ALBERT

Description: System studies have shown the benefits of reducing blade tip clearances in modern turbine engines. Minimizing blade tip clearances throughout the engine will contribute materially to meeting NASA s Ultra-Efficient Engine Technology (UEET) turbine engine project goals. NASA GRC is examining two candidate approaches including rub-avoidance and regeneration which are explained in subsequent slides.

Description: As the aviation industry moves toward higher efficiency electrical power generation, all electric aircraft, or zero emissions and more quiet aircraft, fuel cells are sought as the technology that can deliver on these high expectations. The hybrid solid oxide fuel cell system combines the fuel cell with a micro-turbine to obtain up to 70% cycle efficiency, and then distributes the electrical power to the loads via a power distribution system. The challenge is to understand the dynamics of this complex multidiscipline system and the design distributed controls that take the system through its operating conditions in a stable and safe manner while maintaining the system performance. This particular system is a power generation and a distribution system, and the fuel cell and micro-turbine model fidelity should be compatible with the dynamics of the power distribution system in order to allow proper stability and distributed controls design. The novelty in this paper is that, first, the case is made why a high fidelity fuel cell mode is needed for systems control and stability designs. Second, a novel modeling approach is proposed for the fuel cell that will allow the fuel cell and the power system to be integrated and designed for stability, distributed controls, and other interface specifications. This investigation shows that for the fuel cell, the voltage characteristic should be modeled but in addition, conservation equation dynamics, ion diffusion, charge transfer kinetics, and the electron flow inherent impedance should also be included.

Description: The usage and integrated vehicle health management of the NASA C-17. Propulsion health management flight objectives for the aircraft include mapping of the High Pressure Compressor in order to calibrate a Pratt and Whitney engine model and the fusion of data collected from existing sensors and signals to develop models, analysis methods and information fusion algorithms. An additional health manage flight objective is to demonstrate that the Commercial Modular Aero-Propulsion Systems Simulation engine model can successfully execute in real time onboard the C-17 T-1 aircraft using engine and aircraft flight data as inputs. Future work will address aircraft durability and aging, airframe health management, and propulsion health management research in the areas of gas path and engine vibration.

Description: A vane has an airfoil shell and a spar within the shell. The vane has an outboard shroud at an outboard end of the shell and an inboard platform at an inboard end of the shell. The spar has a first chamber essentially along the suction side and a second chamber along the pressure side opposite the first chamber.

Description: Key aspects of the design of sealing systems for On Rotor Combustion/Wave Rotor (ORC/WR) systems were addressed. ORC/WR systems generally fit within a broad class of pressure gain Constant Volume Combustors (CVCs) or Pulse Detonation Combustors (PDCs) which are currently being considered for use in many classes of turbine engines for dramatic efficiency improvement. Technology readiness level of this ORC/WR approaches are presently at 2.0. The results of detailed modeling of an ORC/WR system as applied to a regional jet engine application were shown to capture a high degree of pressure gain capabilities. The results of engine cycle analysis indicated the level of specific fuel consumption (SFC) benefits to be 17 percent. The potential losses in pressure gain due to leakage were found to be closely coupled to the wave processes at the rotor endpoints of the ORC/WR system. Extensive investigation into the sealing approaches is reported. Sensitivity studies show that SFC gains of 10 percent remain available even when pressure gain levels are highly penalized. This indicates ORC/WR systems to have a high degree of tolerance to rotor leakage effects but also emphasizes their importance. An engine demonstration of an ORC/WR system is seen as key to progressing the TRL of this technology. An industrial engine was judged to be a highly advantageous platform for demonstration of a first generation ORC/WR system. Prior to such a demonstration, the existing NASA pressure exchanger wave rotor rig was identified as an opportunity to apply both expanded analytical modeling capabilities developed within this program and to identify and fix identified leakage issues existing within this rig. Extensive leakage analysis of the rig was performed and a detailed design of additional sealing strategies for this rig was generated.

Description: Computational fluid dynamics (CFD) analysis has been performed to study the plume effects on sonic boom signature for isolated nozzle configurations. The objectives of these analyses were to provide comparison to past work using modern CFD analysis tools, to investigate the differences of high aspect ratio nozzles to circular (axisymmetric) nozzles, and to report the effects of underexpanded nozzle operation on boom signature. CFD analysis was used to address the plume effects on sonic boom signature from a baseline exhaust nozzle. Near-field pressure signatures were collected for nozzle pressure ratios (NPRs) between 6 and 10. A computer code was used to extrapolate these signatures to a ground-observed sonic boom N-wave. Trends show that there is a reduction in sonic boom N-wave signature as NPR is increased from 6 to 10. The performance curve for this supersonic nozzle is flat, so there is not a significant loss in thrust coefficient as the NPR is increased. As a result, this benefit could be realized without significant loss of performance. Analyses were also collected for a high aspect ratio nozzle based on the baseline design for comparison. Pressure signatures were collected for nozzle pressure ratios from 8 to 12. Signatures were nearly twice as strong for the two-dimensional case, and trends also show a reduction in sonic boom signature as NPR is increased from 8 to 12. As low boom designs are developed and improved, there will be a need for understanding the interaction between the aircraft boat tail shocks and the exhaust nozzle plume. These CFD analyses will provide a baseline study for future analysis efforts.

Description: Helicopter Health Usage Monitoring Systems (HUMS) have potential for providing data to support increasing the service life of a dynamic mechanical component in the transmission of a helicopter. Data collected can demonstrate the HUMS condition indicator responds to a specific component fault with appropriate alert limits and minimal false alarms. Defining thresholds for specific faults requires a tradeoff between the sensitivity of the condition indicator (CI) limit to indicate damage and the number of false alarms. A method using Receiver Operating Characteristic (ROC) curves to assess CI performance was demonstrated using CI data collected from accelerometers installed on several UH60 Black Hawk and AH64 Apache helicopters and an AH64 helicopter component test stand. Results of the analysis indicate ROC curves can be used to reliably assess the performance of commercial HUMS condition indicators to detect damaged gears and bearings in a helicopter transmission.

Description: Flaps (or half wedges) attached to the sides of a pylon are shown to result in a small but clear noise benefit. Noise radiated towards the ground is reduced apparently through a deflection and thickening of the fan stream underneath. Based on results from the current as well as concurrent investigations at the University of California at Irvine, it is recommended that further tests in a larger facility simulating realistic engine conditions be considered.

Description: The objective is to provide turbine-cooling technologies to meet Propulsion 21 goals related to engine fuel burn, emissions, safety, and reliability. Specifically, the GE Aviation (GEA) Advanced Turbine Cooling and Thermal Management program seeks to develop advanced cooling and flow distribution methods for HP turbines, while achieving a substantial reduction in total cooling flow and assuring acceptable turbine component safety and reliability. Enhanced cooling techniques, such as fluidic devices, controlled-vortex cooling, and directed impingement jets, offer the opportunity to incorporate both active and passive schemes. Coolant heat transfer enhancement also can be achieved from advanced designs that incorporate multi-disciplinary optimization of external film and internal cooling passage geometry.

Description: The objective of the Advanced Turbine Cooling and Thermal Management program is to develop intelligent control and distribution methods for turbine cooling, while achieving a reduction in total cooling flow and assuring acceptable turbine component safety and reliability. The program also will develop embedded sensor technologies and cooling system models for real-time engine diagnostics and health management. Both active and passive control strategies will be investigated that include the capability of intelligent modulation of flow quantities, pressures, and temperatures both within the supply system and at the turbine component level. Thermal management system concepts were studied, with a goal of reducing HPT blade cooling air supply temperature. An assessment will be made of the use of this air by the active clearance control system as well. Turbine component cooling designs incorporating advanced, high-effectiveness cooling features, will be evaluated. Turbine cooling flow control concepts will be studied at the cooling system level and the component level. Specific cooling features or sub-elements of an advanced HPT blade cooling design will be downselected for core fabrication and casting demonstrations.

Description: For the Intelligent Engine System (Propulsion 21) study, each technology was evaluated to determine the impact to fuel burn, acoustics, and NOx emissions. The optimum combination of technologies and their overall benefits to the system were also evaluated, resulting in noise improvement potential of 1.89 EPNdB cumulative margin,-1.34 percent fuel burn, and 50 percent NOx reduction from the 2015 UEET-QAT baseline. All the technology evaluations, except T18-20D, were based on newengines, where the engine was resized to obtain the maximum system benefit while maintaining the same cycle parameters as the 2015 UEET-QAT baseline. The impact of turbine clearance control on deteriorated engines, T18-20D, was also evaluated. Recommendations for future system study work include, but were not limited to, validation of a university-developed engine deterioration model and customer value analysis as figures of merit beside fuel burn, emissions, and acoustics.

Description: An adaptive controls method for instability suppression in gas turbine engine combustors has been developed and successfully tested with a realistic aircraft engine combustor rig. This testing was part of a program that demonstrated, for the first time, successful active combustor instability control in an aircraft gas turbine engine-like environment. The controls method is called Adaptive Sliding Phasor Averaged Control. Testing of the control method has been conducted in an experimental rig with different configurations designed to simulate combustors with instabilities of about 530 and 315 Hz. Results demonstrate the effectiveness of this method in suppressing combustor instabilities. In addition, a dramatic improvement in suppression of the instability was achieved by focusing control on the second harmonic of the instability. This is believed to be due to a phenomena discovered and reported earlier, the so called Intra-Harmonic Coupling. These results may have implications for future research in combustor instability control.

Description: Low-noise fan exit guide vanes are disclosed. According to the present invention a fan exit guide vane has an outer shell substantially shaped as an airfoil and defining an interior cavity. A porous portion of the outer shell allows communication between the fluctuations in the air passing over the guide vane and the interior cavity. At least one acoustically resonant chamber is located within the interior cavity. The resonant chamber is in communication with the porous portion of the outer perimeter. The resonant chamber is configured to reduce the noise generated at a predetermined frequency. In various preferred embodiments, there is a plurality of acoustically resonant chambers located within the interior cavity. The resonant chambers can be separated by one or more partitions within the interior cavity. In these embodiments, the resonant chambers can be configured to reduce the noise generated over a range of predetermined frequencies.

Description: Pulsed combustion is receiving renewed interest as a potential route to higher performance in air breathing propulsion systems. Pulsejets offer a simple experimental device with which to study unsteady combustion phenomena and validate simulations. Previous computational fluid dynamic (CFD) simulation work focused primarily on the pulsejet combustion and exhaust processes. This paper describes a new inlet sub-model which simulates the fluidic and mechanical operation of a valved pulsejet head. The governing equations for this sub-model are described. Sub-model validation is provided through comparisons of simulated and experimentally measured reed valve motion, and time averaged inlet mass flow rate. The updated pulsejet simulation, with the inlet sub-model implemented, is validated through comparison with experimentally measured combustion chamber pressure, inlet mass flow rate, operational frequency, and thrust. Additionally, the simulated pulsejet exhaust flowfield, which is dominated by a starting vortex ring, is compared with particle imaging velocimetry (PIV) measurements on the bases of velocity, vorticity, and vortex location. The results show good agreement between simulated and experimental data. The inlet sub-model is shown to be critical for the successful modeling of pulsejet operation. This sub-model correctly predicts both the inlet mass flow rate and its phase relationship with the combustion chamber pressure. As a result, the predicted pulsejet thrust agrees very well with experimental data.

Description: Mission Support Features: a) Shirtsleeve environment, . 18 scientists; b) worldwide deployment experience; c) Extensive modifications to support in-situ and remote sensing instruments 1) zenith and nadir viewports; 2) modified power systems; 3) 19 inch rack mounting; 4) on-board data acquisition network.

Description: The Model Based Fault Tolerant Control (MBFTC) task was conducted under the NASA Aviation Safety and Security Program. The goal of MBFTC is to develop and demonstrate real-time strategies to diagnose and accommodate anomalous aircraft engine events such as sensor faults, actuator faults, or turbine gas-path component damage that can lead to in-flight shutdowns, aborted take offs, asymmetric thrust/loss of thrust control, or engine surge/stall events. A suite of model-based fault detection algorithms were developed and evaluated. Based on the performance and maturity of the developed algorithms two approaches were selected for further analysis: (i) multiple-hypothesis testing, and (ii) neural networks; both used residuals from an Extended Kalman Filter to detect the occurrence of the selected faults. A simple fusion algorithm was implemented to combine the results from each algorithm to obtain an overall estimate of the identified fault type and magnitude. The identification of the fault type and magnitude enabled the use of an online fault accommodation strategy to correct for the adverse impact of these faults on engine operability thereby enabling continued engine operation in the presence of these faults. The performance of the fault detection and accommodation algorithm was extensively tested in a simulation environment.

Description: An assessment was made of the capability of jet noise prediction codes over a broad range of jet flows, with the objective of quantifying current capabilities and identifying areas requiring future research investment. Three separate codes in NASA s possession, representative of two classes of jet noise prediction codes, were evaluated, one empirical and two statistical. The empirical code is the Stone Jet Noise Module (ST2JET) contained within the ANOPP aircraft noise prediction code. It is well documented, and represents the state of the art in semi-empirical acoustic prediction codes where virtual sources are attributed to various aspects of noise generation in each jet. These sources, in combination, predict the spectral directivity of a jet plume. A total of 258 jet noise cases were examined on the ST2JET code, each run requiring only fractions of a second to complete. Two statistical jet noise prediction codes were also evaluated, JeNo v1, and Jet3D. Fewer cases were run for the statistical prediction methods because they require substantially more resources, typically a Reynolds-Averaged Navier-Stokes solution of the jet, volume integration of the source statistical models over the entire plume, and a numerical solution of the governing propagation equation within the jet. In the evaluation process, substantial justification of experimental datasets used in the evaluations was made. In the end, none of the current codes can predict jet noise within experimental uncertainty. The empirical code came within 2dB on a 1/3 octave spectral basis for a wide range of flows. The statistical code Jet3D was within experimental uncertainty at broadside angles for hot supersonic jets, but errors in peak frequency and amplitude put it out of experimental uncertainty at cooler, lower speed conditions. Jet3D did not predict changes in directivity in the downstream angles. The statistical code JeNo,v1 was within experimental uncertainty predicting noise from cold subsonic jets at all angles, but did not predict changes with heating of the jet and did not account for directivity changes at supersonic conditions. Shortcomings addressed here give direction for future work relevant to the statistical-based prediction methods. A full report will be released as a chapter in a NASA publication assessing the state of the art in aircraft noise prediction.

Description: A concept for mitigating the adverse effects of jet vorticity and liftoff at high blowing ratios for turbine film cooling flows has been developed and studied at NASA Glenn Research Center. This "anti-vortex" film cooling concept proposes the addition of two branched holes from each primary hole in order to produce a vorticity counter to the detrimental kidney vortices from the main jet. These vortices typically entrain hot freestream gas and are associated with jet separation from the turbine blade surface. The anti-vortex design is unique in that it requires only easily machinable round holes, unlike shaped film cooling holes and other advanced concepts. The anti-vortex film cooling hole concept has been modeled computationally for a single row of 30deg angled holes on a flat surface using the 3D Navier-Stokes solver Glenn-HT. A modification of the anti-vortex concept whereby the branched holes exit adjacent to the main hole has been studied computationally for blowing ratios of 1.0 and 2.0 and at density ratios of 1.0 and 2.0. This modified concept was selected because it has shown the most promise in recent experimental studies. The computational results show that the modified design improves the film cooling effectiveness relative to the round hole baseline and previous anti-vortex cases, in confirmation of the experimental studies.

Description: Recent technology reviews have identified the need for objective assessments of engine health management (EHM) technology. The need is two-fold: technology developers require relevant data and problems to design and validate new algorithms and techniques while engine system integrators and operators need practical tools to direct development and then evaluate the effectiveness of proposed solutions. This paper presents a publicly available gas path diagnostic benchmark problem that has been developed by the Propulsion and Power Systems Panel of The Technical Cooperation Program (TTCP) to help address these needs. The problem is coded in MATLAB (The MathWorks, Inc.) and coupled with a non-linear turbofan engine simulation to produce "snap-shot" measurements, with relevant noise levels, as if collected from a fleet of engines over their lifetime of use. Each engine within the fleet will experience unique operating and deterioration profiles, and may encounter randomly occurring relevant gas path faults including sensor, actuator and component faults. The challenge to the EHM community is to develop gas path diagnostic algorithms to reliably perform fault detection and isolation. An example solution to the benchmark problem is provided along with associated evaluation metrics. A plan is presented to disseminate this benchmark problem to the engine health management technical community and invite technology solutions.

Description: An accurate indication of available power is required for helicopter mission planning purposes. Available power is currently estimated on U.S. Army Blackhawk helicopters by performing a Maximum Power Check (MPC), a manual procedure performed by maintenance pilots on a periodic basis. The MPC establishes Engine Torque Factor (ETF), an indication of available power. It is desirable to replace the current manual MPC procedure with an automated approach that will enable continuous real-time assessment of available power utilizing normal mission data. This report presents an automated power assessment approach which processes data currently collected within helicopter Health and Usage Monitoring System (HUMS) units. The overall approach consists of: 1) a steady-state data filter which identifies and extracts steady-state operating points within HUMS data sets; 2) engine performance curve trend monitoring and updating; and 3) automated ETF calculation. The algorithm is coded in MATLAB (The MathWorks, Inc.) and currently runs on a PC. Results from the application of this technique to HUMS mission data collected from UH-60L aircraft equipped with T700-GE-701C engines are presented and compared to manually calculated ETF values. Potential future enhancements are discussed.

Description: In this paper, a baseline system which utilizes dual-channel sensor measurements for aircraft engine on-line diagnostics is developed. This system is composed of a linear on-board engine model (LOBEM) and fault detection and isolation (FDI) logic. The LOBEM provides the analytical third channel against which the dual-channel measurements are compared. When the discrepancy among the triplex channels exceeds a tolerance level, the FDI logic determines the cause of the discrepancy. Through this approach, the baseline system achieves the following objectives: (1) anomaly detection, (2) component fault detection, and (3) sensor fault detection and isolation. The performance of the baseline system is evaluated in a simulation environment using faults in sensors and components.

Description: Recent developments in gas foil bearing technology have led to numerous advanced high-speed rotating system concepts, many of which have become either commercial products or experimental test articles. Examples include Oil-Free microturbines, motors, generators and turbochargers. The driving forces for integrating gas foil bearings into these high-speed systems are the benefits promised by removing the oil lubrication system. Elimination of the oil system leads to reduced emissions, increased reliability, and decreased maintenance costs. Another benefit is reduced power plant weight. For rotorcraft applications, this would be a major advantage, as every pound removed from the propulsion system results in a payload benefit. Implementing foil gas bearings throughout a rotorcraft gas turbine engine is an important long-term goal that requires overcoming numerous technological hurdles. Adequate thrust bearing load capacity and potentially large gearbox applied radial loads are among them. However, by replacing the turbine end, or hot section, rolling element bearing with a gas foil bearing many of the above benefits can be realized. To this end, engine manufacturers are beginning to explore the possibilities of hot section gas foil bearings in propulsion engines. This paper presents a logical follow-on activity by analyzing a conceptual rotorcraft engine to determine the feasibility of a foil bearing supported core. Using a combination of rotordynamic analyses and a load capacity model, it is shown to be reasonable to consider a gas foil bearing core section.

Description: A methodology for the design and construction of simple foil thrust bearings intended for parametric performance testing and low marginal costs is presented. Features drawn from a review of the open literature are discussed as they relate to bearing performance. The design of fixtures and tooling required to fabricate foil thrust bearings is presented, using conventional machining processes where possible. A prototype bearing with dimensions drawn from the literature is constructed, with all fabrication steps described. A load-deflection curve for the bearing is presented to illustrate structural stiffness characteristics. Start-top cycles are performed on the bearing at a temperature of 425 C to demonstrate early-life wear patterns. A test of bearing load capacity demonstrates useful performance when compared with data obtained from the open literature.

Description: CFD calculations using high-performance parallel computing were conducted to simulate the pre-stall flow of a transonic compressor stage, NASA compressor Stage 35. The simulations were run with a full-annulus grid that models the 3D, viscous, unsteady blade row interaction without the need for an artificial inlet distortion to induce stall. The simulation demonstrates the development of the rotating stall from the growth of instabilities. Pressure-rise performance and pressure traces are compared with published experimental data before the study of flow evolution prior to the rotating stall. Spatial FFT analysis of the flow indicates a rotating long-length disturbance of one rotor circumference, which is followed by a spike-type breakdown. The analysis also links the long-length wave disturbance with the initiation of the spike inception. The spike instabilities occur when the trajectory of the tip clearance flow becomes perpendicular to the axial direction. When approaching stall, the passage shock changes from a single oblique shock to a dual-shock, which distorts the perpendicular trajectory of the tip clearance vortex but shows no evidence of flow separation that may contribute to stall.

Description: A model has been developed to simulate a fixed-exit porous bleed system for supersonic inlets. The fixed-exit model allows the amount of bleed flow to vary according to local flow conditions and fixed-exit characteristics of the bleed system. This variation is important for the control of shock-wave/boundary-layer interactions within the inlet. The model computes the bleed plenum static pressure rather than requiring its specification. The model was implemented in the Wind-US computational fluid dynamics code. The model was then verified and validated against experimental data for bleed on a flat plate with and without an impinging oblique shock and for bleed in a Mach 3.0 axisymmetric, mixed-compression inlet. The model was able to accurately correlate the plenum pressures with bleed rates and simulate the effect of the bleed on the downstream boundary layer. Further, the model provided a realistic simulation of the initiation of inlet unstart. The results provide the most in-depth examination to date of bleed models for use in the simulation of supersonic inlets. The results also highlight the limitations of the models and aspects that require further research.

Description: Meeting future goals for aircraft and air traffic system performance will require new airframes with more highly integrated propulsion. Previous studies have evaluated hybrid wing body (HWB) configurations with various numbers of engines and with increasing degrees of propulsion-airframe integration. A recently published configuration with 12 small engines partially embedded in a HWB aircraft, reviewed herein, serves as the airframe baseline for the new concept aircraft that is the subject of this paper. To achieve high cruise efficiency, a high lift-to-drag ratio HWB was adopted as the baseline airframe along with boundary layer ingestion inlets and distributed thrust nozzles to fill in the wakes generated by the vehicle. The distributed powered-lift propulsion concept for the baseline vehicle used a simple, high-lift-capable internally blown flap or jet flap system with a number of small high bypass ratio turbofan engines in the airframe. In that concept, the engine flow path from the inlet to the nozzle is direct and does not involve complicated internal ducts through the airframe to redistribute the engine flow. In addition, partially embedded engines, distributed along the upper surface of the HWB airframe, provide noise reduction through airframe shielding and promote jet flow mixing with the ambient airflow. To improve performance and to reduce noise and environmental impact even further, a drastic change in the propulsion system is proposed in this paper. The new concept adopts the previous baseline cruise-efficient short take-off and landing (CESTOL) airframe but employs a number of superconducting motors to drive the distributed fans rather than using many small conventional engines. The power to drive these electric fans is generated by two remotely located gas-turbine-driven superconducting generators. This arrangement allows many small partially embedded fans while retaining the superior efficiency of large core engines, which are physically separated but connected through electric power lines to the fans. This paper presents a brief description of the earlier CESTOL vehicle concept and the newly proposed electrically driven fan concept vehicle, using the previous CESTOL vehicle as a baseline.

Description: The purpose of this cooperative agreement was to develop a foundation of intelligent propulsion technologies for NASA and industry that will have an impact on safety, noise, emissions, and cost. These intelligent engine technologies included sensors, electronics, communications, control logic, actuators, smart materials and structures, and system studies. Furthermore, this cooperative agreement helped prepare future graduates to develop the revolutionary intelligent propulsion technologies that will be needed to ensure pre-eminence of the U.S. aerospace industry. This Propulsion 21 - Phase 11 program consisted of four primary research areas and associated work elements at Ohio universities: 1.0 Turbine Engine Prognostics, 2.0 Active Controls for Emissions and Noise Reduction, 3.0 Active Structural Controls and Performance, and 4.0 System Studies and Integration. Phase l, which was conducted during the period August 1, 2003, through September 30, 2004, has been reported separately.

Description: The Advanced Thermally Actuated Clearance Control System underwent several studies. Improved flow path isolation quantified what can be gained by making the HPT case nearly adiabatic. The best method of heat transfer was established, and finally two different borrowed air cooling circuits were evaluated to be used for the HPT Active Clearance Control System.

Description: This presentation reviews recent progress made under NASA s Subsonic Rotary Wing (SRW) propulsion research activities. Advances in engines, drive systems and optimized propulsion systems are discussed. Progress in wide operability compressors, modeling of variable geometry turbine performance, foil gas bearings and multi-speed transmissions are presented.

Description: Far-field noise sound power level (PWL) spectra and overall sound pressure level (OASPL) directivities were compared for three significantly different model fan stages which were tested in the NASA Glenn 9 15 Low Speed Wind Tunnel. The test fans included the Advanced Ducted Propulsor (ADP) Fan1, the baseline Source Diagnostic Test (SDT) fan, and the Quiet High Speed Fan2 (QHSF2). These fans had design rotor tangential tip speeds from 840 to 1474 ft/s and stage pressure ratios from 1.29 to 1.82. Additional parameters included rotor-stator spacing, stator sweep, and downstream support struts. Acoustic comparison points were selected on the basis of stage thrust. Acoustic results for the low tip speed/low pressure ratio fan (ADP Fan1) were thrust-adjusted to show how a geometrically-scaled version of this fan might compare at the higher design thrust levels of the other two fans. Lowest noise levels were typically observed for ADP Fan1 (which had a radial stator) and for the intermediate tip speed fan (Source Diagnostics Test, SDT, R4 rotor) with a swept stator. Projected noise levels for the ADP fan to the SDT swept stator configuration at design point conditions showed the fans to have similar noise levels. However, it is possible that the ADP fan could be 2 to 3 dB quieter with incorporation of a swept stator. Benefits of a scaled ADP fan include avoidance of multiple pure tones associated with transonic and higher blade tip speeds. Penalties of a larger size ADP fan would include increased nacelle size and drag.

Description: Control of jet noise continues to be an important research topic. Exhaust-nozzle chevrons have been shown to reduce jet noise, but parametric effects are not well understood. Additionally, thrust loss due to chevrons at cruise suggests significant benefit from active chevrons. The focus of this study is development of an active chevron concept for the primary purpose of parametric studies for jet noise reduction in the laboratory and secondarily for technology development to leverage for full scale systems. The active chevron concept employed in this work consists of a laminated composite structure with embedded shape memory alloy (SMA) actuators, termed a SMA hybrid composite (SMAHC). SMA actuators are embedded on one side of the neutral axis of the structure such that thermal excitation, via joule heating, generates a moment and deflects the structure. The performance of two active chevron concepts is demonstrated in the presence of representative flow conditions. One of the concepts is shown to possess significant advantages for the proposed application and is selected for further development. Fabrication and design changes are described and shown to produce a chevron prototype that meets the performance objectives.

Description: Researchers at NASA Glenn Research Center have been investigating high temperature shape memory alloys as potential damping materials for turbomachinery rotor blades. Analysis shows that a thin layer of SMA with a loss factor of 0.04 or more would be effective at reducing the resonant response of a titanium alloy beam. Two NiTiHf shape memory alloy compositions were tested to determine their loss factors at frequencies from 0.1 to 100 Hz, at temperatures from room temperature to 300 C, and at alternating strain levels of 34-35x10(exp -6). Elevated damping was demonstrated between the M(sub s) and M(sub f) phase transformation temperatures and between the A(sub s) and A(sub f) temperatures. The highest damping occurred at the lowest frequencies, with a loss factor of 0.2-0.26 at 0.1 Hz. However, the peak damping decreased with increasing frequency, and showed significant temperature hysteresis in heating and cooling. Keywords: High-temperature, shape memory alloy, damping, aircraft engine blades, NiTiHf

Description: Ceramic thermal and environmental barrier coatings (TEBCs) are used in gas turbine engines to protect engine hot-section components in the harsh combustion environments, and extend component lifetimes. Advanced TEBCs that have significantly lower thermal conductivity, better thermal stability and higher toughness than current coatings will be beneficial for future low emission and high performance propulsion engine systems. In this paper, ceramic coating design and testing considerations will be described for turbine engine high temperature and high-heat-flux applications. Thermal barrier coatings for metallic turbine airfoils and thermal/environmental barrier coatings for SiC/SiC ceramic matrix composite (CMC) components for future supersonic aircraft propulsion engines will be emphasized. Further coating capability and durability improvements for the engine hot-section component applications can be expected by utilizing advanced modeling and design tools.

Description: A pressure-gain combustor comprised of a mechanically valved, liquid fueled pulsejet, an ejector, and an enclosing shroud, was coupled to a small automotive turbocharger to form a self-aspirating, thrust producing gas turbine engine. The system was constructed in order to investigate issues associated with the interaction of pulsed combustion devices and turbomachinery. Installed instrumentation allowed for sensing of distributed low frequency pressure and temperature, high frequency pressure in the shroud, fuel flow rate, rotational speed, thrust, and laboratory noise. The engine ran successfully and reliably, achieving a sustained thrust of 5 to 6 lbf, and maintaining a rotor speed of approximately 90,000 rpm, with a combustor pressure gain of approximately 4 percent. Numerical simulations of the system without pressure-gain combustion indicated that the turbocharger would not operate. Thus, the new combustor represented a substantial improvement in system performance. Acoustic measurements in the shroud and laboratory indicated turbine stage sound pressure level attenuation of 20 dB. This is consistent with published results from detonative combustion experiments. As expected, the mechanical reed valves suffered considerable damage under the higher pressure and thermal loading characteristics of this system. This result underscores the need for development of more robust valve systems for this application. The efficiency of the turbomachinery components did not appear to be significantly affected by unsteadiness associated with pulsed combustion, though the steady component efficiencies were already low, and thus not expected to be particularly sensitive.

Description: Reliable engine-weight estimation at the conceptual design stage is critical to the development of new aircraft engines. It helps to identify the best engine concept amongst several candidates. At NASA Glenn (GRC), the Weight Analysis of Turbine Engines (WATE) computer code, originally developed by Boeing Aircraft, has been used to estimate the engine weight of various conceptual engine designs. The code, written in FORTRAN, was originally developed for NASA in 1979. Since then, substantial improvements have been made to the code to improve the weight calculations for most of the engine components. Most recently, to improve the maintainability and extensibility of WATE, the FORTRAN code has been converted into an object-oriented version. The conversion was done within the NASA s NPSS (Numerical Propulsion System Simulation) framework. This enables WATE to interact seamlessly with the thermodynamic cycle model which provides component flow data such as airflows, temperatures, and pressures, etc. that are required for sizing the components and weight calculations. The tighter integration between the NPSS and WATE would greatly enhance system-level analysis and optimization capabilities. It also would facilitate the enhancement of the WATE code for next-generation aircraft and space propulsion systems. In this paper, the architecture of the object-oriented WATE code (or WATE++) is described. Both the FORTRAN and object-oriented versions of the code are employed to compute the dimensions and weight of a 300- passenger aircraft engine (GE90 class). Both versions of the code produce essentially identical results as should be the case. Keywords: NASA, aircraft engine, weight, object-oriented

Description: The NASA Glenn Research Center (GRC) is developing a high-power-density switched-reluctance cryogenic motor for all-electric and pollution-free flight. However, cryogenic operation at higher rotational speeds markedly shortens the life of mechanical rolling element bearings. Thus, to demonstrate the practical feasibility of using this motor for future flights, a non-contact rotor-bearing system is a crucial technology to circumvent poor bearing life that ordinarily accompanies cryogenic operation. In this paper, a bearingless motor control technology for a 12-8 (12 poles in the stator and 8 poles in the rotor) switched-reluctance motor operating in liquid nitrogen (boiling point, 77 K (-196 C or -321 F)) was presented. We pushed previous disciplinary limits of electromagnetic controller technique by extending the state-of-the-art bearingless motor operating at liquid nitrogen for high-specific-power applications. The motor was levitated even in its nonlinear region of magnetic saturation, which is believed to be a world first for the motor type. Also we used only motoring coils to generate motoring torque and levitation force, which is an important feature for developing a high specific power motor.

Description: The General Aviation Propulsion (GAP) Program Turbine Engine Element focused on the development of an advanced small turbofan engine. Goals were good fuel consumption and thrust-to-weight ratio, and very low production cost. The resulting FJX-2 turbofan engine showed the potential to meet all of these goals. The development of the engine was carried through to proof of concept testing of a complete engine system. The proof of concept engine was ground tested at sea level and in altitude test chambers. A turboprop derivative was also sea-level tested.

Description: A new hypersonic inlet for a turbine-based combined-cycle (TBCC) engine has been designed. This split-flow inlet is designed to provide flow to an over-under propulsion system with turbofan and dual-mode scramjet engines for flight from takeoff to Mach 7. It utilizes a variable-geometry ramp, high-speed cowl lip rotation, and a rotating low-speed cowl that serves as a splitter to divide the flow between the low-speed turbofan and the high-speed scramjet and to isolate the turbofan at high Mach numbers. The low-speed inlet was designed for Mach 4, the maximum mode transition Mach number. Integration of the Mach 4 inlet into the Mach 7 inlet imposed significant constraints on the low-speed inlet design, including a large amount of internal compression. The inlet design was used to develop mechanical designs for two inlet mode transition test models: small-scale (IMX) and large-scale (LIMX) research models. The large-scale model is designed to facilitate multi-phase testing including inlet mode transition and inlet performance assessment, controls development, and integrated systems testing with turbofan and scramjet engines.

Description: The primary objective of this effort was to demonstrate active control of combustion instabilities in a direct-injection gas turbine combustor that accurately simulates engine operating conditions and reproduces an engine-type instability. This report documents the second phase of a two-phase effort. The first phase involved the analysis of an instability observed in a developmental aeroengine and the design of a single-nozzle test rig to replicate that phenomenon. This was successfully completed in 2001 and is documented in the Phase I report. This second phase was directed toward demonstration of active control strategies to mitigate this instability and thereby demonstrate the viability of active control for aircraft engine combustors. This involved development of high-speed actuator technology, testing and analysis of how the actuation system was integrated with the combustion system, control algorithm development, and demonstration testing in the single-nozzle test rig. A 30 percent reduction in the amplitude of the high-frequency (570 Hz) instability was achieved using actuation systems and control algorithms developed within this effort. Even larger reductions were shown with a low-frequency (270 Hz) instability. This represents a unique achievement in the development and practical demonstration of active combustion control systems for gas turbine applications.

Description: The acoustic liner system designed for use in the High Speed Civil Transport (HSCT) was tested in a thermal-acoustic environment. Five ceramic matrix composite (CMC) acoustic tile configurations, five bulk acoustic absorbers, and one thermal protection system design were tested. The CMC acoustic tiles were subjected to two 2 3/4 hr ambient temperature acoustic exposures to measure their dynamic response. One exposure was conducted on the tiles alone and the second exposure included the tiles and the T-foam bulk absorber. The measured tile RMS strains were small. With or without the T-foam absorber, the dynamic strains were below strain levels that would cause damage during fatigue loading. After the ambient exposure, a 75-hr durability test of the entire acoustic liner system was conducted using a thermal-acoustic cycle that approximated the anticipated service cycle. Acoustic loads up to 139 dB/Hz and temperatures up to 1670 F (910 C) were employed during this 60 cycle test. During the durability test, the CMC tiles were exposed to temperatures up to 1780 F and a transient through thickness gradient up to 490 F. The TPS peak temperatures on the hot side of the panels ranged from 750 to 1000 F during the 60 cycles. The through thickness delta T ranged from 450 to 650 F, varying with TPS location and cycle number. No damage, such as cracks or chipping, was observed in the CMC tiles after completion of the testing. However, on tile warped during the durability test and was replaced after 43 or 60 cycles. No externally observed damage was found in this tile. No failure of the CMC fasteners occurred, but damage was observed. Cracks and missing material occurred, only in the fastener head region. No indication of damage was observed in the T-foam acoustic absorbers. The SiC foam acoustic absorber experienced damage after about 43 cycles. Cracking in the TPS occurred around the attachment holes and under a vent. In spite of the development of damage, the TPS maintained its insulative capability throughout the durability test. The durability test results demonstrate damage-tolerant CMC tile, CMC fastener, TPS, and T-foam absorber designs for the combined thermal and acoustic engine nozzle environment.

Description: This study was motivated by a goal to understand the mixing and emissions in the Rich-burn/Quick-mix/Lean-burn (RQL) combustor scheme that has been proposed to minimize the formation of oxides of nitrogen (NOx) in gas turbine combustors. The study reported herein was a reacting jet-in-crossflow experiment at atmospheric pressure. The jets were injected from the perimeter of a cylindrical duct through round-hole orifices into a fuel-rich mainstream flow. The number of orifices investigated in this study gave over- to optimum to underpenetrating jets at a jet-to-mainstream momentum-flux ratio of J = 57. The size of individual orifices was decreased as the number of orifices increased to maintain a constant total area; the jet-to-mainstream mass-flow ratio was constant at MR = 2.5. The experiments focused on the effects of the number of orifices and inlet air preheat and were conducted in a facility that provided the capability for independent variation of jet and main inlet air preheat temperature. The number of orifices was found to have a significant effect on mixing and the distributions of species, but very little effect on overall NOx emissions, suggesting that an aerodynamically optimum mixer might not minimize NOx emissions. Air preheat was found to have very little effect on mixing and the distributions of major species, but preheating both main and jet air did increase NOx emissions significantly. Although the air jets injected in the quick-mix section of an RQL combustor may comprise over 70 percent of the total air flow, the overall NOx emission levels were found to be more sensitive to main stream air preheat than to jet stream air preheat.

Description: The NASA John H. Glenn Research Center has a wealth of experience in Halbach array technology through the Fundamental Aeronautics Program. The goals of the program include improving aircraft efficiency, reliability, and safety. The concept of a Halbach magnetically levitated electric aircraft motor will help reduce harmful emissions, reduce the Nation s dependence on fossil fuels, increase efficiency and reliability, reduce maintenance and decrease operating noise levels. Experimental hardware systems were developed in the GRC Engineering Development Division to validate the basic principles described herein and the theoretical work that was performed. A number of Halbach Magnetic rotors have been developed and tested under this program. A separate test hardware setup was developed to characterize each of the rotors. A second hardware setup was developed to test the levitation characteristics of the rotors. Each system focused around a unique Halbach array rotor. Each rotor required original design and fabrication techniques. A 4 in. diameter rotor was developed to test the radial levitation effects for use as a magnetic bearing. To show scalability from the 4 in. rotor, a 1 in. rotor was developed to also test radial levitation effects. The next rotor to be developed was 20 in. in diameter again to show scalability from the 4 in. rotor. An axial rotor was developed to determine the force that could be generated to position the rotor axially while it is rotating. With both radial and axial magnetic bearings, the rotor would be completely suspended magnetically. The purpose of this report is to document the development of a series of Halbach magnetic rotors to be used in testing. The design, fabrication and assembly of the rotors will be discussed as well as the hardware developed to test the rotors.

Description: A key technological concept for producing reliable engine diagnostics and prognostics exploits the benefits of fusing sensor data, information, and/or processing algorithms. This report describes the development of a hybrid engine model for a propulsion gas turbine engine, which is the result of fusing two diverse modeling methodologies: a physics-based model approach and an empirical model approach. The report describes the process and methods involved in deriving and implementing a hybrid model configuration for a commercial turbofan engine. Among the intended uses for such a model is to enable real-time, on-board tracking of engine module performance changes and engine parameter synthesis for fault detection and accommodation.

Description: The purpose of this engine feasibility study was to determine the benefits that can be achieved by incorporating positive displacement axial vane compression and expansion stages into high bypass turbofan engines. These positive-displacement stages would replace some or all of the conventional compressor and turbine stages in the turbine engine, but not the fan. The study considered combustion occurring internal to an axial vane component (i.e., Diesel engine replacing the standard turbine engine combustor, burner, and turbine); and external continuous flow combustion with an axial vane compressor and an axial vane turbine replacing conventional compressor and turbine systems.

Description: This paper at first describes the fluid network approach recently implemented into the National Combustion Code (NCC) for the simulation of transport of aerosols (volatile particles and soot) in the particulate sampling systems. This network-based approach complements the other two approaches already in the NCC, namely, the lower-order temporal approach and the CFD-based approach. The accuracy and the computational costs of these three approaches are then investigated in terms of their application to the prediction of particle losses through sample transmission and distribution lines. Their predictive capabilities are assessed by comparing the computed results with the experimental data. The present work will help establish standard methodologies for measuring the size and concentration of particles in high-temperature, high-velocity jet engine exhaust. Furthermore, the present work also represents the first step of a long term effort of validating physics-based tools for the prediction of aircraft particulate emissions.

Description: Model predictive control is a strategy well-suited to handle the highly complex, nonlinear, uncertain, and constrained dynamics involved in aircraft engine control problems. However, it has thus far been infeasible to implement model predictive control in engine control applications, because of the combination of model complexity and the time allotted for the control update calculation. In this paper, a multiplexed implementation is proposed that dramatically reduces the computational burden of the quadratic programming optimization that must be solved online as part of the model-predictive-control algorithm. Actuator updates are calculated sequentially and cyclically in a multiplexed implementation, as opposed to the simultaneous optimization taking place in conventional model predictive control. Theoretical aspects are discussed based on a nominal model, and actual computational savings are demonstrated using a realistic commercial engine model.

Description: In several recent studies and on-going developments for advanced rotorcraft, the need for variable or multi-speed capable rotors has been raised. A speed change of up to 50 percent has been proposed for future rotorcraft to improve overall vehicle performance. Accomplishing rotor speed changes during operation requires both a rotor that can perform effectively over the operation speed/load range, and a propulsion system that can enable these speed changes. A study has been completed to investigate possible drive system arrangements that can accommodate up to the 50 percent speed change. Several concepts will be presented and evaluated. The most promising configurations will be identified and developed for future testing in a sub-scaled test facility to validate operational capability.

Description: Current collaborative research with Pratt & Whitney on Ultra High Bypass Engine Cycle noise, performance and emissions improvements as part of the Subsonic Fixed Wing Project Ultra High Bypass Engine Partnership Element is discussed. The Subsonic Fixed Wing Project goals are reviewed, as well as their relative technology level compared to previous NASA noise program goals. Progress toward achieving the Subsonic Fixed Wing Project goals over the 2008 fiscal year by the UHB Partnership in this area of research are reviewed. The current research activity in Ultra High Bypass Engine Cycle technology, specifically the Pratt & Whitney Geared Turbofan, at NASA and Pratt & Whitney are discussed including the contributions each entity bring toward the research project, and technical plans and objectives. Pratt & Whitney Geared Turbofan current and future technology and business plans are also discussed, including the role the NASA SFW UHB partnership plays toward achieving those goals.

Description: Current collaborative research with General Electric Aviation on Open Rotor propulsion as part of the Subsonic Fixed Wing Project Ultra High Bypass Engine Partnership Element is discussed. The Subsonic Fixed Wing Project goals are reviewed, as well as their relative technology level compared to previous NASA noise program goals. The current Open Rotor propulsion research activity at NASA and GE are discussed including the contributions each entity bring toward the research project, and technical plans and objectives. GE Open Rotor propulsion technology and business plans currently and toward the future are also discussed, including the role the NASA SFW UHB partnership plays toward achieving those goals.

Description: During an aerospace engineer's undergraduate studies, he or she will attend classes in aerodynamics, thermodynamics, structures, stability and control, dynamics, design, propulsion, and computer science, along with the related courses in mathematics, physics, statistics, and chemistry required to understand the material. Upon graduation, the new engineer will have acquired a basic knowledge of how to build an aerospace vehicle. What only comes through experience, however, is the understanding of the inevitable imperfect process through which an aerospace vehicle is built. This is the adventure of turning a basic concept into functional hardware. Engineers working on a project must often deal with ambiguous situations. They are routinely asked by management to provide risk assessments of a project, yet even after careful analysis uncertainties remain. The project must be accomplished within finite limits of time and money. The question an engineer answers is whether the solution to potential problem is worth the cost and schedule delay, or if the solution might actually be worse than the problem it is meant to solve. Review protocols are established to ensure that an unknown has not been overlooked. But these cannot protect against an unknown unknown. Examples of these situations can be found in the history of the X-43A Hyper-X (Hypersonic Experiment) program. In this NASA project, a supersonic combustion ramjet (scramjet) engine was flight tested on a subscale vehicle. The X-43A Hyper-X Research Vehicle (HXRV) was launched from a B-52B mothership, then boosted to the test speed by a modified Pegasus rocket first stage, called the Hyper-X Launch Vehicle (HXLV). Once at the proper speed and altitude, the X-43A separated from the booster, stabilized itself, and then the engine test began. Although wind-tunnel scramjet engine tests had begun in the late 1950s, before the Hyper-X program there had never been an actual in-flight test of such an engine integrated with an appropriate airframe. Thus, while the scramjet had successfully operated in the artificial airflow of wind tunnels, the concept had yet to be proven in real air. These conditions meant changes in density and temperature, as well as changes in angle of attack and sideslip of a free-flying vehicle. A wind tunnel is limited in its ability to simulate these subtle factures, which have a major impact on almost any vehicle, but especially that of a scramjet's performance. The Hyper-X project was to provide a real-world benchmark of the ground test data. The full scale X-43A engine would be operated in the wind tunnel, and then flown, and the data from its operation would then be compared with projections. If these matched, the wind tunnel data would be considered a reliable design tool for future scramjet. If there were significant differences, the reasons for these would have to be identified. Until such information was available, scramjets would lack the technological maturity to be considered for future space launch or high-speed atmospheric flight vehicles.

Description: A wide range of rocket propulsion test work occurs at the NASA John C. Stennis Space Center (SSC) including full-scale engine test activities at test facilities A-1, A-2, B-1 and B-2 as well as combustion device research and development activities at the E-Complex (E-1, E-2, E-3 and E-4) test facilities. The propulsion test engineer at NASA SSC faces many challenges associated with designing and operating a test facility due to the extreme operating conditions (e.g., cryogenic temperatures, high pressures) of the various system components and the uniqueness of many of the components and systems. The purpose of this paper is to briefly describe the NASA SSC Engineering Science Directorate s design and analysis processes, experience, and modeling techniques that are used to design and support the operation of unique rocket propulsion test facilities.

Description: This document is intended as an introduction to the analysis of gas turbine engine cycles using the Numerical Propulsion System Simulation (NPSS) code. It is assumed that the analyst has a firm understanding of fluid flow, gas dynamics, thermodynamics, and turbomachinery theory. The purpose of this paper is to provide for the novice the information necessary to begin cycle analysis using NPSS. This paper and the annotated example serve as a starting point and by no means cover the entire range of information and experience necessary for engine performance simulation. NPSS syntax is presented but for a more detailed explanation of the code the user is referred to the NPSS User Guide and Reference document (ref. 1).

Description: A new linear point design technique is presented for the determination of tuning parameters that enable the optimal estimation of unmeasured engine outputs, such as thrust. The engine s performance is affected by its level of degradation, generally described in terms of unmeasurable health parameters related to each major engine component. Accurate thrust reconstruction depends on knowledge of these health parameters, but there are usually too few sensors to be able to estimate their values. In this new technique, a set of tuning parameters is determined that accounts for degradation by representing the overall effect of the larger set of health parameters as closely as possible in a least-squares sense. The technique takes advantage of the properties of the singular value decomposition of a matrix to generate a tuning parameter vector of low enough dimension that it can be estimated by a Kalman filter. A concise design procedure to generate a tuning vector that specifically takes into account the variables of interest is presented. An example demonstrates the tuning parameters ability to facilitate matching of both measured and unmeasured engine outputs, as well as state variables. Additional properties of the formulation are shown to lend themselves well to diagnostics.

Description: A collection of experimental data acquired in the NASA low-speed multistage axial compressor while operated in rotating stall is presented in this paper. The compressor was instrumented with high-response wall pressure modules and a static pressure disc probe for in-flow measurement, and a split-fiber probe for simultaneous measurements of velocity magnitude and flow direction. The data acquired to-date have indicated that a single fully developed stall cell rotates about the flow annulus at 50.6% of the rotor speed. The stall phenomenon is substantially periodic at a fixed frequency of 8.29 Hz. It was determined that the rotating stall cell extends throughout the entire compressor, primarily in the axial direction. Spanwise distributions of the instantaneous absolute flow angle, axial and tangential velocity components, and static pressure acquired behind the first rotor are presented in the form of contour plots to visualize different patterns in the outer (midspan to casing) and inner (hub to mid-span) flow annuli during rotating stall. In most of the cases observed, the rotating stall started with a single cell. On occasion, rotating stall started with two emerging stall cells. The root cause of the variable stall cell count is unknown, but is not attributed to operating procedures.

Description: Gas turbine engines are designed to provide sufficient safety margins to guarantee robust operation with an exceptionally long life. However, engine performance requirements may be drastically altered during abnormal flight conditions or emergency maneuvers. In some situations, the conservative design of the engine control system may not be in the best interest of overall aircraft safety; it may be advantageous to "sacrifice" the engine to "save" the aircraft. Motivated by this opportunity, the NASA Aviation Safety Program is conducting resilient propulsion research aimed at developing adaptive engine control methodologies to operate the engine beyond the normal domain for emergency operations to maximize the possibility of safely landing the damaged aircraft. Previous research studies and field incident reports show that the propulsion system can be an effective tool to help control and eventually land a damaged aircraft. Building upon the flight-proven Propulsion Controlled Aircraft (PCA) experience, this area of research will focus on how engine control systems can improve aircraft safe-landing probabilities under adverse conditions. This paper describes the proposed research topics in Engine System Requirements, Engine Modeling and Simulation, Engine Enhancement Research, Operational Risk Analysis and Modeling, and Integrated Flight and Propulsion Controller Designs that support the overall goal.

Description: This paper provides an overview of current vibration methods used to identify the health of helicopter transmission gears. The gears are critical to the transmission system that provides propulsion, lift and maneuvering of the helicopter. This paper reviews techniques used to process vibration data to calculate conditions indicators (CI's), guidelines used by the government aviation authorities in developing and certifying the Health and Usage Monitoring System (HUMS), condition and health indicators used in commercial HUMS, and different methods used to set thresholds to detect damage. Initial assessment of a method to set thresholds for vibration based condition indicators applied to flight and test rig data by evaluating differences in distributions between comparable transmissions are also discussed. Gear condition indicator FM4 values are compared on an OH58 helicopter during 14 maneuvers and an OH58 transmission test stand during crack propagation tests. Preliminary results show the distributions between healthy helicopter and rig data are comparable and distributions between healthy and damaged gears show significant differences.

Description: An optimized rotorcraft propulsion system incorporating a foil air bearing supported Oil-Free engine coupled to a high power density gearbox using high viscosity gear oil is explored. Foil air bearings have adequate load capacity and temperature capability for the highspeed gas generator shaft of a rotorcraft engine. Managing the axial loads of the power turbine shaft (low speed spool) will likely require thrust load support from the gearbox through a suitable coupling or other design. Employing specially formulated, high viscosity gear oil for the transmission can yield significant improvements (approx. 2X) in allowable gear loading. Though a completely new propulsion system design is needed to implement such a system, improved performance is possible.

Description: This paper describes recent research into the advancement of small, electric powered unmanned aerial vehicle (UAV) capabilities. Specifically, topics include the improvements made in battery technology, design methodologies, avionics architectures and algorithms, materials and structural concepts, propulsion system performance prediction, and others. The results of prototype vehicle designs and flight tests are discussed in the context of their usefulness in defining and validating progress in the various technology areas. Further areas of research need are also identified. These include the need for more robust operating regimes (wind, gust, etc.), and continued improvement in payload fraction vs. endurance.

Description: This paper discusses the progress of work to model high-speed supersonic reacting flow. The purpose of the work is to improve the state of the art of CFD capabilities for predicting the flow in high-speed propulsion systems, particularly combustor flowpaths. The program has several components including the development of advanced algorithms and models for simulating engine flowpaths as well as a fundamental experimental and diagnostic development effort to support the formulation and validation of the mathematical models. The paper will provide details of current work on experiments that will provide data for the modeling efforts along with the associated nonintrusive diagnostics used to collect the data from the experimental flowfield. Simulation of a recent experiment to partially validate the accuracy of a combustion code is also described.

Description: The propulsion system of rotorcraft vehicles is the most critical system to the vehicle in terms of safety and performance. The propulsion system must provide both vertical lift and forward flight propulsion during the entire mission. Whereas propulsion is a critical element for all flight vehicles, it is particularly critical for rotorcraft due to their limited safe, un-powered landing capability. This unparalleled reliability requirement has led rotorcraft power plants down a certain evolutionary path in which the system looks and performs quite similarly to those of the 1960 s. By and large the advancements in rotorcraft propulsion have come in terms of safety and reliability and not in terms of performance. The concept of the optimized propulsion system is a means by which both reliability and performance can be improved for rotorcraft vehicles. The optimized rotorcraft propulsion system which couples an oil-free turboshaft engine to a highly loaded gearbox that provides axial load support for the power turbine can be designed with current laboratory proven technology. Such a system can provide up to 60% weight reduction of the propulsion system of rotorcraft vehicles. Several technical challenges are apparent at the conceptual design level and should be addressed with current research.

Description: This presentation provides a brief overview of the research underway in the Cruise Efficiency -- Propulsion technical challenge area of NASA's Fundamental Aeronautics Supersonics project. The research involves both computational and experimental efforts in the areas of Advanced Inlet Concepts, High Performance/Wide Operability Fan and Compressors, Advanced Nozzle Concepts and Intelligent Sensors/Actuators. The work consists of both internal NASA research and external efforts funded through the NASA Research Announcement process.

Description: At supersonic cruise conditions, high fuel temperatures, coupled with low pressures in the combustor, create potential for superheated fuel injection leading to shorter fuel jet break-up time and reduced spray penetration. Another issue particularly important to the supersonic cruise is the aircraft emissions contributing to the climate change in the atmosphere. Needless to say, aircraft emissions in general also contribute to the air pollution in the neighborhood of airports. The objectives of the present efforts are to establish baseline for prediction methods and experimental data for (a) liquid fuel atomization and vaporization at superheated conditions and (b) particle sampling systems and laboratory or engine testing environments, as well as to document current capabilities and identify gaps for future research.

Description: An overview of the emissions related research being conducted as part of the Fundamental Aeronautics Subsonics Fixed Wing Project is presented. The overview includes project metrics, milestones, and descriptions of major research areas. The overview also includes information on some of the emissions research being conducted under NASA Research Announcements. Objective: Development of comprehensive detailed and reduced kinetic mechanisms of jet fuels for chemically-reacting flow modeling. Scientific Challenges: 1) Developing experimental facilities capable of handling higher hydrocarbons and providing benchmark combustion data. 2) Determining and understanding ignition and combustion characteristics, such as laminar flame speeds, extinction stretch rates, and autoignition delays, of jet fuels and hydrocarbons relevant to jet surrogates. 3) Developing comprehensive kinetic models for jet fuels.

Description: An overview of emissions related research being conducted as part of the Fundamental Aeronautics Supersonics Project is presented. The overview includes project objectives, milestones, and descriptions of major research areas. The overview also includes information on the emissions research being conducted under NASA Research Announcements. Technical challenges include: 1) Environmental impact of supersonic cruise emissions is greater due to higher flight altitudes which makes emissions reduction increasingly important. 2) Accurate prediction tools to enable combustor designs that reduce emissions at supersonic cruise are needed as well as intelligent systems to minimize emissions. 3) Combustor operating conditions at supersonic cruise are different than at subsonic cruise since inlet fuel and air temperatures are considerably increased.

Description: New powder metallurgy (PM) disk superalloys, such as ME3, LSHR, and Alloy 10, have been developed in recent years which enable rim temperatures in turbine disk applications to approach 1300 F. Before these alloys can be utilized at 1300 F their long term durability must be ensured. One of the key requirements for disk rims is notch fatigue strength. This issue is extremely important and is a direct result of the blade attachment geometry employed at the disk rim. Further, the imposition of a dwell at maximum load, associated with take off and landing, can also affect notch fatigue strength. For these reasons a study has been undertaken to assess the notch dwell fatigue strength of a modern PM disk alloy through spin pit evaluation of a prototypical disk. The first element of this program involves screening potential heat treatments with respect to notch fatigue strength at 1300 F utilizing a conventional notch fatigue specimen with a stress concentration factor (K(sub t)) of 2 and a 90 sec dwell at peak load. The results of this effort are reported in this paper including the downselect of an optimal heat treatment, from a notch fatigue standpoint.

Description: Gas turbine engines for aero-propulsion systems are found to be highly optimized machines after over 70 years of development. Still, additional performance improvements are sought while reduction in the overall cost is increasingly a driving factor. Control systems play a vitally important part in these metrics but are severely constrained by the operating environment and the consequences of system failure. The considerable challenges facing future engine control system design have been investigated. A preliminary analysis has been conducted of the potential benefits of distributed control architecture when applied to aero-engines. In particular, reductions in size, weight, and cost of the control system are possible. NASA is conducting research to further explore these benefits, with emphasis on the particular benefits enabled by high temperature electronics and an open-systems approach to standardized communications interfaces.

Description: High temperature seals are required for advanced hypersonic airframe applications. In this study, both spring tube thermal barriers and innovative wafer seal systems were evaluated under relevant hypersonic test conditions (temperatures, pressures, etc.) via high temperature compression testing and room temperature flow assessments. Thermal barriers composed of a Rene 41 spring tube filled with Saffil insulation and overbraided with a Nextel 312 sheath showed acceptable performance at 1500 F in both short term and longer term compression testing. Nextel 440 thermal barriers with Rene 41 spring tubes and Saffil insulation demonstrated good compression performance up to 1750 F. A silicon nitride wafer seal/compression spring system displayed excellent load performance at temperatures as high as 2200 F and exhibited room temperature leakage values that were only 1/3 those for the spring tube rope seals. For all seal candidates evaluated, no significant degradation in leakage resistance was noted after high temperature compression testing. In addition to these tests, a superalloy seal suitable for dynamic seal applications was optimized through finite element techniques.

Description: 14. ABSTRACT In a gas turbine engine, the turbine rotor blades are buffeted by the wakes of the vanes located upstream. There is a transient effect from the passing of wakes on the blade heat transfer. This transient effect has been computed for a representative rotor by introducing a wake upstream via an unsteady inlet flow boundary condition, or "gust" condition. Two cases of turbulent flow and laminar flow with Reynolds numbers of 385,000 and 385 respectively were considered. For the turbulent flow case a quasi-steady calculation was also performed. The variation in the unsteady heat transfer coefficient was found to be as high as 120 percent of the mean. For the turbulent flow case a quasisteady calculation was also performed. The time mean of the unsteady heat transfer, the mean of the quasi-steady variations and the steady results agree reasonably well on all blade locations except for the turbulent results which differ near the leading edge. The quasi-steady heat transfer results do not agree with the instantaneous unsteady results, although the time-mean values are similar.

Description: This paper presents a preliminary demonstration of an automated health assessment tool, capable of real-time on-board operation using existing engine control hardware. The tool allows operators to discern how rapidly individual turboshaft engines are degrading. As the compressor erodes, performance is lost, and with it the ability to generate power. Thus, such a tool would provide an instant assessment of the engine s fitness to perform a mission, and would help to pinpoint any abnormal wear or performance anomalies before they became serious, thereby decreasing uncertainty and enabling improved maintenance scheduling. The research described in the paper utilized test stand data from a T700-GE-401 turboshaft engine that underwent sand-ingestion testing to scale a model-based compressor efficiency degradation estimation algorithm. This algorithm was then applied to real-time Health Usage and Monitoring System (HUMS) data from a T700-GE-701C to track compressor efficiency on-line. The approach uses an optimal estimator called a Kalman filter. The filter is designed to estimate the compressor efficiency using only data from the engine s sensors as input.

Description: This document addresses the modeling task plan for the hypersonic GN&C GRC team members. The overall propulsion system modeling task plan is a multi-step process and the task plan identified in this document addresses the first steps (short term modeling goals). The procedures and tools produced from this effort will be useful for creating simplified dynamic models applicable to a hypersonic vehicle propulsion system. The document continues with the GRC short term modeling goal. Next, a general description of the desired simplified model is presented along with simulations that are available to varying degrees. The simulations may be available in electronic form (FORTRAN, CFD, MatLab,...) or in paper form in published documents. Finally, roadmaps outlining possible avenues towards realizing simplified model are presented.

Description: An outer loop retrofit engine control architecture is presented which modifies fan speed command to obtain a desired thrust based on throttle position. This maintains the throttle-to-thrust relationship in the presence of engine degradation, which has the effect of changing the engine s thrust output for a given fan speed. Such an approach can minimize thrust asymmetry in multi-engine aircraft, and reduce pilot workload. The outer loop control is demonstrated under various levels of engine deterioration using a standard deterioration profile as well as an atypical profile. It is evaluated across various transients covering a wide operating range. The modified fan speed command still utilizes the standard engine control logic so all original life and operability limits remain in place. In all cases it is shown that with the outer loop thrust control in place, the deteriorated engine is able to match the thrust performance of a new engine up to the limits the controller will allow.

Description: This paper investigates the integration of on-line and off-line diagnostic algorithms for aircraft gas turbine engines. The on-line diagnostic algorithm is designed for in-flight fault detection. It continuously monitors engine outputs for anomalous signatures induced by faults. The off-line diagnostic algorithm is designed to track engine health degradation over the lifetime of an engine. It estimates engine health degradation periodically over the course of the engine s life. The estimate generated by the off-line algorithm is used to update the on-line algorithm. Through this integration, the on-line algorithm becomes aware of engine health degradation, and its effectiveness to detect faults can be maintained while the engine continues to degrade. The benefit of this integration is investigated in a simulation environment using a nonlinear engine model.

Description: An advanced model turbofan was tested in the NASA Glenn 9- by 15-Foot Low Speed Wind Tunnel (9x15 LSWT) to explore far field acoustic effects associated with rotor Trailing-Edge-Blowing (TEB) for a modern, 1.294 stage pressure ratio turbofan model. The TEB rotor (Fan9) was designed to be aerodynamically similar to the previously tested Fan1, and used the same stator and nacelle hardware. Fan9 was designed with trailing edge blowing slots using an external air supply directed through the rotor hub. The TEB flow was heated to approximate the average fan exit temperature at each fan test speed. Rotor root blockage inserts were used to block TEB to all but the outer 40 and 20% span in addition to full-span blowing. A configuration with full-span TEB on alternate rotor blades was also tested. Far field acoustic data were taken at takeoff/approach conditions at 0.10 tunnel Mach. Far-field acoustic results showed that full-span blowing near 2.0% of the total flow could reduce the overall sound power level by about 2 dB. This noise reduction was observed in both the rotor-stator interaction tones and for the spectral broadband noise levels. Blowing only the outer span region was not very effective for lowering noise, and actually increased the far field noise level in some instances. Full-span blowing of alternate blades at 1.0% of the overall flow rate (equivalent to full-span blowing of all blades at 2.0% flow) showed a more modest noise decrease relative to full-span blowing of all blades. Detailed hot film measurements of the TEB rotor wake at 2.0% flow showed that TEB was not every effective for filling in the wake defect at approach fan speed toward the tip region, but did result in overfilling the wake toward the hub. Downstream turbulence measurements supported this finding, and support the observed reduction in spectral broadband noise.

Description: With the increased emphasis on aircraft safety, enhanced performance and affordability, and the need to reduce the environmental impact of aircraft, there are many new challenges being faced by the designers of aircraft propulsion systems. The Controls and Dynamics Branch at NASA (National Aeronautics and Space Administration) Glenn Research Center (GRC) in Cleveland, Ohio, is leading and participating in various projects in partnership with other organizations within GRC and across NASA, the U.S. aerospace industry, and academia to develop advanced controls and health management technologies that will help meet these challenges through the concept of Intelligent Propulsion Systems. The key enabling technologies for an Intelligent Propulsion System are the increased efficiencies of components through active control, advanced diagnostics and prognostics integrated with intelligent engine control to enhance operational reliability and component life, and distributed control with smart sensors and actuators in an adaptive fault tolerant architecture. This presentation describes the current activities of the Controls and Dynamics Branch in the areas of active component control and propulsion system intelligent control, and presents some recent analytical and experimental results in these areas.

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

Description: NASA s Intelligent Propulsion System Technology (Propulsion 21) project focuses on developing adaptive technologies that will enable commercial gas turbine engines to produce fewer emissions and less noise while increasing reliability. It features adaptive technologies that have included active tip-clearance control for turbine and compressor, active combustion control, turbine aero-thermal and flow control, and enabling technologies such as sensors which are reliable at high operating temperatures and are minimally intrusive. A probabilistic system analysis is performed to evaluate the impact of these technologies on aircraft CO2 (directly proportional to fuel burn) and LTO (landing and takeoff) NO(x) reductions. A 300-passenger aircraft, with two 396-kN thrust (85,000-pound) engines is chosen for the study. The results show that NASA s Intelligent Propulsion System technologies have the potential to significantly reduce the CO2 and NO(x) emissions. The results are used to support informed decisionmaking on the development of the intelligent propulsion system technology portfolio for CO2 and NO(x) reductions.

Description: The engineering design and analysis of air-breathing propulsion systems relies heavily on zero- or one-dimensional properties (e:g: thrust, total pressure recovery, mixing and combustion efficiency, etc.) for figures of merit. The extraction of these parameters from experimental data sets and/or multi-dimensional computational data sets is therefore an important aspect of the design process. A variety of methods exist for extracting performance measures from multi-dimensional data sets. Some of the information contained in the multi-dimensional flow is inevitably lost when any one-dimensionalization technique is applied. Hence, the unique assumptions associated with a given approach may result in one-dimensional properties that are significantly different than those extracted using alternative approaches. The purpose of this effort is to examine some of the more popular methods used for the extraction of performance measures from multi-dimensional data sets, reveal the strengths and weaknesses of each approach, and highlight various numerical issues that result when mapping data from a multi-dimensional space to a space of one dimension.

Description: A method for modeling engine operation comprising the steps of: 1. collecting a first plurality of sensory data, 2. partitioning a flight envelope into a plurality of sub-regions, 3. assigning the first plurality of sensory data into the plurality of sub-regions, 4. generating an empirical model of at least one of the plurality of sub-regions, 5. generating a statistical summary model for at least one of the plurality of sub-regions, 6. collecting an additional plurality of sensory data, 7. partitioning the second plurality of sensory data into the plurality of sub-regions, 8. generating a plurality of pseudo-data using the empirical model, and 9. concatenating the plurality of pseudo-data and the additional plurality of sensory data to generate an updated empirical model and an updated statistical summary model for at least one of the plurality of sub-regions.

Description: A turbine stage includes a row of airfoils joined to corresponding platforms to define flow passages therebetween. Each airfoil includes opposite pressure and suction sides and extends in chord between opposite leading and trailing edges. Each platform includes a crescentic ramp increasing in height from the leading and trailing edges toward the midchord of the airfoil along the pressure side thereof.

Description: This report is a Users Guide for the NASA-developed Commercial Modular Aero-Propulsion System Simulation (C-MAPSS) software, which is a transient simulation of a large commercial turbofan engine (up to 90,000-lb thrust) with a realistic engine control system. The software supports easy access to health, control, and engine parameters through a graphical user interface (GUI). C-MAPSS provides the user with a graphical turbofan engine simulation environment in which advanced algorithms can be implemented and tested. C-MAPSS can run user-specified transient simulations, and it can generate state-space linear models of the nonlinear engine model at an operating point. The code has a number of GUI screens that allow point-and-click operation, and have editable fields for user-specified input. The software includes an atmospheric model which allows simulation of engine operation at altitudes from sea level to 40,000 ft, Mach numbers from 0 to 0.90, and ambient temperatures from -60 to 103 F. The package also includes a power-management system that allows the engine to be operated over a wide range of thrust levels throughout the full range of flight conditions.

Description: This study evaluated the feasibility of a hybrid solid oxide fuel cell (SOFC) auxiliary power unit (APU) and the impact in a 90-passenger More-Electric Regional Jet application. The study established realistic hybrid SOFC APU system weight and system efficiencies, and evaluated the impact on the aircraft total weight, fuel burn, and emissions from the main engine and the APU during cruise, landing and take-off (LTO) cycle, and at the gate. Although the SOFC APU may be heavier than the current conventional APU, its weight disadvantage can be offset by fuel savings in the higher SOFC APU system efficiencies against the main engine bleed and extraction during cruise. The higher SOFC APU system efficiency compared to the conventional APU on the ground can also provide considerable fuel saving and emissions reduction, particularly at the gate, but is limited by the fuel cell stack thermal fatigue characteristic.

Description: One of the key technology challenges for the use of hydrogen in gas turbine engines is the performance of the combustion system, in particular the fuel injectors. To investigate the combustion performance of gaseous hydrogen fuel injectors flame tube combustor experiments were performed. Tests were conducted to measure the nitrogen oxide (NO(x)) emissions and combustion performance at inlet conditions of 588 to 811 K, 0.4 to 1.4 MPa, and equivalence ratios up to 0.48. All the injectors were based on Lean Direct Injection (LDI) technology with multiple injection points and quick mixing. One challenge to hydrogen-based premixing combustion systems is flashback since hydrogen has a reaction rate over 7 times that of Jet-A. To reduce the risk, design mixing times were kept short and velocities high to minimize flashback. Five fuel injector designs were tested in 6.35- and 8.9-cm-diameter flame tubes with non-vitiated heated air and gaseous hydrogen. Data is presented on measurements of NO(x) emissions and combustion efficiency for the hydrogen injectors at 2.540, 7.937, and 13.652 cm from the injector face. Results show that for some configurations, NO(x) emissions are comparable to that of state of the art Jet-A LDI combustor concepts.

Description: Recent developments in gas foil bearing technology have led to numerous advanced high-speed rotating system concepts, many of which have become either commercial products or experimental test articles. Examples include oil-free microturbines, motors, generators and turbochargers. The driving forces for integrating gas foil bearings into these high-speed systems are the benefits promised by removing the oil lubrication system. Elimination of the oil system leads to reduced emissions, increased reliability, and decreased maintenance costs. Another benefit is reduced power plant weight. For rotorcraft applications, this would be a major advantage, as every pound removed from the propulsion system results in a payload benefit.. Implementing foil gas bearings throughout a rotorcraft gas turbine engine is an important long-term goal that requires overcoming numerous technological hurdles. Adequate thrust bearing load capacity and potentially large gearbox applied radial loads are among them. However, by replacing the turbine end, or hot section, rolling element bearing with a gas foil bearing many of the above benefits can be realized. To this end, engine manufacturers are beginning to explore the possibilities of hot section gas foil bearings in propulsion engines. This overview presents a logical follow-on activity by analyzing a conceptual rotorcraft engine to determine the feasibility of a foil bearing supported core. Using a combination of rotordynamic analyses and a load capacity model, it is shown to be reasonable to consider a gas foil bearing core section. In addition, system level foil bearing testing capabilities at NASA Glenn Research Center are presented along with analysis work being conducted under NRA Cooperative Agreements.

Description: This document provides a model specification for the rework and/or repair of bearings used in aircraft engines, helicopter main power train transmissions, and auxiliary bearings determined to be critical by virtue of performance, function, or availability. The rolling-element bearings to be processed under the provisions of this model specification may be used bearings removed after service, unused bearings returned from the field, or certain rejected bearings returned for reinspection and salvage. In commercial and military aircraft application, it has been a practice that rolling-element bearings removed at maintenance or overhaul be reworked and returned to service. Depending on the extent of rework and based upon theoretical analysis, representative life factors (LF) for bearings subject to rework ranged from 0.87 to 0.99 the lives of new bearings. Based on bearing endurance data, 92 percent of the bearing sets that would be subject to rework would result in L(sub 10) lives equaling and/or exceeding that predicted for new bearings. The remaining 8 percent of the bearings have the potential to achieve the analytically predicted life of new bearings when one of the rings is replaced at rework. The potential savings from bearing rework varies from 53 to 82 percent of that of new bearings depending on the cost, size, and complexity of the bearing

Description: This study was motivated by a goal to understand the mixing and emissions in the rich-burn/quick-mix/lean-burn (RQL) combustor scheme that has been proposed to minimize the formation of oxides of nitrogen (NOx) in gas turbine combustors. The study reported in this paper was a reacting jet-in-crossflow experiment at atmospheric pressure in a cylindrical duct. The jets were injected from the perimeter of the duct through round-hole orifices into a fuel-rich mainstream flow. The number of orifices investigated in this study gave over- to optimum to underpenetrating jets at a jet-to-mainstream momentum-flux ratio of 57. The size of individual orifices was decreased as their number increased to maintain a constant total area. The jet-to-mainstream mass-flow ratio was held constant at 2.5. The experiments focused on the effects of the number of orifices and inlet air preheat and were conducted in a facility that provided the capability for independent variation of jet and main inlet air preheat temperature. The number of orifices was found to have a significant effect on mixing and the distributions of species, but very little effect on overall NOx emissions, suggesting that an aerodynamically optimum mixer may not minimize NOx emissions. Air preheat was found to have very little effect on mixing and the distributions of major species, but preheat did increase NOx emissions significantly. Although the air jets injected in the quick-mix section of a RQL combustor may comprise over 70% of the total air flow, the overall NOx emission levels were found to be more sensitive to mainstream air preheat than to jet stream air preheat.

Description: NASA s Fundamental Aeronautics Program is investigating turbine-based propulsion systems for access to space because it provides the potential for aircraft-like, space-launch operations that may significantly reduce launch costs and improve safety. Studies performed under NASA s NGLT and the NASP High Speed Propulsion Assessment (HiSPA) program indicated a variable cycle turbofan/ramjet was the best configuration to satisfy access-to-space mission requirements because this configuration maximizes the engine thrust-to-weight ratio while minimizing frontal area. To this end, NASA and GE teamed to design a Mach 4 variable cycle turbofan/ramjet engine for access to space. To enable the wide operating range of a Mach 4+ variable cycle turbofan ramjet required the development of a unique fan stage design capable of multi-point operation to accommodate variations in bypass ratio (10X), fan speed (7X), inlet mass flow (3.5X), inlet pressure (8X), and inlet temperature (3X). The primary goal of the fan stage was to provide a high pressure ratio level with good efficiency at takeoff through the mid range of engine operation, while avoiding stall and losses at the higher flight Mach numbers, without the use of variable inlet guide vanes. Overall fan performance and operability therefore requires major consideration, as competing goals at different operating points and aeromechanical issues become major drivers in the design. To mitigate risk of meeting the unique design requirements for the fan stage, NASA and GE teamed to design and build a 57% engine scaled fan stage to be tested in NASA s transonic compressor facility. The objectives of this test are to assess the aerodynamic and aero mechanic performance and operability characteristics of the fan stage over the entire range of engine operation including: 1) sea level static take-off, 2) transition over large swings in fan bypass ratio, 3) transition from turbofan to ramjet, and 4) fan windmilling operation at high Mach flight conditions. In addition, the fan stage design was validated by performing pre-test CFD analysis using both GE proprietary and NASA s APNASA codes. Herein we will discuss 1) the fan stage design, 2) the experiment including the unique facility and instrumentation, and 3) the comparison of pre-test CFD analysis to initial aerodynamic test results for the baseline fan stage configuration. Measurements and pre-test analysis will be compared at 37%, 50%, 80%, 90%, and 100% of design speed to assess the ability of state-of-the-art design and analysis tools to meet the fan stage performance and operability requirements for turbine based propulsion for access to space.