hydrocarbons (UHC). These regulations have forced the gas turbine industry to examine non-conventional combustion strategies, such as the lean burn approach. The reasoning behind operating under lean conditions is to maintain the temperature of combustion near and below temperatures required for the formation of thermal nitric oxide (NO). To be successful, however, the lean processes require careful preparation of the fuel/air mixture to preclude formation of either locally rich reaction zones, which may give rise to NO formation, or locally lean reaction zones, which may give rise to inefficient fuel processing. As a result fuel preparation is crucial to the development and success of new aeroengine combustor technologies. A key element of the fuel preparation process is the fuel nozzle. As nozzle technologies have developed, airblast atomization has been adopted for both industrial and aircraft gas turbine applications. However, the majority of the work to date has focused on prefilming nozzles, which despite their complexity and high cost have become an industry standard for conventional combustion strategies. It is likely that the new strategies required to meet future emissions goals will utilize novel fuel injector approaches, such as radial injection. This thesis proposes and demonstrates an experiment to examine, on a mechanistic level (i.e., the physics of the action), the processes associated with the atomization, evaporation, and dispersion of a liquid jet introduced, from a radial, plain-jet airblast injector, into a crossflow of air. This understanding requires the knowledge not only of what factors influence atomization, but also the underlying mechanism associated with liquid breakup and dispersion. The experimental data acquired identify conditions and geometries for improved performance of radial airblast injectors.

Description: Results of an analytical and experimental investigation, directed at the understanding of the impact of periodic unsteadiness on the time-averaged flows in axial flow turbomachines, are presented. Analysis of available experimental data, from a large-scale rotating rig (LSRR) (low speed rig), shows that in the time-averaged axisymmetric equations the magnitude of the terms representing the effect of periodic unsteadiness (deterministic stresses) are as large or larger than those due to random unsteadiness (turbulence). Numerical experiments, conducted to highlight physical mechanisms associated with the migration of combustor generated hot-streaks in turbine rotors, indicated that the effect can be simulated by accounting for deterministic stress like terms in the time-averaged mass and energy conservation equations. The experimental portion of this program shows that the aerodynamic loss for the second stator in a 1-1/2 stage turbine are influenced by the axial spacing between the second stator leading edge and the rotor trailing edge. However, the axial spacing has little impact on the heat transfer coefficient. These performance changes are believed to be associated with the change in deterministic stress at the inlet to the second stator. Data were also acquired to quantify the impact of indexing the first stator relative to the second stator. For the range of parameters examined, this effect was found to be of the same order as the effect of axial spacing.

Description: The low-NO(x) emitting potential of rich-burn/quick-mix/lean-burn )RQL) combustion makes it an attractive option for engines of future stratospheric aircraft. Because NO(x) formation is exponentially dependent on temperature, the success of the RQL combustor depends on minimizing high temperature stoichiometric pocket formation in the quick-mixing section. An experiment was designed and built, and tests were performed to characterize reaction and mixing properties of jets issuing from round orifices into a hot, fuel-rich crossflow confined in a cylindrical duct. The reactor operates on propane and presents a uniform, non-swirling mixture to the mixing modules. Modules consisting of round orifice configurations of 8, 9, 10, 12, 14, and 18 holes were evaluated at a momentum-flux ratio of 57 and jet-to-mainstream mass-flaw ratio of 2.5. Temperatures and concentrations of O2, CO2, CO, HC, and NO(x) were obtained upstream, down-stream, and within the orifice plane to determine jet penetration as well as reaction processes. Jet penetration was a function of the number of orifices and affected the mixing in the reacting system. Of the six configurations tested, the 14-hole module produced jet penetration close to the module half-radius and yielded the best mixing and most complete combustion at a plane one duct diameter from the orifice leading edge. The results reveal that substantial reaction and heat release occur in the jet mixing zone when the entering effluent is hot and rich, and that the experiment as designed will serve to explore satisfactorily jet mixing behavior under realistic reacting conditions in future studies.

Description: A procedure has been developed for predicting peak dynamic inlet distortion. This procedure combines Computational Fluid Dynamics (CFD) and distortion synthesis analysis to obtain a prediction of peak dynamic distortion intensity and the associated instantaneous total pressure pattern. A prediction of the steady state total pressure pattern at the Aerodynamic Interface Plane is first obtained using an appropriate CFD flow solver. A corresponding inlet turbulence pattern is obtained from the CFD solution via a correlation linking root mean square (RMS) inlet turbulence to a formulation of several CFD parameters representative of flow turbulence intensity. This correlation was derived using flight data obtained from the NASA High Alpha Research Vehicle flight test program and several CFD solutions at conditions matching the flight test data. A distortion synthesis analysis is then performed on the predicted steady state total pressure and RMS turbulence patterns to yield a predicted value of dynamic distortion intensity and the associated instantaneous total pressure pattern.

Description: The objective of the program was to determine a wave rotor demonstrator engine concept using the Allison 250 series engine. The results of the NASA LERC wave rotor effort were used as a basis for the wave rotor design. A wave rotor topped gas turbine engine was identified which incorporates five basic requirements of a successful demonstrator engine. Predicted performance maps of the wave rotor cycle were used along with maps of existing gas turbine hardware in a design point study. The effects of wave rotor topping on the engine cycle and the subsequent need to rematch compressor and turbine sections in the topped engine were addressed. Comparison of performance of the resulting engine is made on the basis of wave rotor topped engine versus an appropriate baseline engine using common shaft compressor hardware. The topped engine design clearly demonstrates an impressive improvement in shaft horsepower (+11.4%) and SFC (-22%). Off design part power engine performance for the wave rotor topped engine was similarly improved including that at engine idle conditions. Operation of the engine at off design was closely examined with wave rotor operation at less than design burner outlet temperatures and rotor speeds. Challenges identified in the development of a demonstrator engine are discussed. A preliminary design was made of the demonstrator engine including wave rotor to engine transition ducts. Program cost and schedule for a wave rotor demonstrator engine fabrication and test program were developed.

Description: A convertible engine called the CEST TF34, using the variable inlet guide vane method of power change, was tested on an outdoor stand at the NASA Lewis Research Center with a waterbrake dynamometer for the shaft load. A new digital electronic system, in conjunction with a modified standard TF34 hydromechanical fuel control, kept engine operation stable and safely within limits. All planned testing was completed successfully. Steady-state performance and acoustic characteristics were reported previously and are referenced. This report presents results of transient and dynamic tests. The transient tests measured engine response to several rapid changes in thrust and torque commands at constant fan (shaft) speed. Limited results from dynamic tests using the pseudorandom binary noise technique are also presented. Performance of the waterbrake dynamometer is discussed in an appendix.

Description: Theoretical analyses and computer codes are being developed for predicting compressible unsteady inviscid and viscous flows through blade rows of axial-flow turbomachines. Such analyses are needed to determine the impact of unsteady flow phenomena on the structural durability and noise generation characteristics of the blading. The emphasis has been placed on developing analyses based on asymptotic representations of unsteady flow phenomena. Thus, high Reynolds number flows driven by small amplitude unsteady excitations have been considered. The resulting analyses should apply in many practical situations and lead to a better understanding of the relevant flow physics. In addition, they will be efficient computationally, and therefore, appropriate for use in aeroelastic and aeroacoustic design studies. Under the present effort, inviscid interaction and linearized inviscid unsteady flow models have been formulated, and inviscid and viscid prediction capabilities for subsonic steady and unsteady cascade flows have been developed. In this report, we describe the linearized inviscid unsteady analysis, LINFLO, the steady inviscid/viscid interaction analysis, SFLOW-IVI, and the unsteady viscous layer analysis, UNSVIS. These analyses are demonstrated via application to unsteady flows through compressor and turbine cascades that are excited by prescribed vortical and acoustic excitations and by prescribed blade vibrations. Recommendations are also given for the future research needed for extending and improving the foregoing asymptotic analyses, and to meet the goal of providing efficient inviscid/viscid interaction capabilities for subsonic and transonic unsteady cascade flows.

Description: This report describes the design of the Advanced Low-Noise Research Fan stage. The fan is a variable pitch design, which is designed at the cruise pitch condition. Relative to the cruise setting, the blade is closed at takeoff and opened for reverse thrust operation. The fan stage is a split flow design with fan exit guide vanes (FEGVs) and core stators. The fan stage design is combined with a nacelle and engine core duct to form a powered fan/nacelle subscale model. This model is intended for use in combined aerodynamic, acoustic, and structural testing in a wind tunnel. The fan has an outer diameter of 22 in. and a hub-to-tip of 0.426 in., which allows the use of existing NASA fan and cowl force balance and rig drive systems. The design parameters were selected to permit valid acoustic and aerodynamic comparisons with the Pratt & Whitney (P&W) 17- and 22-in. rigs previously tested under NASA contract. The fan stage design is described in detail. The results of the design axisymmetric and Navier-Stokes aerodynamic analysis are presented at the critical design conditions. The structural analysis of the fan rotor and attachment is included. The blade and attachment are predicted to have adequate low-cycle fatigue life and an acceptable operating range without resonant stress or flutter. The stage was acoustically designed with airfoil counts in the FEGV and core stator to minimize noise. A fan/FEGV tone analysis developed separately under NASA contract was used to determine the optimum airfoil counts. The fan stage was matched to the existing nacelle, designed under the previous P&W low-noise contract, to form a fan/nacelle model for wind tunnel testing. It is an axisymmetric nacelle for convenience in testing and analysis. Previous testing confirmed that the nacelle performed as required at various aircraft operating conditions.

Description: The goal of this program was to evaluate the low cycle fatigue behavior of an SCS-6/Ti-6Al-4V sub-component under bi-axial loading conditions at 316 C(600 F). A simulated bladed TMC disk was designed having thirty four blades representing the number that would be used in Allied Signal's JTAGG II impeller. The outer diameter of the bladed ring was 254 mm (10.0 inch) and the inner diameter 114.3 mm (4.50 inch). The outer and inner diameter of the composite zone was 177.8 mm (7.00 inch) and 127.O mm(5.00 inch) respectively. Stress analysis showed that the fatigue life of the bladed composite ring would be about 12000 cycles for the test conditions applied. A modal analysis was conducted which showed that the blades would have sufficient life margin from dynamic excitation. The arbor design was the same as that employed in the spin-to burst test of NAS3-27027. A systematic stress analysis of each part making up the arbor was undertaken to assure the design would meet the low cycle fatigue requirements of the program. The Textron Systems grooved foil-fiber process was chosen to make the SCS-6/Ti-6Al-4V core ring based on the success they had in contract NAS3-27027. Fiber buckling, however, was observed at several locations in the first ring made which rendered it unsuitable for spin testing. The fiber buckling was attributed to cracking of the graphite tooling during the consolidation process. On this basis a second ring was made but it too contained fiber buckling defects. Analysis by Textron indicated that the fiber buckling was most likely due to poor placement of the SCS-6 fiber in the etched grooves of the Ti-6Al-4V foil. This was also a contributor to the defects in the first ring. Since there was little indication of control in the process to manufacture a quality ring a third attempt at making a ring was not undertaken.

Description: Ballistic impact tests were conducted on fabric made from both Poly(phenylene benzobizoxazole) (PBO) and Kevlar 29 which were selected to be similar in weave pattern, areal density, and fiber denier. The projectiles were 2.54-cm- (1-in.-) long aluminum cylinders with a diameter of 1.27 cm (0.5 in.). The fabric specimens were clamped on four sides in a 30.5-cm- (12-in.-) square frame. Tests on PBO were conducted at room temperature and at 260 C (500 F). A number of PBO specimens were aged in air at 204 and 260 C (400 and 500 F) before impact testing. Kevlar specimens were tested only at room temperature and with no aging. The PBO absorbed significantly more energy than the Kevlar at both room and elevated temperatures. However, after aging at temperatures of 204 C (400 F) and above, the PBO fabric lost almost all of its energy absorbing ability. It was concluded that PBO fabric is not a feasible candidate for fan containment system applications in supersonic jet engines where operating temperatures exceed this level.

Description: Integrators are included in the feedback loop of a control system to eliminate the steady state errors in the commanded variables. The integrator windup problem arises if the control actuators encounter operational limits before the steady state errors are driven to zero by the integrator. The typical effects of windup are large system oscillations, high steady state error, and a delayed system response following the windup. In this study, methods to prevent the integrator windup are examined to provide Integrator Windup Protection (IW) for an engine controller of a Short Take-Off and Vertical Landing (STOVL) aircraft. An unified performance index is defined to optimize the performance of the Conventional Anti-Windup (CAW) and the Modified Anti-Windup (MAW) methods. A modified Genetic Algorithm search procedure with stochastic parameter encoding is implemented to obtain the optimal parameters of the CAW scheme. The advantages and drawbacks of the CAW and MAW techniques are discussed and recommendations are made for the choice of the IWP scheme, given some characteristics of the system.

Description: The Advanced Film Cooling Rig (AFCR) test was a program to develop a thermal imaging camera system to measure the effectiveness of multihole film cooling. This system was used to evaluate low-thermal conductivity ceramic test panels incorporating a wide range of multihole cooling geometries. Tests were conducted with a mainstream flow of 600 F at Mach numbers of 0.37 and 1.57. Coolant flow was controlled to give a blowing parameter variation of up to 2.9. Data reduction techniques are given for camera calibration, rig temperature reflection correction, and deriving surface temperatures from video images. Results are presented in terms of overall and film cooling effectiveness distributions. A one-dimensional analysis of a unit film cooling hole was performed to generate film temperature.

Description: The exhaust flow properties (mass flow, pressure, temperature, velocity, and Mach number) of the F110-GE-129 engine in an F-16XL airplane were determined from a series of flight tests flown at NASA Dryden Flight Research Center, Edwards, California. These tests were performed in conjunction with NASA Langley Research Center, Hampton, Virginia (LARC) as part of a study to investigate the acoustic characteristics of jet engines operating at high nozzle pressure conditions. The range of interest for both objectives was from Mach 0.3 to Mach 0.9. NASA Dryden flew the airplane and acquired and analyzed the engine data to determine the exhaust characteristics. NASA Langley collected the flyover acoustic measurements and correlated these results with their current predictive codes. This paper describes the airplane, tests, and methods used to determine the exhaust flow properties and presents the exhaust flow properties. No acoustics results are presented.

Description: The higher temperature and pressure cycles of future aviation gas turbine combustors challenge designers to produce combustors that minimize their environmental impact while maintaining high operation efficiency. The development of low emissions combustors includes the reduction of unburned hydrocarbons, smoke, and particulates, as well as the reduction of oxides of nitrogen (NO(x)). In order to better understand and control the mechanisms that produce emissions, tools are needed to aid the development of combustor hardware. Current methods of measuring species within gas turbine combustors use extractive sampling of combustion gases to determine major species concentrations and to infer the bulk flame temperature. These methods cannot be used to measure unstable combustion products and have poor spatial and temporal resolution. The intrusive nature of gas sampling may also disturb the flow structure within a combustor. Planar laser-induced fluorescence (PLIF) is an optical technique for the measurement of combustion species. In addition to its non-intrusive nature, PLIF offers these advantages over gas sampling: high spatial resolution, high temporal resolution, the ability to measure unstable species, and the potential to measure combustion temperature. This thesis considers PLIF for in-situ visualization of combustion species as a tool for the design and evaluation of gas turbine combustor subcomponents. This work constitutes the first application of PLIF to the severe environment found in liquid-fueled, aviation gas turbine combustors. Technical and applied challenges are discussed. PLIF of OH was used to observe the flame structure within the post flame zone of a flame tube combustor, and within the flame zone of a sector combustor, for a variety of fuel injector configurations. OH was selected for measurement because it is a major combustion intermediate, playing a key role in the chemistry of combustion, and because its presence within the flame zone can serve as a qualitative marker of flame temperature. All images were taken in the environment of actual engines during flight, using actual jet fuel. The results of the PLIF study led directly to the modification of a fuel injector.

Description: The NASA Numerical Propulsion System Simulation (NPSS) project is exploring the use of computer simulation to facilitate the design of new jet engines. Several key issues raised in this research are being examined in an NPSS-related research project: zooming, monitoring and control, and support for heterogeneity. The design of a simulation executive that addresses each of these issues is described. In this work, the strategy of zooming, which allows codes that model at different levels of fidelity to be integrated within a single simulation, is applied to the fan component of a turbofan propulsion system. A prototype monitoring and control system has been designed for this simulation to support experimentation with expert system techniques for active control of the simulation. An interconnection system provides a transparent means of connecting the heterogeneous systems that comprise the prototype.

Description: An experimental study was conducted (1) to experimentally measure, assess and analyze the heat transfer within the internal cooling configuration of a radial turbine rotor blade and (2) to obtain heat transfer data to evaluate and improve computational fluid dynamics (CFD) procedures and turbulent transport models of internal coolant flows. A 1.15 times scale model of the coolant passages within the NASA LERC High Temperature Radial Turbine was designed, fabricated of Lucite and instrumented for transient beat transfer tests using thin film surface thermocouples and liquid crystals to indicate temperatures. Transient heat transfer tests were conducted for Reynolds numbers of one-fourth, one-half, and equal to the operating Reynolds number for the NASA Turbine. Tests were conducted for stationary and rotating conditions with rotation numbers in the range occurring in the NASA Turbine. Results from the experiments showed the heat transfer characteristics within the coolant passage were affected by rotation. In general, the heat transfer increased and decreased on the sides of the straight radial passages with rotation as previously reported from NASA-HOST-sponsored experiments. The heat transfer in the tri-passage axial flow region adjacent to the blade exit was relatively unaffected by rotation. However, the heat transfer on one surface, in the transitional region between the radial inflow passage and axial, constant radius passages, decreased to approximately 20 percent of the values without rotation. Comparisons with previous 3-D numerical studies indicated regions where the heat transfer characteristics agreed and disagreed with the present experiment.

Description: The flow in a planar shear layer of hydrogen reacting with hot air was measured with a two-component laser Doppler velocimeter (LDV) system, a schlieren system, and OH fluorescence imaging. It was compared with a similar air-to-air case without combustion. The high-speed stream's flow speed was about 390 m/s, or Mach 0.71, and the flow speed ratio was 0.34. The results showed that a shear layer with reaction grows faster than one without; both cases are within the range of data scatter presented by the established data base. The coupling between the streamwise and the cross-stream turbulence components inside the shear layers was low, and reaction only increased it slightly. However, the shear layer shifted laterally into the lower speed fuel stream, and a more organized pattern of Reynolds stress was present in the reaction shear layer, likely as a result of the formation of a larger scale structure associated with shear layer corrugation from heat release. Dynamic pressure measurements suggest that coherent flow perturbations existed inside the shear layer and that this flow became more chaotic as the flow advected downstream. Velocity and thermal variable values are listed in this report for a computational fluid dynamics (CFD) benchmark.

Description: This investigation summarizes a comparative study of two high-speed engine performance assessment techniques based on energy (available work) and thrust-potential (thrust availability). Simple flow-fields utilizing Rayleigh heat addition and one-dimensional flow with friction are used to demonstrate the fundamental inability of conventional energy techniques to predict engine component performance, aid in component design, or accurately assess flow losses. The use of the thrust-based method on these same examples demonstrates its ability to yield useful information in all these categories. Energy and thrust are related and discussed from the stand-point of their fundamental thermodynamic and fluid dynamic definitions in order to explain the differences in information obtained using the two methods. The conventional definition of energy is shown to include work which is inherently unavailable to an aerospace Brayton engine. An engine-based energy is then developed which accurately accounts for this inherently unavailable work; performance parameters based on this quantity are then shown to yield design and loss information equivalent to the thrust-based method.

Description: This work addresses the significant differences in compressor rotor wake mixing loss which exist in a stage environment relative to a rotor in isolation. The wake decay for a rotor in isolation is due solely to viscous dissipation which is an irreversible process and thus leads to a loss in both total pressure and efficiency. Rotor wake decay in the stage environment is due to both viscous mixing and the inviscid strain imposed on the wake fluid particles by the stator velocity field. This straining process, referred to by Smith (1993) as recovery, is reversible and for a 2D rotor wake leads to an inviscid reduction of the velocity deficit of the wake. A model for the rotor wake decay process is developed and used to quantify the viscous dissipation effects relative to those of inviscid wake stretching. The model is verified using laser anemometer measurements acquired in the wake of a transonic rotor operated in isolation and in a stage configuration at near peak efficiency and near stall operating conditions. Additional insight is provided by a time-accurate 3D Navier-Stokes simulation of the compressor stator flow field at the corresponding stage loading levels. Results from the wake decay model exhibit good agreement with the experimental data. Data from the model, laser anemometer measurements, and numerical simulations indicate that for the rotor/stator spacing used in this work, which is typical of core compressors, rotor wake straining (stretching) is the primary decay process in the stator passage with viscous mixing playing only a minor role. The implications of these results on compressor stage design are discussed.

Description: This report compiles the various research activities conducted under the auspices of the NASA Grant NAG3-1026, "Numerical Investigation of Hot Gas Ingestion by STOVL Aircraft" during the period of April 1989 to April 1994. The effort involved the development of multigrid based algorithms and computer programs for the calculation of the flow and temperature fields generated by Short Take-off and Vertical Landing (STOVL) aircraft, while hovering in ground proximity. Of particular importance has been the interaction of the exhaust jets with the head wind which gives rise to the hot gas ingestion process. The objective of new STOVL designs to reduce the temperature of the gases ingested into the engine. The present work describes a solution algorithm for the multi-dimensional elliptic partial-differential equations governing fluid flow and heat transfer in general curvilinear coordinates. The solution algorithm is based on the multigrid technique which obtains rapid convergence of the iterative numerical procedure for the discrete equations. Initial efforts were concerned with the solution of the Cartesian form of the equations. This algorithm was applied to a simulated STOVL configuration in rectangular coordinates. In the next phase of the work, a computer code for general curvilinear coordinates was constructed. This was applied to model STOVL geometries on curvilinear grids. The code was also validated in model problems. In all these efforts, the standard k-Epsilon model was used.

Description: The broad scope of aeropropulsion multidisciplinary applications necessitates that a collection of approaches, with distinct capabilities, be developed. Three general approaches to multidisciplinary simulations have been identified. The three approaches; loosely coupled, coupled process, and multiphysics, provide a comprehensive collection of capabilities for multidisciplinary aeropropulsion analysis. At the data access level, or loosely coupled approach of coupling, existing disciplinary simulations are run, data is generated, and made available and used for subsequent analysis. The data must be in the correct format for implementation by the subsequent analysis but the subsequent code need not directly communicate with the previous code. At the process level, or coupled process approach of coupling, individual disciplinary codes are used, similarly to the loosely coupled approach, however, in the coupled process approach the disciplinary codes need to run concurrently with each other. The system of equation coupled approach, or multiphysics approach, addresses those applications whose characteristics require that the disciplines be coupled at the fundamental equation level to accurately, or more efficiently, capture the multidisciplinary physics of the problem. No one of these approaches, by itself, addresses all of the community needs in this area. However, collectively the three approaches encompass all of the multidisciplinary applications which have been identified thus far. Multiple approaches to multidisciplinary simulations will be needed as long as the applications and their requirements remain as diverse as they currently are today.

Description: This contact had two main objectives involving both numerical and experimental investigations of a small highly loaded two-stage axial compressor designated Advanced Small Turboshaft Compressor (ASTC) winch had a design pressure ratio goal of 5:1 at a flowrate of 10.53 lbm/s. The first objective was to conduct 3-D Navier Stokes multistage analyses of the ASTC using several different flow modelling schemes. The second main objective was to complete a numerical/experimental investigation into stall range enhancement of the ASTC. This compressor was designed wider a cooperative Space Act Agreement and all testing was completed at NASA Lewis Research Center. For the multistage analyses, four different flow model schemes were used, namely: (1) steady-state ADPAC analysis, (2) unsteady ADPAC analysis, (3) steady-state APNASA analysis, and (4) steady state OCOM3D analysis. The results of all the predictions were compared to the experimental data. The steady-state ADPAC and APNASA codes predicted similar overall performance and produced good agreement with data, however the blade row performance and flowfield details were quite different. In general, it can be concluded that the APNASA average-passage code does a better job of predicting the performance and flowfield details of the highly loaded ASTC compressor.

Description: The objective of this effort is to develop an analytical model for the coupling of active noise control (ANC) piston-type actuators that are mounted flush to the inner and outer walls of an annular duct to the modes in the duct generated by the actuator motion. The analysis will be used to couple the ANC actuators to the modal analysis propagation computer program for the annular duct, to predict the effects of active suppression of fan-generated engine noise sources. This combined program will then be available to assist in the design or evaluation of ANC systems in fan engine annular exhaust ducts. An analysis has been developed to predict the modes generated in an annular duct due to the coupling of flush-mounted ring actuators on the inner and outer walls of the duct. The analysis has been combined with a previous analysis for the coupling of modes to a cylindrical duct in a FORTRAN computer program to perform the computations. The method includes the effects of uniform mean flow in the duct. The program can be used for design or evaluation purposes for active noise control hardware for turbofan engines. Predictions for some sample cases modeled after the geometry of the NASA Lewis ANC Fan indicate very efficient coupling in both the inlet and exhaust ducts for the m = 6 spinning mode at frequencies where only a single radial mode is cut-on. Radial mode content in higher order cut-off modes at the source plane and the required actuator displacement amplitude to achieve 110 dB SPL levels in the desired mode were predicted. Equivalent cases with and without flow were examined for the cylindrical and annular geometry, and little difference was found for a duct flow Mach number of 0.1. The actuator ring coupling program will be adapted as a subroutine to the cylindrical duct modal analysis and the exhaust duct modal analysis. This will allow the fan source to be defined in terms of characteristic modes at the fan source plane and predict the propagation to the arbitrarily-located ANC source plane. The actuator velocities can then be determined to generate the anti-phase mode. The resulting combined fan source/ANC pressure can then be calculated at any desired wall sensor position. The actuator velocities can be determined manually or using a simulation of a control system feedback loop. This will provide a very useful ANC system design and evaluation tool.

Description: An extension of a prior study has been completed to examine the potential reduction of aircraft flyover noise by the method of active noise control (ANC). It is assumed that the ANC system will be designed such that it cancels discrete tones radiating from the engine fan inlet or fan exhaust duct, at least to the extent that they no longer protrude above the surrounding broadband noise levels. Thus, without considering the engineering details of the ANC system design, tone levels am arbitrarily removed from the engine component noise spectrum and the flyover noise EPNL levels are compared with and without the presence of tones. The study was conducted for a range of engine cycles, corresponding to fan pressure ratios of 1.3, 1.45, 1.6, and 1.75. This report is an extension of an effort reported previously. The major conclusions drawn from the prior study, which was restricted to fan pressure ratios of 1.45 and 1.75, are that, for a fan pressure ratio of 1.75, ANC of tones gives about the same suppression as acoustic treatment without ANC. For a fan pressure ratio of 1.45, ANC appears to offer less effectiveness from passive treatment. In the present study, the other two fan pressure ratios are included in a more detailed examination of the benefits of the ANC suppression levels. The key results of this extended study are the following observations: (1) The maximum overall benefit obtained from suppression of BPF alone was 2.5 EPNdB at high fan speeds. The suppression benefit increases with increase in fan pressure ratio (FPR), (2) The maximum overall benefit obtained from suppression of the first three harmonics was 3 EPNdB at high speeds. Suppression benefit increases with increase in FPR, (3) At low FPR, only about 1.0 EPNdB maximum reduction was obtained. Suppression is primarily from reduction of BPF at high FPR values and from the combination of tones at low FPR, (4) The benefit from ANC is about the same as the benefit from passive treatment at fan pressure ratios of 1.75 and 1.60. At the two lower fan pressure ratios, the effectivness of treatment is much greater than that of ANC, and (5) No significant difference in ANC suppression behavior was found from the QCSEE engine database analysis compared to that of the E3 engine database, for the FPR = 1.3 engine cycle. The effects of ANC on EPNL noise reduction are difficult to generalize. It was found that the reduction obtained in any particular case depended upon the frequency of the tones and their shift with rpm, the amount of ANC suppression received by each tone (which depended on its protrusion from the background), and the NOY-value of the tone relative to the NOY-value of other tones and the peak broadband levels, because PNL is determined from the sum of the NOY-values.

Description: Experiments were performed on a low-speed multistage axial-flow compressor to assess the effects of shrouded stator cavity flows on aerodynamic performance. Five configurations, which involved changes in seal-tooth leakage rates and/or elimination of the shrouded stator cavities, were tested. Data collected enabled differences in overall individual stage and the third stage blade element performance parameters to be compared. The results show conclusively that seal-tooth leakage ran have a large impact on compressor aerodynamic performance while the presence of the shrouded stator cavities alone seemed to have little influence. Overall performance data revealed that for every 1% increase in the seal-tooth clearance to blade-height ratio the pressure rise dropped up to 3% while efficiency was reduced by 1 to 1.5 points. These observed efficiency penalty slopes are comparable to those commonly reported for rotor and cantilevered stator tip clearance variations. Therefore, it appears that in order to correctly predict overall performance it is equally important to account for the effects of seal-tooth leakage as it is to include the influence of tip clearance flows. Third stage blade element performance data suggested that the performance degradation observed when leakage was increased was brought about in two distinct ways. First, increasing seal-tooth leakage directly spoiled the near hub performance of the stator row in which leakage occurred. Second, the altered stator exit now conditions caused by increased leakage impaired the performance of the next downstream stage by decreasing the work input of the downstream rotor and increasing total pressure loss of the downstream stator. These trends caused downstream stages to progressively perform worse. Other measurements were acquired to determine spatial and temporal flow field variations within the up-and-downstream shrouded stator cavities. Flow within the cavities involved low momentum fluid traveling primarily in the circumferential direction at about 40% of the hub wheel speed. Measurements indicated that the flow within both cavities was much more complex than first envisioned. A vortical flow structure in the meridional plane, similar to a driven cavity, existed within the upstream cavity Furthermore, other spatial and temporal variations in Row properties existed. the most prominent being caused by the upstream potential influence of the downstream blade. This influence caused the fluid within cavities near the leading edges of either stator blades in space or rotor blades in time to be driven radially inward relative to fluid near blade mid-pitch. This influence also produced large unsteady velocity fluctuations in the downstream cavity because of the passing of the downstream rotor blade.

Description: The effect of spanwise-periodic mean-flow distortions (i.e. streamwise-vortex structures) on the evolution of small-amplitude, single-frequency instability waves in an otherwise two-dimensional shear flow is investigated. The streamwise-vortex structures are taken to be just weak enough so that the spatially growing instability waves behave (locally) like linear perturbations about a uni-directional transversely sheared mean flow. Numerical solutions are computed and discussed for both the mean flow and the instability waves. The influence of the streamwise-vortex wavelength on the properties of the most rapidly growing instability wave is also discussed.

Description: A nonintrusive concentration measurement method is developed for determining the concentration distribution in a complex flow field. The measurement method consists of marking a liquid flow with a water soluble fluorescent dye. The dye is excited by a two dimensional sheet of laser light. The fluorescent intensity is shown to be proportional to the relative concentration level. The fluorescent field is recorded on a video cassette recorder through a video camera. The recorded images are analyzed with image processing hardware and software to obtain intensity levels. Mean and root mean square (rms) values are calculated from these intensity levels. The method is tested on a single round turbulent jet because previous concentration measurements have been made on this configuration by other investigators. The previous results were used to comparison to qualify the current method. These comparisons showed that this method provides satisfactory results. 'Me concentration measurement system was used to measure the concentrations in the complex flow field of a model gas turbine annular combustor. The model annular combustor consists of opposing primary jets and an annular jet which discharges perpendicular to the primary jets. The mixing between the different jet flows can be visualized from the calculated mean and rms profiles. Concentration field visualization images obtained from the processing provide further qualitative information about the flow field.

Description: A series of non-reacting parametric experiments was conducted to investigate the effect of geometric and flow variations on mixing of cold jets in an axis-symmetric, heated cross flow. The confined, cylindrical geometries tested represent the quick mix region of a Rich-Burn/Quick-Mix/Lean-Burn (RQL) combustor. The experiments show that orifice geometry and jet to mainstream momentum-flux ratio significantly impact the mixing characteristic of jets in a cylindrical cross stream. A computational code was used to extrapolate the results of the non-reacting experiments to reacting conditions in order to examine the nitric oxide (NO) formation potential of the configurations examined. The results show that the rate of NO formation is highest immediately downstream of the injection plane. For a given momentum-flux ratio, the orifice geometry that mixes effectively in both the immediate vicinity of the injection plane, and in the wall regions at downstream locations, has the potential to produce the lowest NO emissions. The results suggest that further study may not necessarily lead to a universal guideline for designing a low NO mixer. Instead, an assessment of each application may be required to determine the optimum combination of momentum-flux ratio and orifice geometry to minimize NO formation. Experiments at reacting conditions are needed to verify the present results.

Description: This paper presents the results of a computational study on the effect of axial spacing between the vane and blade rows of a transonic turbine stage. The study was performed on the mid-span section of a high-pressure turbine stage using a quasi-3D, unsteady Navier-Stokes solver that provides a fully interactive vane-blade unsteady flow solution. Three different cases were considered, corresponding to axial spacings of 20%, 40%, and 60% of the vane axial chord. The calculated vane and blade pressure distributions for the 40 percent case were found to compare favorably with experimental measurements acquired in a short-duration shock tunnel. In addition, the analysis shows a marked increase in the amplitude of the unsteady pressure fluctuations on the vane and blade surfaces as the spacing decreases. Time-averaged stage adiabatic efficiency predictions for each case are presented to show the effect of spacing on aerodynamic performance.

Description: Mixing of gaseous jets in a cross-flow has significant applications in engineering, one example of which is the dilution zone of a gas turbine combustor. Despite years of study, the design of the jet injection in combustors is largely based on practical experience. The emergence of NO(x) regulations for stationary gas turbines and the anticipation of aero-engine regulations requires an improved understanding of jet mixing as new combustor concepts are introduced. For example, the success of the staged combustor to reduce the emission of NO(x) is almost entirely dependent upon the rapid and complete dilution of the rich zone products within the mixing section. It is these mixing challenges to which the present study is directed. A series of experiments was undertaken to delineate the optimal mixer orifice geometry. A cross-flow to core-flow momentum-flux ratio of 40 and a mass flow ratio of 2.5 were selected as representative of a conventional design. An experimental test matrix was designed around three variables: the number of orifices, the orifice length-to- width ratio, and the orifice angle. A regression analysis was performed on the data to arrive at an interpolating equation that predicted the mixing performance of orifice geometry combinations within the range of the test matrix parameters. Results indicate that the best mixing orifice geometry tested involves eight orifices with a long-to-short side aspect ratio of 3.5 at a twenty-three degree inclination from the center-line of the mixing section.

Description: Staged combustion, such as Rich-Burn/Quick-Mix/Lean-Burn (RQL), is a viable strategy to meet nitric oxide (NO) emission goals for both stationary and propulsion gas turbine engines. A critical element of the design is the quick mixer section where the potential for NO production is high. While numerical calculations of the quick mixer under reacting conditions have been conducted, the hostile environment and lack of appropriate diagnostics have, to date, precluded experimental probing of the reacting case. As an alternative to understanding the effect of geometry and flow variations on the production of NO in the quick mixer, the present paper presents (1) a series of non-reacting parametric studies, and (2) a computational method to extrapolate the results of the non-reacting experiments to reacting conditions. The results show that the rate of NO production is highest in the immediate vicinity of the injection plane. For a given momentum flux ratio between the jets and mainstream, the most effective mixing geometry is that which mixes effectively in both (1) the plane of injection, and (2) the wall regions downstream of the plan of injection. The tailoring of the mixing is key to minimize the NO formed. As a result, the best overall mixer with respect to the minimization of NO production may depend on the system specific characteristics of the particular application.

Description: The spread of a flame over solid fuel is not only a fundamental textbook combustion phenomenon, but also the central element of destructive fires that cause tragic loss of life and property each year. Throughout history, practical measures to prevent and fight fires have been developed, but these have often been based on lessons learned in a costly fire. Since the 1960 s, scientists and engineers have employed powerful tools of scientific research to understand the details of flame spread and how a material can be rendered nonflammable. High-speed computers have enabled complex flame simulations, whereasand lasers have provided measurements of the chemical composition, temperature, and air velocities inside flames. The microgravity environment has emerged as the third great tool for these studies. Spreading flames are complex combinations of chemical reactions and several physical processes including the transport of oxygen and fuel vapor to the flame and the transfer of heat from the flame to fresh fuel and to the surroundings. Depending on its speed, air motion in the vicinity of the flame can affect the flame in substantially different ways. For example, consider the difference between blowing on a campfire and blowing out a match. On Earth, gravity induces air motion because of buoyancy (the familiar rising hot gases); this process cannot be controlled experimentally. For theoreticians, buoyant air motion complicates the problem modeling of flame spread beyond the capacity of modern computers to simulate. The microgravity environment provides experimental control of air motion near spreading flames, with results that can be compared with detailed theory. The Solid Surface Combustion Experiment (SSCE) was designed to obtain benchmark flame spreading data in quiescent test atmospheres--the limiting case of flames spreading. Professor Robert Altenkirch, Vice President for Research at Mississippi State University, proposed the experiment concept, and the NASA Lewis Research Center designed, built, and tested the SSCE hardware. It was the first microgravity science experiment built by Lewis for the space shuttle and the first combustion science experiment flown in space.

Description: With the advent of new, more stringent noise regulations in the next century, aircraft engine manufacturers are investigating new technologies to make the current generation of aircraft engines as well as the next generation of advanced engines quieter without sacrificing operating performance. A current NASA initiative called the Advanced Subsonic Technology (AST) Program has set as a goal a 6-EPNdB (effective perceived noise) reduction in aircraft engine noise relative to 1992 technology levels by the year 2000. As part of this noise program, and in cooperation with the Allison Engine Company, an advanced, low-noise, high-bypass-ratio fan stage design and several advanced technology stator vane designs were recently tested in NASA Lewis Research Center's 9- by 15-Foot Low-Speed Wind Tunnel (an anechoic facility). The project was called the NASA/Allison Low Noise Fan.

Description: A Two-Dimensional Bifurcated (2DB) Inlet was successfully tested in NASA Lewis Research Center s 10- by 10-Foot Supersonic Wind Tunnel. These tests were the culmination of a collaborative effort between the Boeing Company, General Electric, Pratt & Whitney, and Lewis. Extensive support in-house at Lewis contributed significantly to the progress and accomplishment of this test. The results, which met or exceeded many of the High-Speed Research (HSR) Program goals, were used to revise system studies within the HSR Program. The HSR Program is focused on developing low-noise, low-polluting, high-efficiency supersonic commercial aircraft. A supersonic inlet is an important component of an efficient, low-noise vehicle.

Description: NASA s Advanced Subsonic Technology (AST) Program seeks to develop new technologies to increase the fuel efficiency of commercial aircraft engines, improve the safety of engine operation, and reduce emissions and engine noise. For new designs of ducted fans, compressors, and turbines to achieve these goals, a basic aeroelastic requirement is that there should be no flutter or high resonant blade stresses in the operating regime. For verifying the aeroelastic soundness of the design, an accurate prediction/analysis code is required. Such a three-dimensional viscous propulsion aeroelastic code, named TURBO-AE, is being developed at the NASA Lewis Research Center. The development and verification of the flutter version of the TURBO-AE code (version 4) has been completed. Validation of the code is partially complete.

Description: The NASA Lewis Research Center is developing an environment for analyzing and designing aircraft engines-the Numerical Propulsion System Simulation (NPSS). NPSS will integrate multiple disciplines, such as aerodynamics, structure, and heat transfer, and will make use of numerical "zooming" on component codes. Zooming is the coupling of analyses at various levels of detail. NPSS uses the latest computing and communication technologies to capture complex physical processes in a timely, cost-effective manner. The vision of NPSS is to create a "numerical test cell" enabling full engine simulations overnight on cost-effective computing platforms. Through the NASA/Industry Cooperative Effort agreement, NASA Lewis and industry partners are developing a new engine simulation called the National Cycle Program (NCP). NCP, which is the first step toward NPSS and is its initial framework, supports the aerothermodynamic system simulation process for the full life cycle of an engine. U.S. aircraft and airframe companies recognize NCP as the future industry standard common analysis tool for aeropropulsion system modeling. The estimated potential payoff for NCP is a $50 million/yr savings to industry through improved engineering productivity.

Description: A new forward-swept rotor designed by Allison Engine Company was tested in NASA Lewis Research Center's CE-18 facility. This testing was a follow-on project sponsored by NASA Lewis to study range enhancements in small turbomachinery. The test was conducted against a baseline rotor design that was also tested in CE-18. The design point for the rotor was a rotor pressure ratio of 2.69, a mass flow of 10.52 lbm/sec, and an adiabatic efficiency of 89.1 percent. Test data indicate that the rotor met the pressure ratio of 2.69 with a 10.77 lbm/sec flow rate, a 87.5-percent adiabatic efficiency, and a 19.5-percent stall margin. The baseline rotor achieved a pressure ratio of 2.69 at a 10.77 lbm/sec flow rate with a stall margin of only 9.2 percent and an adiabatic efficiency of 87.0 percent. The major differences are the significant stall margin increase and the substantially higher off-design peak efficiencies of the forward-swept rotor. The substantially higher performance over the baseline rotor design makes the new design a viable technology candidate for future products.

Description: NASA's Advanced Subsonic Technology (AST) program is developing new technologies to increase the fuel efficiency of commercial aircraft engines, improve the safety of engine operation, and reduce engine emissions and noise. With the development of new designs for ducted fans, compressors, and turbines to achieve these goals, a basic aeroelastic requirement is that there should be no flutter or high resonant blade stresses in the operating regime. To verify the aeroelastic soundness of these designs, we need an accurate prediction and analysis code. Such a two-dimensional viscous propulsion aeroelastic code, named TURBO-AE, is being developed at the NASA Lewis Research Center. The TURBO-AE aeroelastic code is based on a three-dimensional unsteady aerodynamic Euler/Navier-Stokes turbomachinery code TURBO, developed under a grant from NASA Lewis. TURBO-AE can model viscous flow effects that play an important role in certain aeroelastic problems, such as flutter with flow separation (or stall flutter) and flutter in the presence of shock and boundary-layer interaction. The structural dynamics representation of the blade in the TURBO-AE code is based on a normal mode representation. A finite element analysis code, such as NASTRAN, is used to calculate in-vacuum vibration modes and the associated natural frequency. A work-per-cycle approach is used to determine aeroelastic (flutter) stability. With this approach, the motion of the blade is prescribed to be a harmonic vibration in a specified in vacuum normal mode. The aerodynamic forces acting on the vibrating blade and the work done by these forces on the vibrating blade during a cycle of vibration are calculated. If positive work is being done on the blade by the aerodynamic forces, the blade is dynamically unstable, since it will extract energy from the flow, leading to an increase in the amplitude of the blade's oscillation. Initial calculations have been done for a configuration representative of the Energy Efficient Engine fan rotor. The accompanying figure shows the work-per-cycle after each cycle of vibration. It can be seen that the work-per-cycle does not vary much after the fourth cycle. The negative sign of the converged work-per-cycle shows that the fan blade is dynamically stable and will not flutter. TURBO-AE will provide a useful aeroelastic prediction/analysis capability for engine manufacturers. It will reduce design cycle times by allowing new blade designs to be verified for aeroelastic soundness before blades are built and tested. With this prediction capability, it will be possible to build thinner, lighter, and faster rotating blades without encountering aeroelastic problems like stall flutter and high-cycle fatigue due to forced vibrations.

Description: The relative circumferential location of stator rows (stator indexing) is an aspect of multistage compressor design that has not yet been explored for its potential impact on compressor aerodynamic performance. Although the inlet stages of multistage compressors usually have differing stator blade counts, the aft stages of core compressors can often have stage blocks with equal stator blade counts in successive stages. The potential impact of stator indexing is likely greatest in these stages. To assess the performance impact of stator indexing, researchers at the NASA Lewis Research Center used the 4 ft diameter, four-stage NASA Low Speed Axial Compressor for detailed experiments. This compressor has geometrically identical stages that can circumferentially index stator rows relative to each other in a controlled manner; thus it is an ideal test rig for such investigations.

Description: A wave rotor is a device that can boost the pressure and temperature of an airflow. When used as part of the core of a gas turbine engine, a wave rotor can significantly improve the thrust or shaft horsepower by boosting the flow pressure without raising the turbine inlet temperature. The NASA Lewis Research Center's Aeropropulsion Analysis Office, which is identifying technologies and research opportunities that will enhance the technical and economic competitiveness of the U.S. aeronautics industry, is evaluating the wave rotor to quantify the potential benefits of this device. Preliminary studies such as these are critical to identifying technologies that have high payoffs.

Description: Most helicopters now in service have a transmission with a planetary design. Studies have shown that some helicopters would be lighter and more reliable if they had a transmission with a split-torque design instead. However, a split-torque design has never been used by a U.S. helicopter manufacturer because there has been no proven method to ensure equal sharing of the load among the multiple load paths. The Sikorsky/Boeing team has chosen to use a split-torque transmission for the U.S. Army's Comanche helicopter, and Sikorsky Aircraft is designing and manufacturing the transmission. To help reduce the technical risk of fielding this helicopter, NASA and the Army have done the research jointly in cooperation with Sikorsky Aircraft. A theory was developed that equal load sharing could be achieved by proper configuration of the geartrain, and a computer code was completed in-house at the NASA Lewis Research Center to calculate this optimal configuration.

Description: A Navier-Stokes computer code is used to predict one of the ducted-fan engine acoustic modes that results from rotor-wake/stator-blade interaction. A patched sliding-zone interface is employed to pass information between the moving rotor row and the stationary stator row. The code produces averaged aerodynamic results downstream of the rotor that agree well with a widely used average-passage code. The acoustic mode of interest is generated successfully by the code and is propagated well upstream of the rotor; temporal and spatial numerical resolution are fine enough such that attenuation of the signal is small. Two acoustic codes are used to find the far-field noise. Near-field propagation is computed by using Eversman's wave envelope code, which is based on a finite-element model. Propagation to the far field is accomplished by using the Kirchhoff formula for moving surfaces with the results of the wave envelope code as input data. Comparison of measured and computed far-field noise levels show fair agreement in the range of directivity angles where the peak radiation lobes from the inlet are observed. Although only a single acoustic mode is targeted in this study, the main conclusion is a proof-of-concept: Navier-Stokes codes can be used both to generate and propagate rotor/stator acoustic modes forward through an engine, where the results can be coupled to other far-field noise prediction codes.

Description: Steady deflection measurements were taken of two identical NASA/Pratt & Whitney-designed fan blades while they were rotating in a vacuum in NASA Lewis Research Center's Dynamic Spin Facility. The one-fifth-scale fan blades, which have a tip diameter of 22 in. and a pinroot retention, are of sparshell construction and were unducted for this test. The purpose of the test was to measure the change of the radial deflection of the blade tip and blade angle at selected radial stations along the blade span with respect to rotational speed. The procedure for radial deflection measurement had no precedent and was newly developed for this test. Radial deflection measurements were made to assure adequate tip clearance existed between the fan blades and the duct for a follow-on wind tunnel test. Also, blade angle deflection measurements were desired before pitchsetting parts for the wind tunnel test were finish machined. During the test, laser beams were aimed across the blade path into photodiodes to give signals that were used to determine blade angle change or tip radial deflection. These laser beams were set parallel to the spin axis at selected radial stations.

Description: A wave rotor, which acts as a high-technology topping spool in gas turbine engines, can increase the effective pressure ratio as well as the turbine inlet temperature in such engines. The wave rotor topping, in other words, may significantly enhance engine performance by increasing shaft horse power while reducing specific fuel consumption. This performance enhancement requires optimum selection of the wave rotor's adjustable parameters for speed, surge margin, and temperature constraints specified on different engine components. To examine the benefit of the wave rotor concept in engine design, researchers soft coupled NASA Lewis Research Center's multidisciplinary optimization tool COMETBOARDS and the NASA Engine Performance Program (NEPP) analyzer. The COMETBOARDS-NEPP combined design tool has been successfully used to optimize wave-rotor-topped engines. For illustration, the design of a subsonic gas turbine wave-rotor-enhanced engine with four ports for 47 mission points (which are specified by Mach number, altitude, and power-setting combinations) is considered. The engine performance analysis, constraints, and objective formulations were carried out through NEPP, and COMETBOARDS was used for the design optimization. So that the benefits that accrue from wave rotor enhancement could be examined, most baseline variables and constraints were declared to be passive, whereas important parameters directly associated with the wave rotor were considered to be active for the design optimization. The engine thrust was considered as the merit function. The wave rotor engine design, which became a sequence of 47 optimization subproblems, was solved successfully by using a cascade strategy available in COMETBOARDS. The graph depicts the optimum COMETBOARDS solutions for the 47 mission points, which were normalized with respect to standard results. As shown, the combined tool produced higher thrust for all mission points than did the other solution, with maximum benefits around mission points 11, 25, and 31. Such improvements can become critical, especially when engines are sized for these specific mission points.