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Turboshaft engine performance and weight models were developed to support conceptual propulsion and vehicle mission design and performance under the Revolutionary Vertical Lift Technology (RVLT) Project in 2019 by Snyder in Reference 7. These models were developed using open data sources, assuming present and future technology levels, and range from 650 to 5,000 output shaft horsepower (485 to 3,730 kW). This paper expands on the previous research, extending the power ranges from 200 to 15000 output shaft horsepower (150 to 11,200 kW) and documenting the methodology, assumptions, and engine performance realizes important benefits for NASA and the aviation community. NASA concept-vehicle study efforts using these baseline propulsion models can be more readily shared among the government, industry, and university community to support present and future work. Assessing the benefits of advanced technologies and new configurations can be facilitated using these models, which helps guide technology investment. As the various conceptual vehicle and mission analysis simulations are developed, these models can be used directly for broader systems analysis studies, including optimization within the propulsion model itself. To expand on the previous effort, the turboshaft engine is briefly discussed, highlighting the specific components, and expected performance characteristics over the updated power range and technology levels considered. Additional engine configurations will also be discussed as they vary based on power output and assumed technology level. Engine performance, such as airflow, power output and weight are updated, noting important trends for system studies. Finally, the effect of advanced propulsion technologies on public reference models including RVLT concept vehicles are reported along with the tools and software methods used to complete the analysis.

Overview of RVLT Electric Propulsion Testbeds

This paper presents a parametric modeling and integrated aircraft sizing and synthesis approach for a charge depleting parallel hybrid electric architecture. The developed models are integrated within the baseline thin-haul and regional aircraft. In addition to the physical architecture, different modes of operation enabled by propulsion system electrification are also modeled parametrically. The modes of operation presented in this paper are the peak power shaving, climb power electric boost, in-flight battery recharging, and electric taxi. The sizing of the powertrain and the aircraft are performed within the multidisciplinary analysis and optimization environment, E-PASS. The consideration of the physical system and its operation together provides a holistic approach where the propulsion system and the airframe are designed under an optimized power and energy management strategy. The parametric nature of the work enables the design space exploration for electrification and lays the groundwork for future technology projection and uncertainty quantification studies. The developed capability is generic and can be applied on other aircraft classes. The work is done as part of the Electrified Powertrain Flight Demonstration program.

This paper presents a detailed market analysis of the U.S. domestic aviation market in support of the NASA subsonic single aft engine (SUSAN) regional aircraft concept. The current scoping of the SUSAN concept is intended to compete in the medium to large (160-180 seats) narrow body market, with range capabilities of up to 2,500 nautical miles, and expected fuel burn reduction up to 40% relative to conventional 2 engine aircraft. Recent historical trends suggest growth in aviation passenger demand will continue to be met by the narrow body fleet of aircraft; however, a comprehensive review is required to understand how these current fleet trends could evolve in the future. Moreover, estimating and forecasting the potential market size is the critical first step when developing a new aircraft concept to determine commercial viability. To assist in the trade space exploration of the SUSAN concept, a generalized traffic and fleet forecast of the U.S. aviation market is conducted. Using publicly available aviation data from the U.S. Bureau of Transportation Statistics and passenger demand forecasts from FAA, a multinomial logit model is estimated to predict the composition of the future fleet by aircraft size. These fleet forecasts are then used as inputs for a fleet evolution model to provide required operational forecasts at the aircraft specific level. Forecast scenarios with and without the SUSAN concept are compared, and a breakeven analysis is performed to evaluate the commercial viability of the SUSAN aircraft from an operating cost perspective. Results from the multinomial logit fleet forecast indicate the narrow body size category of 150+ seats dominating the market, comprising 87% of the future revenue passenger miles market share in 2050 (up from 60% in 2019). Forecast scenarios with the SUSAN concept see a maximum cumulative decline in fuel cost of 20% by 2050, while a breakeven analysis shows competitive advantage of the SUSAN aircraft, due to expected fuel burn reduction, at moderate levels of increased maintenance and capital costs.

Gaseous emissions were measured in single-cup flametube tests of an advanced low-NOx combustor concept at simulated supersonic cruise conditions. The combustor concept is a low technology readiness level (TRL), lean front-end design developed under the NASA Fundamental Aeronautics/Supersonics project to minimize NOx emissions at supersonic cruise. The flametube conditions matched or approached combustor conditions at supersonic cruise, with combustor inlet temperatures up to 920 K, inlet pressures up to 19 bar, and combusted gas temperatures up to 2,120 K. Whether these conditions met or just approached supersonic cruise conditions depended on the type of engine the combustor would be installed in. Two types of engines were considered here: a "derivative" engine based on a current technology and an "advanced" engine with a higher operating pressure ratio and higher temperature limits. For the "derivative" engine, the combustor is expected to be at least close to meeting the NASA NOx emissions goal of 10 g-NOx/kg-fuel at supersonic cruise. However, with the higher combustor inlet and flame temperatures of the advanced engine, NOx emissions are expected to be well above the goal.

This paper describes a process to develop a flight simulation test capability for the SUbsonic Single Aft eNgine (SUSAN) Electrofan, a subsonic regional jet transport aircraft concept that utilizes electrified propulsion to gain benefits in fuel usage, emissions, and cost. The process, which involves the integration of independently developed models and their subsequent implementation in a flight simulator, is general and can be applied to a variety of aircraft types. However, the use of electrified propulsion architectures has the potential to add complexity beyond that of a traditional aircraft, especially with regard to the pilot interface. The way the pilot interacts with the thrust producing components could vary significantly between architectures, and the information displayed to the pilot will necessarily include additional variables beyond what is normally displayed in a traditional cockpit. This paper describes the integration process in general, as well as specific accommodations made for the architecture under consideration.

Distributed Electric Propulsion technology is expected to yield up to a fivefold increase in high-speed cruise efficiency for NASA’s X-57 “Maxwell” flight demonstrator when compared to a combustion-powered general aviation baseline. A portion of this increased efficiency is due to beneficial aero-propulsive interaction inherent to the distributed propulsion architecture. The measure of the relative increase in efficiency between a conventional and distributed propulsion wing will be extracted from comparisons between flight test data from the electrically powered X-57 Mod II configuration with a conventional wing, and from the electrically powered X-57 Mod III/IV configuration with a distributed propulsion wing. Flight test maneuvers that accommodate errors in instrumentation and the flight test environment are developed to establish the power-off drag characteristics for all X-57 configurations. Analysis of these maneuvers with typical errors, including pilot-in-the-loop simulation data that incorporates simulated atmospheric turbulence effects, shows that the proposed power-off flight maneuvers can generate accurate power-off drag predictions for the X-57. These predictions show that the power-off differences in aerodynamic performance between the conventional and distributed propulsion configurations can be accurately measured from flight test data in the presence of typical data error sources.

A growing demand for higher-power electrical equipment on aircraft is driving the exploration of alternative electrical power generation devices. “Hybrid-electric” aircraft architectures that leverage multiple or alternative power sources are also becoming more popular with rising concerns over efficiency. Fuel cells are particularly interesting as auxiliary or secondary electrical power generation devices, due to their relative efficiency and longevity, and could be used as a “power pod” for vehicles without enough on-board electrical generation capability. This paper examines the feasibility and potential value of using an externally-mounted, hybridelectric, solid-oxide fuel cell (SOFC) power system on a Class IV unmanned aerial vehicle (UAV) for the purpose of power generation in excess of the baseline electrical load. Moreover, the paper will demonstrate that a SOFC auxiliary power unit (APU) provides significant value to an aircraft by producing a relatively large amount of additional electrical power for a relatively low initial investment when compared to producing additional power by scaling the engine.

The adjoint-based design capability in FUN3D is extended to allow efficient gradient-based optimization and design of concepts with highly integrated and distributed aero-propulsive systems. Calculations of propulsive power, along with the derivatives needed to perform adjoint-based design, have been implemented in FUN3D. The design capability is demonstrated by the shape optimization and propulsor sizing of NASA’s PEGASUS aircraft concept. The optimization objective is the minimization of flow power at the aerodynamic interface planes for the wing-mounted propulsors, as well as the tail-cone boundary layer ingestion propulsor, subject to vehicle performance and propulsive constraints.

This paper presents preliminary ice-crystal icing (ICI) altitude scaling evaluation results of a Honeywell Uncertified Research Engine (HURE) that was tested in the NASA Glenn Research Center Propulsion Systems Laboratory (PSL) during January of 2018. This engine geometry features a hidden core design to keep the core less exposed. The engine was fitted with internal video cameras to observe various ice buildup processes at multiple selected locations within the engine core flow path covering the fan stator, the splitter-lip/shroud/strut, and the high pressure compressor (HPC) variable inlet guide vane (IGV) regions. The potential ice accretion risk was pre-determined to occur by using NASA’s in-house 1D Engine Icing Risk assessment code, COMDES-MELT. The code was successful in predicting the risk of ice accretion in adiabatic regions like the fan-stator of the HURE at specific engine operating points. However at several operating points during the test, liquid water was observed running along the shroud toward the variable IGV of the HPC regions with an air temperature well below freezing, thus no particle melting could have occurred due to heating from the air alone. It was reasoned that other sources of heat were present in that region. To account for these heat sources the inlet total temperature was adjusted to give a wet bulb temperature of 24 F below the standard minimum wet bulb temperature of 492 R to allow ice to accrete in the splitter-lip/shroud/strut region, which was determined from a reference case where hard ice was observed in that region. With that adjustment the COMDES-MELT code was successful in providing operating points where there was a risk of ice accretion during the test campaign. In addition to calculating possible conditions at different selected lower altitudes, simulations were run to determine potential inlet conditions that could lead to ice-crystal accretion along the prescribed stations where the cameras were available. From there, scaled test conditions were determined by best matching the following three icing related parameters of the reference condition: (1) the local air total wet bulb temperature, (2) the local ice crystal cloud melt ratio and (3) the engine fan face ice/water to air mass flux ratio of the ice crystal cloud. Instantaneous images taken from the time-lapsed movies of ice buildup were used along with the relevant thermodynamic data of air, water vapor and local icing condition to help evaluate how closely the proposed altitude scaling method could be used in ground based test facility to duplicate selected reference ICI features observed at specific location inside this engine at different scale altitudes. Discussions on observed limitation for engine icing scaling application from this test campaign and needed improvement are provided. A scaling test procedure to help identify potential ICI risk conditions and possible ice accretion locations of a new turbofan engine is evaluated in PSL.

The X-57 Maxwell is NASA’s all-electric demonstration vehicle. The primary demonstration objective of this flight test program is to show a factor of five reduction in energy consumption. The vehicle includes two large wing tip propellers designed to provide propul- sion at cruise conditions and twelve leading edge propellers designed to operate at high lift conditions. The first configuration of the vehicle that will be flight tested has the large wing tip propellers relocated to an inboard wing station. A simplified structural dynamic model of the propulsion system has been generated and coupled with a beam model of the vehicle. Whirl flutter analyses have been performed, examining the stability of the isolated propulsion system and coupled to the beam model of the vehicle. Trimmed flight scenarios for the vehicle include straight and level flight and zero power windmilling conditions. The whirl flutter analyses for this configuration indicate that the configuration will be free of whirl flutter within the required flight envelope.

The Tiltrotor Test Rig (TTR) is a NASA project, joint with the U.S. Army and Air Force, to develop a new, large scale proprotor test system for the National Full-Scale Aerodynamics Complex (NFAC). The first wind-tunnel entry was completed in November 2018 with a modern, 26-ft diameter proprotor. The primary purpose was to complete the development of the TTR, including systems integration with the NFAC. The TTR and rotor were tested up to 273 knots in axial flow. This is the highest airspeed ever achieved by a full-scale proprotor in any wind tunnel. Extensive conversion-mode data were also acquired, and hover/climb conditions were explored. Additional testing included aerodynamic tares, motor tests, thermal tests, modal vibration tests, and other checkout activities. This paper summarizes the results of the test, including examples of the most significant data.

A tendency for excessive exhaust jet mixing noise from low bypass ratio turbofan engines is recognized as a key challenge in the design of commercial supersonic aircraft. In this work we investigate a unique combination of two noise mitigation methods as a novel strategy to reduce jet mixing noise. First, a thermal acoustic shield (TAS) is used to reflect high frequency acoustic waves at small angles to the jet axis; second, a mixer-ejector (ME) nozzle is used to mechanically shield noise propagating at large angles to the axis. The ME shroud also provides a convenient location for a TAS nozzle and improves TAS effectiveness by limiting the downstream extent of high frequency noise generation. In an additional benefit for a velocity-matched TAS stream, the ME allows a reduction in strength of the TAS outer shear layer which could serve as a secondary noise source. The present work provides a quantitative assessment of the ME-TAS concept, using a combination of RANS CFD simulations, acoustic analogy calculations for the farfield Green's function, and surrogate-based modeling and parameter space exploration. We first evaluate a subscale configuration, then use scaling arguments to apply subscale results to the systems-level analysis of a flight configuration; the latter configuration includes a generic low bypass ratio turbofan engine with an engine-driven electric generator for supplementary heating of the TAS stream. Additional RANS CFD calculations are performed for a notional ME-TAS geometry based on the full scale configuration, and various modeling assumptions and operational characteristics are evaluated. The ME-TAS concept is shown to provide effective shielding for high frequency jet noise, and should enable comparable noise suppression to a stand-alone ME of considerably greater length, weight and drag. In addition to investigating the integrated ME-TAS system, the present work differs from previous research into TAS and related fluidic shield concepts through the inclusion of modern numerical analysis tools and the systematic numerical examination of various design parameters.

This paper presents an autonomous path-following control architecture for a tilt-wing, distributed electric propulsion, vertical take-off and landing unmanned aerial system in hover mode and presents indoor flight test results. The test-bed vehicle is a subscale model with the same configuration as the NASA GL-10 aircraft. The control architecture consists of an inner-loop attitude controller, outer-loop trajec­tory controller, and a trajectory generation scheme. The flight test results show that the vehicle can satisfactorily follow a path prescribed by a list of waypoints around the indoor flight room.

Reynolds Averaged Navier-Stokes (RANS) simulations were performed on an N+2 conceptual commercial supersonic transport aircraft to explore the possibility of relocating the engines from below the wings to above the wings in order to capitalize upon potential noise shielding benefits. The simulations focused on the feasibility of the top-mounted propulsion configuration in terms of inlet performance and flow separation around the nacelles at take-off conditions for both 0deg and 8deg angle of attack. The results showed overall good inlet performance and little separation around the nacelles. The results were comparable to the engines in their original underwing configuration.

Reynolds Averaged Navier-Stokes (RANS) simulations were performed on the Lockheed Martin Quiet SuperSonic Technology (QueSST) aircraft preliminary design to assess inlet performance. The FUN3D flow solver and its adjoint-based grid refinement capability was used for the simulations in hopes of determining internal "best practices" for predicting inlet performance on top-aft-mounted inlets. Several parameters were explored including tetrahedral vs. pentahedral cells in/around the boundary-layer regions, an engine axis-aligned linear pressure sensor vs. a pressure box objective as the grid adaptation metric, and the number of grid adaptation cycles performed. Additional simulations were performed on manually refined grids for comparison with the adjoint-based adapted grids

The design and analysis of on-demand mobility class vehicles will require thorough acoustic analysis to ensure that more numerous aircraft can operate in densely populated areas without causing excessive levels of noise. This work is a step towards a comprehensive ODM vehicle analysis capability. The authors use 6DOF equations of motion to model the electric quad-rotor concept developed by NASA's Revolutionary Vertical Lift Technologies program. As a first step towards acoustic analysis, this trajectory is coupled to an acoustic model that tracks the sound pressure level perceived by an acoustic observer on the ground. The results show that the approach is successful in finding trajectories that minimize total propulsive impulse while obeying limits imposed on the sound pressure level. Future work will involve adding acoustic analysis of increasing fidelity and tying the resulting trajectories to the performance of the electric propulsion system.

This paper presents results in which we compare fuel staging and its effect on fuel spray pattern, velocity and speciation during combustion for several inlet conditions using a GE TAPS injector configuration. Planar laser-induced fluorescence (PLIF), particle image velocimetry (PIV) and phase Doppler interferometry (PDI) were used to investigate spray patterns and velocity. The 2D PIV provides slices in the flow of axial-vertical or axial horizontal velocity components. With 3D PDI, we obtained 3 components of velocity, and fuel drop sizes. Chemiluminescence imaging and spontaneous Raman scattering (SRS) were used to investigate flame structure, species location and relative species concentration. Phase Doppler and PIV data were acquired using scatter from fuel droplets; therefore, those data were obtained only at the pilot-only test points. Raman measurements were acquired only at 10/90 split points to avoid droplets.

Lightweight, efficient power production has been a pacing technology for aviation. Advances in airborne electric propulsion technology have enabled new aviation concepts in markets that previously were not dominated by aviation systems, largely because electric propulsion allows for efficient, integrated propulsion/aerodynamic/control solutions that were previously not practical with combustion-based power architectures. These new markets include automated package delivery, urban air mobility, and short-haul transportation. As these markets evolve, the impact of using electricity for airborne propulsion on an expanding mission set, as well as the sheer amount of energy consumed, will begin to challenge the energy harvesting and distribution paradigm as it exists today. Left unaddressed, these challenges could stymie the evolution of these new markets and mobility options. This paper identifies some of the potential systemic issues associated with the expanded use of electric propulsion and explores the requirements associated with alternate aviation power architectures. The recommended path includes the development of a new hybrid-electric aviation power architecture that can be used in conjunction with a portfolio of evolving battery-electric and combustion-based systems.

As electric powertrain and propulsion-airframe integration technologies advance, airborne electric propulsion concepts appear to be on the cusp of disrupting or transforming aviation markets. One of the many challenges to this transformation lies within the onboard energy storage and generation technologies. State-of-the-art battery technology is heavy and lacks support infrastructure; purely combustion-based solutions to electrical power generation suffer from increased inefficiency as compared to a traditional combustion powertrain. This paper explores another alternative: a hybrid-electric, solid oxide fuel cell power system. This power system reforms traditional fuels to feed fuel cells that generate electrical power for the aircraft electric powertrain. A representative power system is designed based on the requirements for NASA’s X-57 Mod II electric flight demonstrator platform, and is shown to exceed system feasibility goals of 300 W/kg and 60% efficiency in at least one configuration. The hybrid-electric fuel cell power system is shown to be competitive with the range performance of a combustion-based power architecture and comparable in mission energy cost to a battery-electric power system.

Contributions from the combustor to the overall propulsion noise of civilian transport aircraft are starting to become important due to turbofan design trends and advances in mitigation of other noise sources. Future propulsion systems for ultra-efficient commercial air vehicles are projected to be of increasingly higher bypass ratio from larger fans combined with much smaller cores, with ultra-clean burning fuel-flexible combustors. Unless effective noise-reduction strategies are developed, combustor noise is likely to become a prominent contributor to overall airport community noise in the future. The new NASA DGEN Aeropropulsion Research Turbofan (DART) is a cost-efficient testbed for the study of core-noise physics and mitigation. This paper describes the recently completed DART core/combustor-noise baseline test in the NASA GRC Aero-Acoustic Propulsion Laboratory (AAPL). Acoustic data were simultaneously acquired using the AAPL overhead microphone array in the engine aft quadrant farfield, a single midfield microphone, and two semi-infinite-tube unsteady pressure sensors at the core-nozzle exit. Combustor-noise components of measured total-noise signatures were educed using a two-signal source-separation method and are found to occur in the expected frequency range. The acoustic data compare well with results from a limited 2014 feasibility test and will serve as a high-quality baseline for future research using the DART. The research described herein is aligned with the NASA Ultra-Efficient Commercial Transport strategic thrust and is supported by the NASA Advanced Air Vehicle Program, Advanced Air Transport Technology Project, under the Aircraft Noise Reduction Subproject.

Reynolds Averaged Navier-Stokes (RANS) simulations were performed on the Lockheed Martin Quiet SuperSonic Technology (QueSST) aircraft preliminary design to assess inlet performance. The FUN3D flow solver and its adjoint-based grid refinement capability was used for the simulations in hopes of determining internal "best practices" for predicting inlet performance on top-aft-mounted inlets. Several parameters were explored including tetrahedral vs. pentahedral cells in/around the boundary-layer regions, an engine axis-aligned linear pressure sensor vs. a pressure box objective as the grid adaptation metric, and the number of grid adaptation cycles performed. Additional simulations were performed on manually refined grids for comparison with the adjoint-based adapted grids

This paper investigates the effects of cavity diameter on the acoustic performance of conventional perforate-over-honeycomb liners. This investigation is a follow up to a previous study where it was demonstrated that the cavity diameter has a relatively small effect on the acoustic impedance for core cavity diameters up to 2.0 inches. In the current study, liner concepts are exposed to a complex acoustic environment, where higher-order modes are present. Four liner configurations are explored in this investigation, which include two cavity cores of different cavity diameters combined with two facesheets. These configurations are tested in the Grazing Flow Impedance Tube (GFIT) at Mach 0.0, 0.3, and 0.5 at a source sound pressure level of 150 dB. A comparison of the acoustic performance of these samples suggests that the cavity diameter has a relatively small effect on the acoustic impedance for core cavity diameters up to 2.0 inches at these conditions. These configurations are exposed to a more complex acoustic environment (e.g., higher-order modes) in the Curved Duct Test Rig (CDTR) at Mach 0.0 and 0.3.

This paper presents a conceptual design of a tail-cone thruster system which is operating under an axisymmetric inlet distortion. An effort to realize the targeted fuel burn saving that was proposed in NASA's STARC_ABL aircraft design is made through a CFD (Computational Fluid Dynamics)-based design approach. This method employs three iterative steps to exploit the CFD tools until the design requirements are met: a quasi-2D through-flow model to design the fan/EGV (Exit Guide Vane), a 3-D RANS (Reynolds Averaging Navier-Stokes) simulation of the single blade row to account for the inlet/fan and the EGV/nozzle interaction, and a 3-D RANS simulation of the airframe with a propulsor installed - propulsion airframe integration (PAI). The design requirements which include the thrust, and shaft power of the propulsor are matched throughout the evaluations coming from two CFD domains, i.e., the turbo-machinery and the PAI. During the switch between these different computational domains, the inlet and exit profiles are matched via the correction factors of the body-force model. The present tail-cone thruster (TCT) aerodynamic design leverages a low-pressure ratio fan (FPR=1.2 to approximately 1.25) of which the camber-line angles are predicted by a quasi-2D through-flow model. The quasi-2D model is derived to analyze the radially distorted flow resulting from the ingested boundary layer at the inlet. It also estimates the appropriate velocity vectors of the metal angles of the fan and EGV which is subjected to different types of vortex at the fan exit. The baseline geometry is revisited and its internal flow-path and exhaust cone are redesigned to illustrate the strong correlation among the components of the propulsor in the PAI domain. The peak efficiency point of the fan/EGV with respect to the blade counts, also known as solidity, and rotational speed is chosen for the cruise condition via parametric studies. The corresponding performance maps are presented. The resulting performance metrics of the new conceptual design of the BLI (Boundary Layer Ingestion) propulsor are analyzed and compared with these of the baseline in the PAI aspect. Finally, ideas of the CFD based design of a BLI propulsor are discussed based on the observations drawn from the numerical results.

This presentation is an overview of the current Air Mobility programs being studied by NASA.

An aircraft has wings configured to twist during flight. Inboard and outboard propulsion devices, such as turbofans or other propulsors, are connected to each wing, and are spaced along the wing span. A flight controller independently controls thrust of the inboard and outboard propulsion devices to significantly change flight dynamics, including changing thrust of outboard propulsion devices to twist the wing, and to differentially apply thrust on each wing to change yaw and other aspects of the aircraft during various stages of a flight mission. One or more generators can be positioned upon the wing to provide power for propulsion devices on the same wing, and on an opposite wing.

Studies of various injector configurations in a 7-point Lean Direct Injector (LDI) array are reported for both non-reacting (cold) flow and for Jet-A/air reacting flows. For cold flow, central recirculation zone (CRZ) formation is investigated and for reacting flows, combustor operability and dynamics are of interest. 2D Particle Image Velocimetry (PIV) measurements are described for the cold flow experiments and flame chemiluminescence imaging and dynamic pressure results are discussed for the reacting flow cases. PIV results indicate that for this configuration the close spacing between swirler elements leads to strong interaction that affects whether a CRZ forms, and pilot recess and counter-swirl helps to isolate swirlers from one another. Dynamics results focus on features identified near 500-Hz.

An engine system has a gas generator, a bi-fi wall surrounding at least a portion of the gas generator, a casing surrounding a fan, and the casing having first and second thrust reverser doors which in a deployed position abut each other and the bi-fi wall.

This paper describes the design process for the control system of an advanced geared turbofan engine. This process is applied to a simulation that is representative of a 30,000 pound-force thrust class concept engine with two main spools, ultra-high bypass ratio, and a variable area fan nozzle. Control system requirements constrain the non-linear engine model as it operates throughout its flight envelope of sea level to 40,000 feet and from 0 to 0.8 Mach. The purpose of this paper is to review the engine control design process for an advanced turbofan engine configuration. The control architecture selected for this project was developed from literature and reflects a configuration that utilizes a proportional integral controller with sets of limiters that enable the engine to operate safely throughout its flight envelope. Simulation results show the overall system meets performance requirements without exceeding operational limits.

Assessing the potential to bring 100 years of aeronautics knowledge to the entrepreneurs desktop to enable a design environment for emerging vertical lift vehicles is one goal for the NASAs Design Environment for Novel Vertical Lift Vehicles (DELIVER). As part of this effort, a system study was performed using a notional, urban aerial taxi system to better understand vehicle requirements along with the tools and methods capability to assess these vehicles and their subsystems using cryogenic cooled components. The baseline was a vertical take-off and landing (VTOL) aircraft, with all-electric propulsion system assuming 15 year technology performance levels and its capability limited to a pilot with one or two people and cargo. Hydrocarbon-fueled hybrid concepts were developed to improve mission capabilities. The hybrid systems resulted in significant improvements in maximum range and number of on demand mobility (ODM) missions that could be completed before refuel or recharge. An important consideration was thermal management, including the choice for air-cooled or cryogenic cooling using liquid natural gas (LNG) fuel. Cryogenic cooling for critical components can have important implications on component performance and size. Thermal loads were also estimated, subsequent effort will be required to verify feasibility for cooling airflow and packaging. LNG cryogenic cooling of selected components further improved vehicle range and reduced thermal loads, but the same concerns for airflow and packaging still need to be addressed. The use of the NASA Design and Analysis of Rotorcraft (NDARC) tool for vehicle sizing and mission analysis appears to be capable of supporting analyses for present and future types of vehicles, missions, propulsion, and energy sources. Further efforts are required to develop verified models for these new types of propulsion and energy sources in the size and use envisioned for these emerging vehicle and mission classes.

Assessing the potential to bring 100 years of aeronautics knowledge to the entrepreneurs desktop to enable a design environment for emerging vertical lift vehicles is one goal for the NASA's Design Environment for Novel Vertical Lift Vehicles (DELIVER). As part of this effort, a system study was performed using a notional, urban aerial taxi system to better understand vehicle requirements along with the tools and methods capability to assess these vehicles and their subsystems using cryogenic cooled components. The baseline was a vertical take-off and landing (VTOL) aircraft, with all-electric propulsion system assuming 15 year technology performance levels and its capability limited to a pilot with one or two people and cargo. Hydrocarbon-fueled hybrid concepts were developed to improve mission capabilities. The hybrid systems resulted in significant improvements in maximum range and number of on demand mobility (ODM) missions that could be completed before refuel or recharge. An important consideration was thermal management, including the choice for air-cooled or cryogenic cooling using liquid natural gas (LNG) fuel. Cryogenic cooling for critical components can have important implications on component performance and size. Thermal loads were also estimated, subsequent effort will be required to verify feasibility for cooling airflow and packaging. LNG cryogenic cooling of selected components further improved vehicle range and reduced thermal loads, but the same concerns for airflow and packaging still need to be addressed. The use of the NASA Design and Analysis of Rotorcraft (NDARC) tool for vehicle sizing and mission analysis appears to be capable of supporting analyses for present and future types of vehicles, missions, propulsion, and energy sources. Further efforts are required to develop verified models for these new types of propulsion and energy sources in the size and use envisioned for these emerging vehicle and mission classes.

A computational study of the wing for the distributed electric propulsion X-57 Maxwell airplane configuration at cruise and takeoff/landing conditions was completed. Two unstructured-mesh, Navier-Stokes computational fluid dynamics methods, FUN3D and USM3D, were used to predict the wing performance. The goal of the X-57 wing and distributed electric propulsion system design was to meet or exceed the required lift coefficient 3.95 for a stall speed of 58 knots, with a cruise speed of 150 knots at an altitude of 8,000 ft. The X-57 Maxwell airplane was designed with a small, high aspect ratio cruise wing that was designed for a high cruise lift coefficient (0.75) at angle of attack of 0deg. The cruise propulsors at the wingtip rotate counter to the wingtip vortex and reduce induced drag by 7.5 percent at an angle of attack of 0.6deg. The unblown maximum lift coefficient of the high-lift wing (with the 30deg flap setting) is 2.439. The stall speed goal performance metric was confirmed with a blown wing computed effective lift coefficient of 4.202. The lift augmentation from the high-lift, distributed electric propulsion system is 1.7. The predicted cruise wing drag coefficient of 0.02191 is 0.00076 above the drag allotted for the wing in the original estimate. However, the predicted drag overage for the wing would only use 10.1 percent of the original estimated drag margin, which is 0.00749.

Electric aircraft pose a unique design challenge in that they lack a simple way to reject waste heat from the power train. While conventional aircraft reject most of their excess heat in the exhaust stream, for electric aircraft this is not an option. To examine the implications of this challenge on electric aircraft design and performance, we developed a model of the electric subsystems for the NASA X-57 electric testbed aircraft. We then coupled this model with a model of simple 2D aircraft dynamics and used a Legendre-Gauss-Lobatto collocation optimal control approach to find optimal trajectories for the aircraft with and without thermal constraints. The results show that the X-57 heat rejection systems are well designed for maximum-range and maximum-efficiency flight, without the need to deviate from an optimal trajectory. Stressing the thermal constraints by reducing the cooling capacity or requiring faster flight has a minimal impact on performance, as the trajectory optimization technique is able to find flight paths which honor the thermal constraints with relatively minor deviations from the nominal optimal trajectory.

As efficiency, emissions, and noise become increasingly prominent considerations in aircraft design, turning to an electric propulsion system is a desirable solution. Achieving the intended benefits of distributed electric propulsion (DEP) requires thermally demanding high power systems, presenting a different set of challenges compared to traditional aircraft propulsion. The embedded nature of these heat sources often preclude the use of traditional thermal management systems in order to maximize performance, with less opportunity to exhaust waste heat to the surrounding environment. This paper summarizes the thermal analyses of X-57 vehicle subsystems that don't employ externally air-cooled heat sinks. The high-power battery, wires, high-lift motors, and aircraft outer surface are subjected to heat loads with stringent thermal constraints. The temperature of these components are tracked transiently, since they never reach a steady-state equilibrium. Through analysis and testing, this report demonstrates that properly characterizing the material properties is key to accurately modeling peak temperature of these systems, with less concern for spatial thermal gradients. Experimentally validated results show the thermal profile of these systems can be sufficiently estimated using reduced order approximations.

Aerodynamic shape optimization is demonstrated to tailor the under-track pressure signature of a conceptual low-boom supersonic aircraft. Primarily, the optimization reduces nearfield pressure waveforms induced by propulsion integration effects. For computational efficiency, gradient-based optimization is used and coupled to the discrete adjoint formulation of the Reynolds-averaged Navier Stokes equations. The engine outer nacelle, nozzle, and vertical tail fairing are axi-symmetrically parameterized, while the horizontal tail is shaped using a wing-based parameterization. Overall, 48 design variables are coupled to the geometry and used to deform the outer mold line. During the design process, an inequality drag constraint is enforced to avoid major compromise in aerodynamic performance. Linear elastic mesh morphing is used to deform volume grids between design iterations. The optimization is performed at Mach 1.6 cruise, assuming standard day altitude conditions at 51,707-ft. To reduce uncertainty, a coupled thermodynamic engine cycle model is employed that captures installed inlet performance effects on engine operation.

A gas turbine engine has a fairing and an air intake that includes an air inlet embedded within the fairing for supplying free stream atmospheric air to a gas generator.

This paper describes the power and command system architecture of the X-57 Maxwell flight demonstrator aircraft. The X-57 is an experimental aircraft designed to demonstrate radically improved aircraft efficiency with a 3.5 times aero-propulsive efficiency gain at a "high-speed cruise" flight condition for comparable general aviation aircraft. These gains are enabled by integrating the design of a new, optimized wing and a new electric propulsion system. As a result, the X-57 vehicle takes advantage of the new capabilities afforded by electric motors as primary propulsors. Integrating new technologies into critical systems in experimental aircraft poses unique challenges that require careful design considerations across the entire vehicle system, such as qualification of new propulsors (motors, in the case of the X-57 aircraft), compatibility of existing systems with a new electric power distribution bus, and instrumentation and monitoring of newly qualified propulsion system devices.

An N+3 technology level engine, suitable as a propulsion system for an advanced single-aisle transport, was developed as a reference cycle for use in technology assessment and decision-making efforts. This reference engine serves three main purposes: it provides thermodynamic quantities at each major engine station, it provides overall propulsion system performance data for vehicle designers to use in their analyses, and it can be used for comparison against other proposed N+3 technology-level propulsion systems on an equal basis. This reference cycle is meant to represent the expected capability of gas turbine engines in the N+3 timeframe given reasonable extrapolations of technology improvements and the ability to take full advantage of those improvements.

There a few NASA funded electric and hybrid electric projects from different NASA Centers, including the NASA Armstrong Flight Research Center (AFRC) (Edwards, California). Each project identifies a specific technology gap that is currently inhibiting the growth and proliferation of relevant technologies in commercial aviation. This paper describes the design and development of a turbo-electric distributed propulsion (TeDP) hardware-in-the-loop (HIL) simulation bench, which is a test bed for discovering turbo-electric control, distributed electric control, power management control, and integration competencies while providing risk mitigation for future turbo-electric flying demonstrators.

Combustion dynamics data were collected at the NASA Glenn Research Center's CE-5 flame tube test facility under combustor outlet choked conditions. Two 9-point Swirl-Venturi Lean Direct Injection (SV-LDI) configurations were tested in a rectangular cuboid combustor geometry. Combustion dynamic data were measured at different engine operational conditions up to inlet air pressure and temperature of 24.13 bar and 828 K, respectively. In this study, the effects of acoustic cavity resonance, precessing vortex core (PVC), and non-uniform thermal expansion on the dynamic noise spectrum are identified by comparing the dynamic data that collected at various combustor inlet conditions along with combustor geometric calculations. The results show that the acoustic cavity resonance noises were seen in the counter-rotating pilot configuration but not in the co-rotating pilot configuration. Dynamic pressure noise band at around 0.9 kHz was only detected at the P'41 location (9.8 cm after fuel injector face) but not at the P'42 location (29 cm after the fuel injector face); the amplitude of this noise band depended on the thermal expansion ratio (T4/T3). The noise band at around 1.8 kHz was found to depend on the inlet air pressure or the air density inside the combustor. The PVC frequency was not observed in these two configurations.

Advancements in aircraft electric propulsion may enable an expanded operational envelope for electrically powered vehicles compared to their internal combustion engine counterparts. High aspect ratio wings provide additional lift and drag reduction for a proposed multi-propulsor design, however, the challenge is to reduce the weight of wing structures while maintaining adequate structural and aeroelastic margins. Design exploration using a conventional design-and-build philosophy coupled with a finite element method (FEM)-based design of experiments (DOE) strategy are presented to examine high aspect ratio wing structures that have spanwise distributed electric motors. Multiple leading-edge-mounted engine masses presented a challenge to design a wing within acceptable limits for dynamic and aeroelastic stability. Because the first four primary bending eigenmodes of the proposed wing structure are very sensitive to outboard motor placement, safety-of-flight requirements drove the need for multiple spars, rib attachments, and outboard structural reinforcements in the design. Global aeroelasticity became an increasingly important design constraint during the on-going design process, with outboard motor pod flutter ultimately becoming a primary design constraint. Designers successively generated models to examine stress, dynamics, and aeroelasticity concurrently. This research specifically addressed satisfying multi-disciplinary design criteria to generate fluid-structure interaction solution sets, and produced high aspect ratio primary structure designs for the NASA Scalable Convergent Electric Propulsion Technology and Operations Research (SCEPTOR) project in the Aeronautic Research Mission Directorate at NASA. In this paper, a dynamics-driven, quasi-inverse design methodology is presented to address aerodynamic performance goals and structural challenges encountered for the SCEPTOR demonstrator vehicle. These results are compared with a traditional computer aided design based approach.

NASA's X-57 "Maxwell" flight demonstrator incorporates distributed electric propulsion technologies in a design that will achieve a significant reduction in energy used in cruise flight. A substantial portion of these energy savings come from beneficial aerodynamic-propulsion interaction. Previous research has shown the benefits of particular instantiations of distributed propulsion, such as the use of wingtip-mounted cruise propellers and leading edge high-lift propellers. However, these benefits have not been reduced to a generalized design or analysis approach suitable for large-scale design exploration. This paper discusses the rapid, "design-order" toolchains developed to investigate the large, complex tradespace of candidate geometries for the X-57. Due to the lack of an appropriate, rigorous set of validation data, the results of these tools were compared to three different computational flow solvers for selected wing and propulsion geometries. The comparisons were conducted using a common input geometry, but otherwise different input grids and, when appropriate, different flow assumptions to bound the comparisons. The results of these studies showed that the X-57 distributed propulsion wing should be able to meet the as-designed performance in cruise flight, while also meeting or exceeding targets for high-lift generation in low-speed flight.

The present invention is a turbojet design that integrates power generation into the turbojet itself, rather than use separate generators attached to the turbojet for power generation. By integrating the power generation within the jet engine, the weight of the overall system is significantly reduced, increasing system efficiency. Also, by integrating the power generating elements of the system within the air flow of the jet engine, the present invention can use the heat generated by the power generating elements (which is simply expelled waste heat in current designs) to increase the engine performance.

Systems, methods, and devices are provided that combine an advance vehicle configuration, such as an advanced aircraft configuration, with the infusion of electric propulsion, thereby enabling a four times increase in range and endurance while maintaining a full vertical takeoff and landing ("VTOL") and hover capability for the vehicle. Embodiments may provide vehicles with both VTOL and cruise efficient capabilities without the use of ground infrastructure. An embodiment vehicle may comprise a wing configured to tilt through a range of motion, a first series of electric motors coupled to the wing and each configured to drive an associated wing propeller, a tail configured to tilt through the range of motion, a second series of electric motors coupled to the tail and each configured to drive an associated tail propeller, and an electric propulsion system connected to the first series of electric motors and the second series of electric motors.

NASA's Small Business Innovation Research (SBIR) program focuses on technological innovation by investing in development of innovative concepts and technologies to help NASA mission directorates address critical research needs for Agency programs. This report highlights 24 of the innovative SBIR 2015 Phase I projects that emphasize one of NASA Glenn Research Center's six core competencies-Air-Breathing Propulsion. The technologies cover a wide spectrum of applications such as hybrid nanocomposites for efficient aerospace structures; plasma flow control for drag reduction; physics-based aeroanalysis methods for open rotor conceptual designs; vertical lift by series hybrid power; fast pressure-sensitive paint systems for production wind tunnel testing; rugged, compact, and inexpensive airborne fiber sensor interrogators based on monolithic tunable lasers; and high sensitivity semiconductor sensor skins for multi-axis surface pressure characterization. Each featured technology describes an innovation and technical objective and highlights NASA commercial and industrial applications. This report provides an opportunity for NASA engineers, researchers, and program managers to learn how NASA SBIR technologies could help their programs and projects, and lead to collaborations and partnerships between the small SBIR companies and NASA that would benefit both.

Set of charts used in technical portion of the meeting, covering NASA's work in hybrid electric aircraft.

This paper covers the application of a model-based engine control (MBEC) methodology featuring a self tuning on-board model for an aircraft turbofan engine simulation. The nonlinear engine model is capable of modeling realistic engine performance, allowing for a verification of the advanced control methodology over a wide range of operating points and life cycle conditions. The on-board model is a piece-wise linear model derived from the nonlinear engine model and updated using an optimal tuner Kalman Filter estimation routine, which enables the on-board model to self-tune to account for engine performance variations. MBEC is used here to show how advanced control architectures can improve efficiency during the design phase of a turbofan engine by reducing conservative operability margins. The operability margins that can be reduced, such as stall margin, can expand the engine design space and offer potential for efficiency improvements. Application of MBEC architecture to a nonlinear engine simulation is shown to reduce the thrust specific fuel consumption by approximately 1% over the baseline design, while maintaining safe operation of the engine across the flight envelope.

Three variations of a low emissions aircraft gas turbine engine combustion concept were developed and tested. The concept is a second generation swirl-venturi lean direct injection (SV-LDI) concept. LDI is a lean-burn combustion concept in which the fuel is injected directly into the flame zone. All three variations were based on the baseline 9- point SV-LDI configuration reported previously. The three second generation SV-LDI variations are called the 5-recess configuration, the flat dome configuration, and the 9- recess configuration. These three configurations were tested in a NASA Glenn Research Center medium pressure flametube. All three second generation variations had better low power operability than the baseline 9-point configuration. All three configurations had low NO(sub x) emissions, with the 5-recess configuration generally having slightly lower NO(x) than the flat dome or 9-recess configurations. Due to the limitations of the flametube that prevented testing at pressures above 20 atm, correlation equations were developed for the at dome and 9-recess configurations so that the landing-takeoff NO(sub x) emissions could be estimated. The flat dome and 9-recess landing-takeoff NO(x) emissions are estimated to be 81-88% below the CAEP/6 standards, exceeding the project goal of 75% reduction.