ALBERT

Description: A fundamental exploratory experiment is conducted assessing the performance of a one-sided ejector with the eventual goal of noise reduction for jet engines. The hardware is comprised of an 8:1 rectangular nozzle together with an ejector box whose lower surface is flush with the lower lip of the nozzle. Secondary flow is allowed through a gap between the upper lip of the nozzle and a flap that constitutes the upper surface of the ejector. Wall static pressures and Pitot probe surveys are conducted to evaluate the performance of the ejector with variation of geometric parameters. It is found that addition of vortex generating tabs at the upper lip of the nozzle significantly increases secondary flow entrainment. The entrainment is further enhanced by a divergence of the ejector upper surface. Limited noise measurements are done. The baseline ejector (without tabs) often encounters flow resonance with accompanying tones. The tabs have the additional benefit of eliminating those tones in all cases. However, for the tabbed case, addition of the ejector produces insignificant further noise reduction. This is due to the fact that the flow remains unmixed on the lower half of the ejector. The focus of ongoing and future efforts is to achieve sufficient mixing of the flow so that the exhaust velocities are uniformly low, while keeping the ejector hardware short and lightweight.

Description: Charts are presented for computing the thrust, fuel consumption, and other performance values of a turbojet engine for any given set of operating conditions and component efficiencies. The effects of the pressure losses in the inlet duct and combustion chamber, the variation in the physical properties of the gas as it passes through the cycle, and the change in mass flow by the addition of fuel are included. The principle performance charts show the effects of the primary variables and correction charts provide the effects of the secondary variables.

Description: The performance of hypothetical turbojet systems, without thrust augmentation, as power plants for supersonic airplanes has been calculated. The thrust, thrust power, air-fuel ratio, 1 specific fuel consumption, cross-sectional area, and thrust coefficient are shown for free-stream Mach numbers from 1.2 to 3. For comparison, the performance of ram-jet systems over the same Mach number range has also been calculated. For Mach numbers between 1.2 and 2 the calculated thrust coefficient of the turbojet system was found to be larger than the estimated drag coefficient, and the specific fuel consumption was calculated to be considerably less than the specific fuel consumption of the ram-jet system. The turbojet system therefore appears to merit consideration as a propulsion method for free-stream Mach numbers between approximately 1.2 and 2.

Description: Convenient charts are presented for computing the thrust, fuel consumption, and other performance values of a turbojet system. These charts take into account the effects of ram pressure, compressor pressure ratio, ratio of combustion-chamber-outlet temperature to atmospheric temperature, compressor efficiency, turbine efficiency, combustion efficiency, discharge-nozzle coefficient, losses in total pressure in the inlet to the jet-propulsion unit and in the combustion chamber, and variation in specific heats with temperature. The principal performance charts show clearly the effects of the primary variables and correction charts provide the effects of the secondary variables. The performance of illustrative cases of turbojet systems is given. It is shown that maximum thrust per unit mass rate of air flow occurs at a lower compressor pressure ratio than minimum specific fuel consumption. The thrust per unit mass rate of air flow increases as the combustion-chamber discharge temperature increases. For minimum specific fuel consumption, however, an optimum combustion-chamber discharge temperature exists, which in some cases may be less than the limiting temperature imposed by the strength temperature characteristics of present materials.

Description: Outline - Introduction: X-57 CFD task overview; Motivation. Part I, Computational simulations without propulsion: Establishing CFD (Computational Fluid Dynamics) Best Practices - Grid generation - Mesh refinement study - Numerical methods - Wind tunnel validation study; Power-Off Aerodynamic Database Results. Part II, Computational simulations with propulsion: Cruise Power-On Database; High-Lift Power-On Database. Summary.

Description: A rotor blade comprises an airfoil extending radially from a root section to a tip section and axially from a leading edge to a trailing edge, the leading and trailing edges defining a curvature therebetween. The curvature determines a relative exit angle at a relative span height between the root section and the tip section, based on an incident flow velocity at the leading edge of the airfoil and a rotational velocity at the relative span height. In operation of the rotor blade, the relative exit angle determines a substantially flat exit pressure ratio profile for relative span heights from 75% to 95%, wherein the exit pressure ratio profile is constant within a tolerance of 10% of a maximum value of the exit pressure ratio profile.

Description: A simulator to artificially generate turbofan broadband signatures using the ANCF (Advanced Noise Control Fan) test article is presented. [Development of a Broadband Acoustic Emulator to Mature Propulsion Noise Reduction (CFANS-BB: Configurable Fan Artificial Noise Source- Broadband)]

Description: Reynolds-Averaged Navier-Stokes simulations have been performed on a three-stream inverted velocity profile nozzle with and without various configurations of chevrons attached. The nozzle was mounted on a planform to imitate an engine mounted above a wing, shielding ground observers from engine noise. Several chevron designs intended to aggressively mix the jet and move noise sources upstream for shielding were examined to investigate there effects on noise and thrust. Numerical results for the baseline nozzle and one chevron configuration were compared with far-field noise and particle image velocimetry data obtained in NASA Glenn Research Center's Aero-Acoustic Propulsion Laboratory. A configuration in which chevrons alternate penetration into the primary stream and tertiary fan stream was explored using the Modern Design of Experiments approach. Short, high-penetration chevrons demonstrated a significant noise reduction for a relatively small thrust penalty.

Description: This paper describes the design of a turboshaft engine for a tiltwing air taxi application. In this case, the tiltwing air taxi is intended to fly a 400 nm mission with up to fifteen passengers. Engine requirements for the concept engine are taken from aircraft system studies where thrust is produced by four propellers driven by electric motors and powered by a single gas turbine engine. The purpose of this paper is to perform a cycle design optimization that minimizes fuel consumption and weight while respecting current technology limitations to meet mission requirements. To achieve results, the engine overall pressure ratio and maximum temperature at the exit of the combustor are set as the design parameters. Several sensitivity studies are also performed to visualize optimization trends. Results of the optimization study show solutions are heavily dependent on engine cooling flow requirements and exact mission requirements. This engine is intended for use in large system optimization research.

Description: Time accurate simulation of non-equilibrium flows inside shock tube facilities presents several challenges from both physical and mathematical aspects. Furthermore, the large computational cost makes it impractical to support a real-time experimental test campaign. In this work, we explore other methods for modeling the shock tube problem with the main focus on the post-shock region and the absolute radiation emanating from it. The proposed alternative approach is several orders of magnitude less computationally expensive while still accurate enough with regards to the quantities of interest. Excellent agreement is found with the established stagnation-line approach. Comparison with time-accurate simulations shows good agreement close to the peak values and disagreement of the temperatures relaxation and radiance profiles toward equilibrium.

Description: Magnetic gearing is being investigated at NASA as a replacement to conventional mechanical gearing in aerospace applications. Some potential benefits of magnetic gears over mechanical gearing are torque transmission without mechanical contact, decreased transmission noise, and no required lubrication. However, in order to be a viable alternative for aerospace applications, magnetic gearing must be shown to provide high enough specific torque (torque per unit mass). NASA's second magnetic gearing prototype (PT-2) was able to achieve promising specific torque on par with low torque mechanical gearboxes. This work will briefly review the electromagnetic and structural design of PT-2, provide detailed information on fabrication and assembly, examine build errors, walk through rebuild efforts to improve operation, and conclude with remarks on build difficulties and opportunities for improvement in future prototypes.

Description: NASA Acoustic Stirling IRAD (Internal Research and Development) Thermal Recovery Energy Efficient System (TREES) Energy Conversion and Management in Aircraft. Presentation on energy conversion on aircraft. Thermal energy recovery changes aircraft thermal management from being a necessary burden on aircraft performance to a desirable asset. It improves the engine performance by recycling waste heat and ultimately rejecting all collected aircraft heat out through the engine nozzle.

Description: Turboshaft engine performance and weight models were developed to support conceptual propulsion and vehicle mission design in support of the National Aeronautics and Space Administration's (NASA) Aeronautics Mission Research Directorate's (ARMD) Revolutionary Vertical Lift Technology (RVLT) Project. These models were developed using open data sources, assuming current and advanced technology levels, and range from 650 to 7,500 shaft output horsepower (485 to 5,600 kilowatts). Documenting the methodology, assumptions, and resulting performance realizes important benefits for NASA and the aviation community. NASA concept vehicle efforts using these propulsion models can more readily shared among the government, industry and university community as common baselines to support current 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 modeling conceptual vehicle and mission analysis environments advance, these models can be used directly for broader systems analysis studies, including optimization within the propulsion model itself. To perform this effort, the turboshaft engine is briefly discussed, highlighting the specific components and their expected performance characteristics over the power range and technology levels considered. Engine configurations will also be discussed as they will vary based on power output and assumed technology level. Engine performance, such as airflow, power output and weight will be reported, noting trends that are important for system studies. The effect of advanced propulsion technologies on RVLT-concept vehicles are also reported. Finally, potential future propulsion modeling work will be proposed.

Description: A model-scale exhaust system was tested to validate low-noise concepts and noise prediction methods. The tests involved far-field acoustics, translating phased array, and particle image velocimetry; this report covers the far-field acoustic measurements. Data were acquired for a series of nozzles with different chevron designs, both uninstalled and installed on a representative aircraft planform. The impact of the various chevron treatments on the far-field noise was documented, along with the impact of the pylon and planform. For the baseline nozzle, installation produced a 2 EPNdB (Effective Perceived Noise in deciBels) reduction, as assumed in system studies. Chevrons were used to shift noise sources upstream to maximize the installation benefits and to reduce unshielded sources downstream. These resulted in reductions of 4-5 EPNdB...

Description: 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. These models were developed using open data sources, assuming current and advanced technology levels, and range from 650 to 7,500 shaft output horsepower (485 to 5,600 kW). Documenting the methodology, assumptions, and resulting performance realizes important benefits NASA and the aviation community. NASA concept vehicle efforts using these propulsion models can be more readily shared among the government, industry and university community as common baselines to support current 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 modeling conceptual vehicle and mission analysis environments advanced, these models can be used directly for broader systems analysis studies, including optimization within the propulsion model itself. To perform this effort, the turboshaft engine is briefly discussed, highlighting the specific components and their expected performance characteristics over the power range and technology levels considered. Engine configurations will also be discussed as they will vary based on power output and assumed technology level. Engine performance, such as airflow, power output and weight will be reported, noting trends that are important for system studies. The effect of advanced propulsion technologies on RVLT concept vehicles are also reported. Finally, potential future propulsion modeling work will be proposed.

Description: This paper presents the design, development, operation, and test capabilities of a proposed superconducting coil testbed to measure alternating current (AC) losses at the NASA Glenn Research Center. Superconducting AC losses are important in the design of electric stators and rotors, power transmission lines, transformers, fault current limiters, magnets, and superconducting energy storage (not batteries). The new liquid-hydrogen-based rig will allow superconducting testing across a wide range of test parameters, including injected current up to 400 A, frequency (0 to 400 Hz), magnetic field (0 to 0.6 T), phase angle between induced voltage and injected current (180 to 180), coil coolant temperature (18 to 28 K), and AC power loss (5 to 30 W). While the target application of interest is 20 K superconducting MgB2 (the only superconductor that can presently be made with low losses) stator coils for future electric machines, the rig can accommodate test articles (TAs) with straight wire, tape, cables, coils of any shape, any allowable combination of superconducting wire and fluid (e.g., yttrium barium copper oxide (YBCO) coils and liquid nitrogen), and AC or direct current (DC) testing. The new spin rig builds upon the existing Air Force spin rig through a more flexible mode of fluid control, a wider gap space (up to 10.2 cm) for TAs, and the ability to accommodate TAs over a wider range of operating temperatures (18 to 95 K) using liquid hydrogen, gaseous helium, or liquid nitrogen as the working fluid, thus supporting direct cooled machines below 77 K.

Description: A theoretical analysis of the radial temperature distribution through the rotor and constant cross sectional area blades near the coolant passages of liquid cooled gas turbines was made. The analysis was applied to obtain the rotor and blade temperatures of a specific turbine using a gas flow of 55 pounds per second, a coolant flow of 6.42 pounds per second, and an average coolant temperature of 200 degrees F. The effect of using kerosene, water, and ethylene glycol was determined. The effect of varying blade length and coolant passage lengths with water as the coolant was also determined. The effective gas temperature was varied from 2000 degrees to 5000 degrees F in each investigation.

Description: A theoretical analysis of the cross-sectional temperature distribution of a water-cooled turbine blade was made using the relaxation method to solve the differential equation derived from the analysis. The analysis was applied to specific turbine blade and the studies icluded investigations of the accuracy of simple methods to determine the temperature distribution along the mean line of the rear part of the blade, of the possible effect of varying the perimetric distribution of the hot gas-to -metal heat transfer coefficient, and of the effect of changing the thermal conductivity of the blade metal for a constant cross sectional area blade with two quarter inch diameter coolant passages.

Description: The performance at inlet pressure of 21 inches mercury absolute and inlet temperature of 538 R for the 10-stage axial-flow X24C-2 compressor from the X24C-2 turbojet engine was investigated. the peak adiabatic temperature-rise efficiency for a given speed generally occurred at values of pressure coefficient fairly close to 0.35.For this compressor, the efficiency data at various speeds could be correlated on two converging curves by the use of a polytropic loss factor derived.

Description: On the basis of the investigations so far completed on the behavior of PTL power plants under various operating conditions, in which the influence of the propeller characteristics is of considerable importance, the most important aspects of a control system for turbine-propeller jet power plants are deduced. A simple possible means for its concrete realization, which is also applicable to TL [NACA comment: TL, jet] power plants, is presented by means of examples. A control device of this kind is now being developed.

Description: A theoretical analysis of the temperature distribution through the trailing portion of a blade near the coolant passages of liquid cooled gas turbines was made. The analysis was applied to obtain the hot spot temperatures at the trailing edge and influence of design variables. The effective gas temperature was varied from 2000 degrees to 5000 degrees F in each investigation.

Description: Axial blowers are gaining importance as aircraft engine superchargers. However, the pressure head obtainable per stage is small. Due to the necessary great number of stages, the physical length of the blower becomes too great for an airworthy device. This report discusses several types of construction that permit a reduction in the length of the blower.

Description: An altitude-wind-tunnel investigation of a TG-100A gas turbine-propeller engine was performed. Pressure and temperature data were obtained at altitudes from 5000 to 35000 feet, compressor inlet ram-pressure ratios from 1.00 to 1.17, and engine speeds from 800 to 13000 rpm. The effect of engine speed, shaft horsepower, and compressor-inlet ram-pressure ratio on pressure and temperature distribution at each measuring station are presented graphically.

Description: Rim cracking in turbine wheels with welded blades was evaluated. The problem is explained on the basis of the occurrence of plastic flow in the rim during transient starting conditions when thermal compressive stresses resulting from high-temperature gradients exceed the proportional elastic limit of the material.

Description: While the gas turbine by itself has been applied in particular cases for power generation and is in a state of promising development in this field, it has already met with considerable success in two cases when used as an exhaust turbine in connection with a centrifugal compressor, namely, in the supercharging of combustion engines and in the Velox process, which is of particular application for furnaces. In the present paper the most important possibilities of combining a combustion engine with a gas turbine are considered. These "combination engines " are compared with the simple gas turbine on whose state of development a brief review will first be given. The critical evaluation of the possibilities of development and fields of application of the various combustion engine systems, wherever it is not clearly expressed in the publications referred to, represents the opinion of the author. The state of development of the internal-combustion engine is in its main features generally known. It is used predominantly at the present time for the propulsion of aircraft and road vehicles and, except for certain restrictions due to war conditions, has been used to an increasing extent in ships and rail cars and in some fields applied as stationary power generators. In the Diesel engine a most economical heat engine with a useful efficiency of about 40 percent exists and in the Otto aircraft engine a heat engine of greatest power per unit weight of about 0.5 kilogram per horsepower.

Description: After defining the aims and requirements to be set for a control system of gas-turbine power plants for aircraft, the report will deal with devices that prevent the quantity of fuel supplied per unit of time from exceeding the value permissible at a given moment. The general principles of the actuation of the adjustable parts of the power plant are also discussed.

Description: The performance of the 11-stage axial-flow compressor, modified to improve the compressor-outlet velocity, in a revised X24C-4B turbojet engine is presented and compared with the performance of the compressor in the original engine. Performance data were obtained from an investigation of the revised engine in the MACA Cleveland altitude wind tunnel. Compressor performance data were obtained for engine operation with four exhaust nozzles of different outlet area at simulated altitudes from 15,OOO to 45,000 feet, simulated flight Mach numbers from 0.24 to 1.07, and engine speeds from 4000 to 12,500 rpm. The data cover a range of corrected engine speeds from 4100 to 13,500 rpm, which correspond to compressor Mach numbers from 0.30 to 1.00.

Description: Requirements of an automatic engine control, as affected by engine characteristics, have been analyzed for a direct-coupled turbojet engine. Control parameters for various conditions of engine operation are discussed. A hypothetical engine control is presented to illustrate the use of these parameters. An adjustable speed governor was found to offer a desirable method of over-all engine control. The selection of a minimum value of fuel flow was found to offer a means of preventing unstable burner operation during steady-state operation. Until satisfactory high-temperature-measuring devices are developed, air-fuel ratio is considered to be a satisfactory acceleration-control parameter for the attainment of the maximum acceleration rates consistent with safe turbine temperatures. No danger of unstable burner operation exists during acceleration if a temperature-limiting acceleration control is assumed to be effective. Deceleration was found to be accompanied by the possibility of burner blow-out even if a minimum fuel-flow control that prevents burner blow-out during steady-state operation is assumed to be effective. Burner blow-out during deceleration may be eliminated by varying the value of minimum fuel flow as a function of compressor-discharge pressure, but in no case should the fuel flow be allowed to fall below the value required for steady-state burner operation.

Description: The pressure gain combustion (PGC) community is currently investigating rotating detonation engine (RDE) configurations where the flow direction is predominantly radial while the detonation travels circumferentially. These configurations are sometimes referred to as disk rotating detonation engines (DRDE) due to their nominal appearance as two disks in parallel with a gap between them. Having radial flow between disks, as opposed to the conventional RDE with axial flow in an annulus, may have profound effects on both the flow field and the performance. It may also yield extraordinarily compact devices which are well suited to particular propulsion and power applications. This presentation describes a preliminary effort to model the DRDE using a modified computational fluid dynamics (CFD) code originally written for analyzing ordinary RDE's. The quasi-two-dimensional code modifications are described, and some simple test flows are analyzed to insure that the modifications are functioning as envisioned. The code is then used to examine several DRDE scenarios such as radially inward and radially outward devices to see if stable operation is possible and if so, to assess the performance in terms of pressure gain. It is found that several flow scenarios are not only stable, but show superior performance to the ordinary RDE.

Description: No abstract available

Description: The present treatise reports on theoretical investigations and test-stand measurements which were carried out in the BMW Flugmotoren GMbH in developing the hollow blade for exhaust gas turbines. As an introduction the temperature variation and the stress on a turbine blade for a gas temperature of 900 degrees and circumferential velocities of 600 meters per second are discussed. The assumptions onthe heat transfer coefficients at the blade profile are supported by tests on an electrically heated blade model. The temperature distribution in the cross section of a blade Is thoroughly investigated and the temperature field determined for a special case. A method for calculation of the thermal stresses in turbine blades for a given temperature distribution is indicated. The effect of the heat radiation on the blade temperature also is dealt with. Test-stand experiments on turbine blades are evaluated, particularly with respect to temperature distribution in the cross section; maximum and minimum temperature in the cross section are ascertained. Finally, the application of the hollow blade for a stationary gas turbine is investigated. Starting from a setup for 550 C gas temperature the improvement of the thermal efficiency and the fuel consumption are considered as well as the increase of the useful power by use of high temperatures. The power required for blade cooling is taken into account.

Description: Computations were made to determine the temperature distribution and cooling of solid gas-turbine blades.A range of temperatures was used from 1500 degrees to 2500 degrees F, blade-root temperatures from 100 degrees to 1000 degrees F, blade thermal conductivity from 8 to 220 BTU/(hr)(sq ft)(degrees F/ft), and net gas to metal heat transfer coefficients from 75 to 250 BTU/(hr)(sq ft)(degrees F).

Description: Effect of inlet-air pressure and temperature on the performance of the X24-2 10-Stage Axial-Flow Compressor from the X24C-2 turbojet engine was evaluated. Speeds of 80, 89, and 100 percent of equivalent design speed with inlet-air pressures of 6 and 12 inches of mercury absolute and inlet-air temperaures of approximately 538 degrees, 459 degrees,and 419 degrees R ( 79 degrees, 0 degrees, and minus 40 degrees F). Results were compared with prior investigations.

Description: A calulation of the flow in turbine blading is reported that includes the calculation of effect of centrifugal force. Frictional losses on the stator blades and rotor blades are allowed.

Description: An investigation of the antiknock effectiveness of various additive-water solutions when used as internal coolants has been conducted at the NACA Cleveland laboratory. Nine compounds have been previously run in a CFR engine and the results are presented. In an effort to find a good anti-knock-coolant additive with more desirable physical properties than those of the nine compounds previously investigated, water solutions of four alkyl amines, three alkanolamines, six amides, and eight heterocyclic compounds were investigated and the results are presented.

Description: An investigation was conducted to determine the operational and performance characteristics of the TG-100A gas turbine-propeller engine II. Windmilling characteristics were deterined for a range of altitudes from 5000 to 35,000 feet, true airspeeds from 100 to 273 miles per hour, and propeller blade angles from 4 degrees to 46 degrees.

Description: The sea-level performance of I-16 turbojet engine at zero ram was investigated to determine the effects of an intake duct, shroud, and tail pipe intended for installation in an XFR-1 airplane. Engine speeds ranged from 8000 to 16,500 rpm for several variations of the intake duct and tail pipes.

Description: The performance of a mixed-flow impeller in combination with a semivaneless diffuser were experimentally investigated. The diameter of the impeller was 11.0 inches and a maximum tip diameter of 14.74 inches. The semivaneless diffuser had an overall diameter of 28.00 inches. The performance properties of the mixed-flow impeller were also investigated with a 34.00 inch vane loss diffuser having a transition section of the same geometry as the semivaneless diffuser.

Description: The calculation of infinitesimal conical supersonic flow has been applied first to the simplest examples that have also been calculated in another way. Except for the discovery of a miscalculation in an older report, there was found the expected conformity. The new method of calculation is limited more definitely to the conical case.

Description: The Russian AM 35 and AM 38 aircraft engines have superchargers with a swirl throttle, which appears to be a purely Russian development. This paper gives the results of test runs of the two engines, including the effects of the swirl throttle on engine performance.

Description: A study was made of heat transfer in turbine blades and the effects on blade temperature of cooling the blade root and tip, changing the dimensions of the blades, raising the cycle temperatures, insulating with ceramics, and cooling by circulation of air or water through hollow blades.

Description: Four methods of boundary-layer control were tried during an investigation to improve the flow in the impeller passages of a V-1710-93 engine-stage supercharger. The boundary layer along the impeller front shroud was removed by suction. In one method the removal was accomplished by recirculation of the air to the impeller inlet; in another method, by external removal. In the other methods, slots were cut through the impeller-blade faces first at 30 percent and then at 30 and 70 percent of the mean-flow-path length measured from leading edges of the rotating inlet guide vanes to introduce air from the high-pressure side of the blades into the region where stagnation and separation were suspected. A slight improvement in performance was obtained when the boundary layer was removed through the impeller front shroud. In general, this improvement become more pronounced as the amount of air removed was increased even though the excessive impeller frontal clearance maintained for these tests, together with an exaggerated negative pressure gradient, apparently induced flow separation on the diffuser front and rear walls as well as on the impeller front shroud. The use of slots in the impellers at the locations selected had a detrimental effect on the supercharger performance characteristics.

Description: Magnetic gears are currently being developed for use in a variety of industries such as wind and automotive, because of their higher reliability and lower maintenance cost than their mechanical counterparts. The bulk of magnetic gear development to date has focused on maximizing the technology's volumetric torque density. In contrast, the primary performance metrics for an aircraft's gear box are its mass and efficiency. To that end this paper presents a study of the achievable electromagnetic specific torque and efficiency of concentric magnetic gears. NASA's second magnetic gear prototype is used as the baseline for this study. Achievable electromagnetic specific torque and efficiency trends are presented with respect to higher level design variables such as gear ratio and radius.

Description: Magnetic gears are an attractive alternative to mechanical gears for electrified aircraft drive systems due to their ability to transmit torque without mechanical tooth contact. Consequently, magnetic gears enable electrified aircraft to take advantage of the benefits of gearing without introducing most of the contact-related reliability concerns associated with mechanical gearing. Magnetic gears however, have not been shown to match the specific torque (torque/mass) and efficiency of their mechanical counterparts in an aerospace application to date. In this paper, the design of a concentric magnetic gear for a personal air transport NASA reference vehicle is presented to demonstrate the feasibility of a magnetic gear for aerospace applications.

Description: An investigation was conducted on a multicylinder aircraft engine on a dynamometer stand to determine the effect of induction-system icing on engine operating characteristics and to compare the results with those of a previous laboratory investigation in which only the carburetor and the engine-stage supercharger assembly from the engine were used. The experiments were conducted at simulated glide power, low cruise power, and normal rated power through a range of humidity ratios and air temperatures at approximately sea-level pressure. Induction-system icing was found to occur within approximately the same limits as those established by the previous laboratory investigation after making suitable allowances for the difference in fuel volatility and throttle angles. Rough operation of the engine was experienced when ice caused a marked reduction in the air flow. Photographs of typical ice formations from this investigation indicate close similarity to icing previously observed in the laboratory.

Description: It will be shown that by the use of the concept of similarity a simple representation of the characteristic curves of a compressor operating in combination with a turbine may be obtained with correct allowance for the effect of temperature. Furthermore, it bec~mes possible to simplify considerably the rather tedious investigations of the behavior of gas-turbine power plants under different operating conditions. Characteristic values will be derived for the most important elements of operating behavior of the power plant, which will be independent of the absolute valu:s of pressure and temperature. At the same time, the investigations provide the basis for scale-model tests on compressors and turbines.

Description: An investigation has been conducted in the NACA Cleveland altitude wind tunnel to evaluate the performance characteristics of a modified X24C-4B turbojet engine over a range of simulated altitudes from 5000 to 45,000 feet, simulated flight Mach numbers from 0.25 to 1.07, and engine speeds from 4000 to 12,500 rpm. The engine was modified by the manufacturer to improve the velocity and temperature profiles within the engine. Performance data are graphically presented to show the effect of altitude at a flight Mach number of 0.25 and the effect of flight Mach number at an altitude of 25,000 feet. Original and modified engine performances for several specific operating conditions are compared. A complete tabulation of average pressures and temperatures throughout the engine, performance data, and lubrication and fuel-system data is presented.

Description: An experiment is conducted with hot-wire anemometry to document the exit boundary layer characteristics of two nozzle configurations at jet Mach numbers up to 0.82. Far-field noise and jet plume experimental data from these two configurations have been used in Large Eddy Simulations (LES) of jets by colleagues at other Institutions. The current experiment provides the boundary layer data which have been identified as being critical for validation of the simulations since the initial conditions can significantly affect subsequent jet evolution and its radiated noise. The data exhibit fully turbulent boundary layers for the case with a pipe attached upstream of the nozzle. The case without the pipe involves Blasius-like mean velocity profiles but a highly disturbed laminar state with large turbulence intensities in a range of subsonic Mach numbers.

Description: Electrified aircraft propulsion seeks to address ambitious goals in the commercial airline industry, including significant decreases in fuel burn, emissions, noise, and takeoff field length. In order to move these electrified propulsion concepts forward, analysis tools are needed that can model propulsion systems containing both gas turbine and power system components. This work presents the definition of an electric port, a set of electrical power systems tools, and simulation examples for the Numerical Propulsion System Simulation (NPSS) software. NPSS is the industry standard modeling and simulation package for aircraft propulsion systems, and the ability to design, size, integrate, and analyze electric power systems will enable industry efforts towards the development of electrified aircraft propulsion.

Description: An overview is given of an effort that focused on using CFD analysis to complement design and configuration definition of Lean-Direct Injection (LDI) combustion concepts for NASA's Commercial Supersonic Transport (CST) program. The National Combustion Code (OpenNCC) was used to perform non-reacting and two-phase reacting flow computations for second and third generation LDI configurations at CST cruise conditions. All computations were performed with a consistent approach of mesh-generation, spray modeling, ignition and kinetics modeling. Emissions (EINOx) characteristics were predicted for CST cruise conditions, and compared with emissions data from experimental measurements to evaluate the fidelity of the CFD modeling approach to predict emissions changes in response to changes in supersonic cycle conditions.

Description: Some of the challenges associated with developing electric aircraft propulsion systems include developing powertrain components that are both efficient and light-weight. In particular, electric motors must simultaneously achieve high efficiency by minimizing electrical and mechanical losses while also achieving high specific power by increasing the torque and/or speed. Normally increasing torque or speed will increase electrical and mechanical losses. The High Efficiency Megawatt Machine (HEMM) minimizes electrical losses by incorporating a superconductor to enable increased current on the rotor. And the rotor spins in a vacuum to minimize thermal and mechanical losses. Some organizations have been developing superconducting rotors for similar reasons using either cryogenic fluid transfer systems, fully immersed cryogenic cooling, and in a few cases utilized built-in cryogenic cooling on the rotor using a Brayton or Stirling system but the implementation was too large or inefficient for effective motor integration. Instead, a new approach for cryogenically cooling the superconducting rotor coil with an embedded rotating cryocooler is presented that fits completely within the rotating shaft.

Description: Time accurate simulation of non-equilibrium flows inside shock tube facilities presents several challenges from both physical and mathematical aspects. Furthermore, the drastic computational cost makes it non-practical to support real-time experimental test campaign. In this work, we explore other methods for modeling the shock tube prob- lem with the main focus on the post-shock region and the absolute radiation emanating from it. The proposed alternative approach is several orders of magnitude less computa- tionally expansive while still accurate enough with regards to the quantities of interest. Excellent agreement is found with the well-established stagnation-line approach. Comparison with the time-accurate simulation shows good agreement close to the peak values and disagreement of the temperatures relaxation and radiance profiles toward equilibrium, due to shock speed unsteadiness.

Description: The X-57 60kW Permanent Magnet Synchronous Motor for cruise applications was modeled utilizing a two-dimensional electromagnetics simulation software called Finite Element Method Magnets (FEMM, D. Meeker). Through FEMM, the simulated induction and torque characteristics of the X-57 PMSM were obtained. These parameters and other values were compared to actual static laboratory measurements. A three-dimensional electromagnetic model of the X-57 cruise motor was created utilizing OperaFEA (Dassault Systemes SE, Velizy-Villacoublay, France). Torque, RPM, power, resistance, and inductance characteristics were examined along with establishing work to begin examining heat flow and heat dissipation for efficiency purposes.

Description: The adoption of SiC devices in high power applications enables higher switching speed, which requires lower circuit parasitic inductance to reduce the voltage overshoot. This paper presents the design of a busbar for a 500 kVA three-level active natural clamped converter. The layout of the busbar is discussed in detail based on the analysis of the multiple commutation loops, magnetic cancelling effect, and DC-link capacitor placement. The loop inductance of the designed busbar is verified with simulation, impedance measurements and converter experiment. The results can match with each other and the inductances of small and large loop are 6.5 nH and 17.5 nH respectively, which is significantly lower than the busbars of NPC type converters in other references.

Description: NASA is broadly engaged in Electrified Aircraft Propulsion (EAP) efforts across air vehicle sizes and electric aircraft propulsion approaches. EAP enables a wide range of propulsion airframe integration options as well as the use of rechargeable energy storage in an aircraft. This paper is limited to a discussion of boundary layer ingestion (BLI) systems which are located on the fuselage of the aircraft and use electrical drive systems. We term that combination an "electrical propulsive fuselage". The benefits, challenges, and design parameters of an electrically driven fuselage BLI system are considered. Five existing types of fuselage BLI implementation approaches which can be implemented using either electrical or mechanical drive systems are reviewed. An overview of boundary layer types, fan response to boundary layer, and electrical system for aircraft propulsion is presented. An idea distributed electric propulsive fuselage is proposed.

Description: An overview is given of an effort for the use of CFD analysis to complement design and configuration definition of third generation Lean-Direct Injection combustion concepts (LDI-3) for NASAs N+3 program. The National Combustion Code (OpenNCC) was used to perform non-reacting and two-phase reacting flow computations for a three-cup, nineteen-element flame tube array with redesigned pilot injectors to improve spray and emissions characteristics when compared to a previous LDI-3 design. All computations were performed with a consistent approach to mesh-generation, spray modeling, ignition and kinetics modeling for a medium-power cycle condition. Computational predictions of the aerodynamics of a new pre-filming pilot injector were used to arrive at an optimized aerothermal design that meets effective area and fuel-air mixing criteria. The newly designed pilot injectors were shown to provide considerable improvements in aerodynamic stability, flame-tube pattern factor and NOx emissions, when compared to the original design.

Description: Magnetic gears are currently being explored to replace mechanical gears in various industries such as wind and automotive due to their higher reliability and lower maintenance requirements. In these applications volume minimization has been the goal of magnetic gear development. In contrast, the primary performance metrics for electrified aircraft drives are mass and efficiency. This paper presents the first ever study of design tradeoffs between electromagnetic mass and efficiency of concentric magnetic gears and the feasibility of achieving the low mass and high efficiency required for electrified aircraft applications. Higher level design variables are considered, including gear ratio, number of magnetic pole pairs, and number of magnets per pole pair.

Description: This presentation describes experimental and computational approaches to measuring pressure gain in the various devices currently under investigation wherein the working fluid undergoes a pressure gain combustion (PGC) process. Pressure gain is essentially a measure of the fluid availability for work or thrust production. The devices covered are Resonant Pulse Combustors, Internal Combustion Wave Rotors, Pulse Detonation Engines, and Rotating Detonation Engines. The approaches to pressure gain measurement differ in each device. However, all of the approaches attempt to address the fundamental challenges of PGC system measurement: the extremely harsh environment which makes instrumentation difficult, and the temporal and spatial non-uniformity associated with the exhausting flow which makes assigning a single value to the total pressure difficult. As part of the two-day 2019 International Constant Volume and Detonative Combustion Workshop, held in conjunction with the 2019 AIAA Propulsion and Energy Forum, this presentation is intended to foster discussion and eventual consensus on acceptable measurement methods.

Description: Transition from fossil fuels to synthetic drop-in fuels without the need to change existing combustors is the current research topic. The combustor performances such as cold-day ignition limits, lean blow-out (LBO) limits and altitude relight limits are the main focus points. The objective of this work is to evaluate the effect of different fuel candidates on the operability of gas turbines by comparing a conventional petroleum-based fuel with one other alternative fuel candidate. Time filtered Navier-Stokes simulations (TFNS) and K-LES are performed to examine the performance of these fuels at the stable conditions close to blow-out in a referee combustor rig.

Description: This paper continues a parametric study in which we consider the effect of air swirler configuration on the flame structure and combustor performance using a circular 7-point Lean Direct Injector Array for gas turbine applications. The injector array consists of a center swirler element surrounded by six swirler elements. Parameters considered in this study include swirler angle (60 or 52), handedness (co-swirling or counter-swirling) and center swirler offset. The primary focus considers flame stability, comparing four key air swirler configurations: for 1) fuel-lean flames; 2) high cold flow air reference velocity flames. We determined that the baseline swirler configuration had the best lean stability and could sustain the highest reference velocity. For this baseline configuration, we also compare the lean-blowout limits of four aircraft gas turbine reference fuels. With regard to lean blow-out, we determined that C4 could sustain the leanest flame, followed closely by A2. A1 was a poor performer.

Description: In rotorcraft, one of the main sources of mechanical failure is the gearbox, because of the many wear and failure modes associated with tooth contact in traditional mechanical gear boxes. Magnetic gears transmit torque without mechanical tooth contact between gear bodies and therefore they have none of the tooth contact related failure modes associated with mechanical gearing. As a result, magnetic gears have the potential to enable more reliable rotorcraft gearboxes. However, magnetic gears have not been demonstrated to match the performance of mechanical gearboxes at a high enough technology readiness level (TRL) to be used on an aircraft to date. To that end, NASA's Revolutionary Vertical Lift Technologies project has made an investment in developing magnetic gearboxes specifically for electrified vertical lift vehicles (EVTOL). In this presentation, the results of that investment to date will be discussed.

Description: Presentation to the International Forum on Aviation (IFAR) at the Electric Hybrid Propulsion Workshop #2 in Budapest, Hungary. This presentation is to provide an overview of NASA's investments in electrified propulsion as a starting point for the workshop, which will concentrate on the safety of electrified airplanes and potential for international collaboration.

Description: No abstract available

Description: This is for an invited lecture at Cleveland State University for a combustion course. The presentation gives an overview of some of the optical diagnostic techniques the Combustion Branch uses to characterize research fuel injection concepts to reduce emissions. Examples are provided that come from previously presented conference papers

Description: As a means of preparing for high-altitude flight with spark-ignition engines in conjunction with exhaust-gas turbosuperchargers, various methods of modifying the exhaust-gas temperatures, which are initially higher than a turbine can withstand are mathematically compared. The thermodynamic results first obtained are then examined with respect to the effect on flight speed, climbing speed, ceiling, economy, and cruising range. The results are so presented in a generalized form that they may be applied to every appropriate type of aircraft design and a comparison with the supercharged engine without exhaust-gas turbine can be made.

Description: An investigation to determine the performance and operational characteristics of the TG-1OOA gas turbine-propeller engine was conducted in the Cleveland altitude wind tunnel. As part of this investigation, the combustion-chamber performance was determined at pressure altitudes from 5000 to 35,000 feet, compressor-inlet rm-pressure ratios of 1.00 and 1.09, and engine speeds from 8000 to 13,000 rpm. Combustion-chamber performance is presented as a function of corrected engine speed and.correcte& horsepower. For the range of corrected engine speeds investigated, over-all total-pressure-loss ratio, cycle efficiency, ana the frac%ional loss in cycle efficiency resulting from pressure losses in the combustion chambers were unaffected by a change in altitude or compressor-inlet ram-pressure ratio. The scatter of combustion- efficiency data tended to obscure any effect of altitude or ram-pressure ratio. For the range of corrected horse-powers investigated, the total-pressure-loss ratio an& the fractional loss in cycle efficiency resulting from pressure losses in the combustion chambers decreased with an increase in corrected horsepower at a constant corrected engine speed. The combustion efficiency remained constant for the range of corrected horse-powers investigated at all corrected engine speeds.

Description: Based upon a simplified representation of the mode of operation of the pulse-jet tube, the effect of the influences mentioned in the title were investigated and it will be shown that, for a jet tube with a fccmndesigned to be aerodynamically favorable, the ability to operate is at least questionable. By taking into account the course of the development of pressure by combustion, a new insight has been obtained into the processes of motion within the jet tube, an insight that explains a number of empirical observations, namely: certain particulars of the sequence of pressure variations; the existence of an optimum valve-opening ratio; the occurrence of an intrusion of air; and the existence of a flight speed above lrhichthe jet tube ceases to operate. At too great an opening ratio or at too great a flight s-peed, the continuous flow through the tube is too predominant over the oscilla~ory process to perinitthe occurrence of an explosion powerful enough to maintain continuous operation. Certain possible means of making the operation of the jet tube more independent of the flight speed and of reducing the flow losses were proposed and discussed.

Description: Efficiency investigations have been made on a single-stage modification of the turbine of a Mark 25 aerial torpedo to determine the performance of the unit with five different turbine nozzles. The output of the turbine blades was computed by analyzing the windage and mechanical-friction losses of the unit. The turbine was faund to be most efficient with a cast nozzle having sharp-edged inlets to the nine nozzle ports. An analysis af the effectiveness af the first and second stages of the standard Mark 25 torpedo turbine indicates that the first- stage turbine contributes nearly all the brake power produced at blade-jet speed ratios above 0.26.

Description: The performance of the 11-stage axial-flow compressor in the X24C-4B turbojet engine was analyzed on the basis of results obtained from an investigation of the complete engine in the NACA Cleveland altitude wind tunnel. The engine was operated with four, exhaust nozzles of different outlet area over a range of engine speeds from 6000 to 12,500 rpm, corrected engine speeds from approximately 6100 to 13,600 rpm, and compressor Mach numbers from 0.45 to 1.00. Data are presented for engine operation over a range of simulated altitudes from 15,000 to 45,000 feet and simulated flight Mach numbers from 0.24 to 1.08.

Description: Pressures and temperatures throughout the X24C-4B turbojet engine are presented in both tabular and graphical forms to show the effect of altitude, flight Mach number, and engine speed on the internal operation of the engine. These data were obtained in the NACA Cleveland altitude wind tunnel at simulated altitudes from 5000 to 45,000 feet, simulated flight Mach numbers from 0.25 to 1.08, and engine speeds from 4000 to 12,500 rpm. Location and detail drawings of the instrumentation installed at seven survey stations in the engine are shown. Application of generalization factors to pressures and temperatures at each measuring station for the range of altitudes investigated showed that the data did not generalize above an altitude of 25,000 feet. Total-pressure distribution at the compressor outlet varied only with change in engine speed. At altitudes above 35,000 feet and engine speeds above 11,000 rpm, the peak temperature at the turbine-outlet annulus moved inward toward the root of the blade, which is undesirable from blade-stress considerations. The temperature levels at the turbine outlet and the exhaust-nozzle outlet were lowered as the Mach number was increased. The static-pressure measurements obtained at each stator stage of the compressor showed a pressure drop through the inlet guide vanes and the first-stage rotor at high engine speeds. The average values measured by the manufacturer's instrumentation werein close agreement with the average values obtained with NACA instrumentation.

Description: A preliminary investigation of the over-all performance of a simply constructed, short-life, turbojet engine was conducted. The unit was operated at a pressure altitude of 15,000 feet for ram-pressure ratios of 1.2 t o 1.8. The corrected engine speed was varied from the minimum for good combustion to about 17,000 rpm, which is approximately 75 percent of rated speed. The performance is given by generalized parameters that permit the calculation of performance at any altitude. The corrected net thrust of the turbojet engine increased with ram-pressure ratio for a given corrected engine speed above 14,500 rpm and reached a maximum of 425 pounds at a ram-pressure ratio of 1.8 and a corrected engine speed of 16,650 rpm, The corrected thrust specific fuel consumption decreased with flight speed for corrected engine speeds higher than 13,600 rpm, The minimum corrected thrust specific fuel consumption of 1.48 was obtained at a ram-pressure ratio of 1,8 and a corrected engine speed of 15,000 rpm. For all ram-pressure ratios, choking occurred in the engine for corrected engine speeds greater than 14,500 rpm.

Description: An investigation has been conducted in the Cleveland altitude wind tunnel to determine the operational characteristics of the I-40 jet-propulsion engine over a range of pressure altitudes from 10,000 to 50,000 feet and ram-pressure ratios from 1.00 to 1.76. Engine operational data were obtained with the engine in the standard configuration and with various modifications of the fuel system, the electrical system, and the combustion chambers. The effects of altitude and airspeed on operating speed range, starting, windmilli.ng, acceleration, speed regulation, cooling, and vibration of the standard and modified engines were determined, and damage to parts was noted. Maximum engine speed was obtainable at all altitudes and airspeeds wi th each fuel-control system investigated. The minimum idling speed was raised by increases in altitude and airspeed. The lowest minimum stable speeds were obtained with the standard configuration using 40-gallon nozzles with individual metering plugs. The engine was started normally at altitudes as high as 20,000 feet with all of the fuel systems and ignition combinations except one. Ignition at 70,000 feet was difficult and, although successful ignition occurred, acceleration was slow and usually characterized by excessive tail-pipe temperature. During windmilling investigations of the engine equipped with the standard fuel system, the engine could not be started at ram-pressure ratios of 1.1 to 1.7 at altitudes of 10,000, 20,000 and 30,000 feet. When equipped with the production barometric and Monarch 40-gallon nozzles, the engine accelerated in 12 seconds from an engine speed of 6000 rpm to 11,000 rpm at 20,000 feet and an average tail-pipe temperature of 11000 F. At the same altitude and temperature, all the engine configurations had approximately the same rate of acceleration. The Woodward governor produced the safest accelerations, inasmuch as it could be adjusted to automatically prevent acceleration blow out. The engine speed was held constant by the Woodward governor and the Edwards regulator during simulated dives and climbs at constant throttle position. The bearing cooling system was satisfactory at all altitudes and airspeeds. The engines operated without serious failure, although the exhaust cone, the tail pipe, and the airplane fuselage were damaged during altitude starts.

Description: Performance characteristics of the turbine in the 19B-8 jet propulsion engine were determined from an investigation of the complete engine in the Cleveland altitude wind tunnel. The investigation covered a range of simulated altitudes from 5000 to 30,000 feet and flight Mach numbers from 0.05 to 0.46 for various tail-cone positions over the entire operable range of engine speeds. The characteristics of the turbine are presented as functions of the total-pressure ratio across the turbine and the turbine speed and the gas flow corrected to NACA standard atmospheric conditions at sea level. The effect of changes in altitude, flight Mach number, and tail-cone position on turbine performance is discussed. The turbine efficiency with the tail cone in varied from a maximum of 80.5 percent to minimum of 75 percent over a range of engine speeds from 7500 to 17,500 rpm at a flight Mach number of 0.055. Turbine efficiency was unaffected by changes in altitude up to 15,000 feet but was a function of tail-cone position and flight Mach number. Decreasing the tail-pipe-nozzle outlet area 21 percent reduced the turbine efficiency between 2 and 4.5 percent. The turbine efficiency increased between 1.5 and 3 percent as the flight Mach number changed from 0.055 to 0.297.

Description: An investigation was conducted to compare the knock-limited performance of a 20-percent triptane blend in 28-K fuel with that of 28-R and 33-R fuels at high engine speeds, cruising speeds, and two compression ratios in an K-1830-94 multicylinder engine, Data were obtained with the standard compression ratio of 6.7 and with a compression ratio of 3.0, The three fuels were investigated at engine speeds of 1800, 2250, 2600, and 2800 rpm at high and low blower ratios. A carburetor-air temperature of approximate1y 100 deg F was maintained for the multicylinder-engine runs, Data were obtained on a single R-1830-94 cylinder engine as a means of checking the multicylinder data at the higher speeds. A satisfactory correlation between average mixture temperature and knock-limited manifold pressure was obtained by plotting knock-limited manifold pressure against average mixture temperature for the whole range of engine speeds at constant carburetor air temperature and cylinder-head temperature. The single-cylinder knock-limited performance based on charge-air flow matched that of the multicylinder engine within 6 percent under all the conditions except for 28-R fuel at 2800 rpm; these curves differed from each other by 11 percent in the rich region. The knock rating of 33-R fuel was found to be a little higher than that of the 20-percent triptane blend and 26-R fuel at high mixture temperatures (above 210 deg F) and lean mixtures. The 33-R fuel exhibited rich knock limits appreciably lower than the 20-percent triptane blend, Increasing the compression ratio from 6.7 to 8.0 lowered the knock-limited manifold pressure for all fuels approximately 15 to 18 inches of mercury absolute in the cruising range and 20 to 28 inches of mercury absolute at higher engine speeds. Brake specific fuel consumption was reduced 7 to 9 percent by the increase in compression ratio from 6.7 to 8,0,

Description: A knock-limited performance investigation was conducted on blends of triptane and 28-P fuel with a 12-cylinder, V-type, liquid-cooled aircraft engine of 1710-cubic-inch displacement at three compression ratios: 6.65, 7.93, and 9.68. At each compression ratio, the effect of changes in temperature of the inlet air to the auxiliary-stage supercharger and in fuel-air ratio were investigated at engine speeds of 2280 and. 3000 rpm. The results show that knock-limited engine performance, as improved by the use of triptane, allowed operation at both take-off and cruising power at a compression ratio of 9.68. At an inlet-air temperature of 60 deg F, an engine speed of 3000 rpm ; and a fuel-air ratio of 0,095 (approximately take-off conditions), a knock-limited engine output of 1500 brake horsepower was possible with 100-percent 28-R fuel at a compression ratio of 6.65; 20-percent triptane was required for the same power output at a compression ratio of 7.93, and 75 percent at a compression ratio of 9.68 allowed an output of 1480 brake horsepower. Knock-limited power output was more sensitive to changes in fuel-air ratio as the engine speed was increased from 2280 to 3000 rpm, as the compression ratio is raised from 6.65 to 9.68, or as the inlet-air temperature is raised from 0 deg to 120 deg F.

Description: An investigation has been conducted to determine thermal and pressure-drop performance and the operational characteristics of a Stewart-Warner model 906-B combustion heater. The performance tests covered a range of ventilating-air flows from 500 to 3185 pounds per hour, combustion-air pressure drops from 5 to 35 inches of water, and pressure altitudes from sea level to 41,000 feet. The operational characteristics investigated were the combustion-air flows for sustained combustion and for consistent ignition covering fuel-air ratios ranging from 0.033 to 0.10 and pressure altitudes from sea level to 45,000 feet. Rated heat output of 50,000 Btu per hour was obtained at pressure altitudes up to 27,000 feet for ventilating-air flows greater than 800 pounds per hour; rated output was not obtained at ventilating-air flow below 800 pounds per hour at any altitude. The maximum heater efficiency was found to be 60.7 percent at a fuel-air ratio of 0.050, a sea-level pressure altitude, a ventilating-air temperature of 0 F, combustion-air temperature of 14 F, a ventilating-air flow of 690 pounds per hour, and a combustion-air flow of 72.7 pounds per hour. The minimum combustion-air flow for sustained combustion at a pressure altitude of 25,000 feet was about 9 pounds per hour for fuel-air ratios between 0.037 and 0.099 and at a pressure altitude of 45,000 feet increased to 18 pounds per hour at a fuel-air ratio of 0.099 and 55 pounds per hour at a fuel-air ratio of 0.036. Combustion could be sustained at combustion-air flows above values of practical interest. The maximum flow was limited, however, by excessively high exhaust-gas temperature or high pressure drop. Both maximum and minimum combustion-air flows for consistent ignition decrease with increasing pressure altitude and the two curves intersect at a pressure altitude of approximately 25,000 feet and a combustion-air flow of approximately 28 pounds per hour.

Description: Previous performance data of the 19XB axial-flow compressor indicated that the outlet guide vanes and possibly the inlet guide vanes were stalling. Calculations were made to determine if these adverse conditions could be eliminated and if the manufacturer's design specifications could be more nearly approached by altering the blade angles of the first few compression stages as well as the outlet guide vanes. With the blade angles altered, experimental data were taken at compressor speeds of 8500 to 17,000 rpm with inlet-air conditions of 7.4 inches of mercury absolute and 59 0 F. The temperature-rise efficiency increased with speed from 0.70 at 8500 rpm to 0.74 at 13,600 rpm and dropped gradually to 0.70 at 17,000 rpm. At the design speed of 17,000 rpm, the pressure ratio at the peak efficiency point was 3.63. The maximum pressure ratio at design speed was 4.15 at an equivalent weight flow of 29.8 pounds per second. The altered compressor operated very .near the design specifications of pressure ratio and equivalent weight flow. At the high speeds, the peak adiabatic temperature-rise efficiency was increased 0.02 to 0,06 by altering the blade angles. The peak pressure ratio was increased 0.29 at design speed (17,000 rpm) and 0.05 and 0.13 at 11,900 and 13,600 rpm, respectively. The equivalent weight flow through the altered compressor was reduced 2 pounds per second at 15,300 and 17,000 rpm, as was expected from the design calculations. As extreme caution was taken not to surge the compressor violently, the point of minimum air flow may not have been reached in the present investigation and in a previous investigation. A true comparison of the pressure ratios obtained at the high speeds therefore cannot be made.

Description: Data for a liquid-cooled engine with a displacement volume of 1710 cubic inches were analyzed to determine the effect of exhaust pressure on the engine cooling characteristics. The data covered a range of exhaust pressures from 7 to 62 inches of mercury absolute, inlet-manifold pressures from 30 to 50 inches of mercury absolute, engine speeds from 1600 to 3000 rpm, and fuel-air ratios from 0.063 to 0.100. The effect of exhaust pressure on engine cooling was satisfactorily incorporated in the NACA cooling-correlation method as a variation in effective gas temperature with exhaust pressure. Large variations of cylinder-head temperature with exhaust pressure were obtained for operation at constant charge flow. At a constant charge flow of 2 pounds per second (approximately 1000 bhp) and a fuel-air ratio of 0.085, an increase in exhaust pressure from 10 to 60 inches of mercury absolute resulted in an increase of 40 F in average cylinder-head temperature. For operation at constant engine speed and inlet-manifold pressure and variable exhaust pressure (variable charge flow), however, the effect of exhaust pressure on cylinder-head temperature is small. For example, at an inlet-manifold pressure of 40 inches of mercury absolute, an engine speed of 2400 rpm.- and a fuel-air ratio of 0.085, the average cylinder-head temperature was about the same at exhaust pressures of 10 and 60 inches of,mercury absolute; a rise and a subsequent decrease of about 70 occurred between these extremes.

Description: A study of the data obtained in a flight investigation of an R-2800-21 engine in a P-47G airplane was made to determine the effect of the flight variables on the engine cooling-air pressure distribution. The investigation consisted of level flights at altitudes from 5000 to 35,000 feet for the normal range of engine and airplane operation. The data showed that the average engine front pressures ranged from 0.73 to 0.82 of the impact pressure (velocity head). The average engine rear pressures ranged from 0.50 to 0.55 of the impact pressure for closed cowl flaps and from 0.10 to 0.20 for full-open cowl flaps. In general, the highest front pressures were obtained at the bottom of the engine. The rear pressures for the rear-row cylinders were .lower and the pressure drops correspondingly higher than for the front-row cylinders. The rear-pressure distribution was materially affected by cowl-flap position in that the differences between the rear pressures of the front-row and rear-row cylinders markedly increased as the cowl flaps were opened. For full-open cowl flaps, the pressure drops across the rear-row cylinders were in the order of 0.2 of the impact pressure greater than across the front-row cylinders. Propeller speed and altitude had little effect on the -coolingair pressure distribution, Increase in angle of inclination of the thrust axis decreased the front ?pressures for the cylinders at the top of the engine and increased them for the cylinders at the bottom of the engine. As more auxiliary air was taken from the engine cowling, the front pressures and, to a lesser extent, the rear pressures for the cylinders at the bottom of the engine decreased. No correlation existed between the cooling-air pressure-drop distribution and the cylinder-temperature distribution.

Description: At the request of the Air Materiel Command, Army Air Forces, an investigation is being conducted at the NACA Cleveland laboratory to determine the performance characteristics of the XJ-41-V turbojet-engine compressor. The static-pressure variation in the direction of flow through the compressor was presented in reference 1 for an equivalent speed of 8000 rpm. An analysis of these pressure indicated that the maximum-flow limitation of the compressor was caused by separation, which reduced the effective flow area at the vaned-collector entrance. As a result of this analysis, the flow area at the vaned-collector entrance was increased to obtain larger mass flows. The area increase was obtained by cutting back the entrance edges of the collector vanes, which resulted in an increased vaneless-diffuser radius. Comparative performance of the original and revised compressors at an equivalent speed of 8000 rpm is presented. The static-pressure rise through the compressor, determined from static pressures at the impeller entrance and the vaned-collector exit, is also presented together with the compressor adiabatic efficiency and the mass flow over an equivalent speed range from 5000 to 9000 rpm. These static-pressure data are presented for the purpose of correlating the compressor performance with the turbojet-engine performance.

Description: Porous Microstructure Analysis (PuMA) software is used to perform simulations of molecular beam scattering experiments of hyperthermal atomic oxygen striking FiberFormr, a carbon preform material used commonly as a precursor in thermal protection systems (TPS). The purpose of this study is to investigate the reactive interaction of fibrous carbon with atomic oxygen in a complex microstructure, which is the primary source of carbon removal at lower temperatures. The detailed micro-structure of FiberFormr obtained from X-ray micro-tomography is used in the PuMA simulations to capture the complexity of the porous and fibrous characteristic of FiberFormr. A finite-rate surface chemistry model recently constructed from the molecular beam scattering experiments on vitreous carbon is applied to each fiber of the FiberFormr material. This model consists of detailed surface reaction mechanisms such as adsorption, desorption, and several types of Langmuir-Hinshelwood (LH) reactions to characterize the oxygen-carbon interactions at the surface. Comparison between the experimental and PuMA time-of-flight (TOF) distributions of both O and CO show good agreement. It is also found that a significantly higher amount of CO is generated when the beam interacted with FiberFormr, when compared with vitreous carbon. This is postulated to be primarily a result of multiple collisions of oxygen with the fibers, resulting in an higher effective rate of CO production. Multiple collisions with the different fibers, resulting from the porous nature of FiberFormr is also found to thermalize the O atoms, in addition to the adsorption/desorption process. The effect of micro-structure is concluded to be crucial in determining the final composition and energy distributions of the products. Thus, an effective model for the oxygen interaction with FiberFormr, fully accounting for the detailed micro-structure, for use in Computational Fluid Dynamics (CFD) and material response codes, is presented.

Description: MIT, Aurora Flight Sciences, and USC have collaborated to assess the feasibility of electric, hybridelectric, and turbo-electric propulsion for ultra-efficient commercial transportation. The work has drawn on the team expertise in disciplines related to aircraft design, propulsion-airframe integration, electric machines and systems, engineering system design, and optimization. A parametric trade space analysis has been carried out to assess vehicle performance across a range of transport missions and propulsion architectures to establish how electrified propulsion systems scale. An optimization approach to vehicle conceptual design modeling was taken to enable rapid multidisciplinary design space exploration and sensitivity analysis. The results of the analysis indicate vehicle aero-propulsive integration benefits enabled by electrification are required to offset the increased weight and loss associated with the electric system and achieve enhanced performance; the report describes the conceptual configurations than can offer such enhancements. The main contribution of the present work is the definition of electric vehicle design attributes for potential efficiency improvements at different scales. Based on these results, key areas for future research are identified, and extensions to the trade space analysis suitable for higher fidelity electrified commercial aircraft design and analysis have been developed.

Description: Hybrid electric propulsion architectures provide the infrastructure to enable additional benefits to the propulsion system that are otherwise unrealizable with the sole use of the current, state-of-the-art, gas-driven, turbine engines. The presence of electric machines (EMs) coupled to the shaft(s) of the turbine engine provide the ability to actively alter the operation of the engine to the benefit of the propulsion system and the aircraft it propels. This is the goal of the Turbine Electrified Energy Management (TEEM) concept, which at its broadest level addresses the management of energy across the electrified propulsion system. Prior work has demonstrated the use of this concept to alter steady-state operation and improve transient operability of a hybrid-electric propulsion system. The main benefits previously illustrated include the elimination of stability bleeds and expansion of the turbomachinery design space in order to enable more efficient designs. This paper focuses on the development of control strategies to implement the TEEM concept, and it explores several possible architecture variants for applying this concept. Comparison studies are conducted between a purely gas-driven turbofan (baseline engine configuration) and TEEM augmented variants of the baseline engine. The variants are distinguished by the shaft(s) that possess an EM. The configurations consider EMs on both shafts, an EM on the high pressure spool (HPS) only, and an EM on the low pressure spool (LPS) only. These configurations are referred to as the dual-spool configuration, the HPS configuration, and LPS configuration, respectively. The studies expose several options in configuring and controlling the system, including the use of a single EM coupled to a single shaft of a two-spool engine to positively impact the operability of both shafts. The studies also demonstrate the use of independently designed controllers for the electric machine(s) that allow for a decoupled control design process.

Description: NASA's broad investments in Electrified Aircraft Propulsion (EAP) are reviewed in this paper. NASA investments are guided by an assessment of potential market impacts, technical key performance parameters, and technology readiness attained through a combination of studies, enabling fundamental research, and flight research. NASA has determined that the impact of EAP varies by market and NASA is considering three markets: national/international, on-demand mobility, and short haul regional air transport. Flight research is underway to demonstrate integrated solutions and inform standards and certification processes. This paper focuses on the vehicle related activities, however there are related NASA activities in air space management and vehicle autonomy activities as well as a breakthrough technology project called the Convergent Aeronautics Solutions Project. A key finding is that sufficient technical advances in key areas have been made which indicate EAP is a viable technology for aircraft. Significant progress has been made to reduce EAP adoption barriers and further work is needed to transition the technology to a commercial product and improve the technology so it is applicable to large transonic aircraft. This paper will review the activities of the Hybrid Gas Electric Subproject of the Advanced Air Transport Technology Project, the Revolutionary Vertical Lift Technology Project, and the X-57 Flight Demonstration Project, and discuss the potential EAP benefits for commercial and military applications.

Description: An analysis was developed for calculating the radial temperature distribution in a gas turbine with only the temperatures of the gas and the cooling air and the surface heat-transfer coefficient known. This analysis was applied to determine the temperatures of a complete wheel of a conventional single-stage impulse exhaust-gas turbine. The temperatures were first calculated for the case of the turbine operating at design conditions of speed, gas flow, etc. and with only the customary cooling arising from exposure of the outer blade flange and one face of the rotor to the air. Calculations were next made for the case of fins applied to the outer blade flange and the rotor. Finally the effects of using part of the nozzles (from 0 to 40 percent) for supplying cooling air and the effects of varying the metal thermal conductivity from 12 to 260 Btu per hour per foot per degree Farenheit on the wheel temperatures were determined. The gas temperatures at the nozzle box used in the calculations ranged from 1600F to 2000F. The results showed that if more than a few hundred degrees of cooling of turbine blades are required other means than indirect cooling with fins on the rotor and outer blade flange would be necessary. The amount of cooling indicated for the type of finning used could produce some improvement in efficiency and a large increase in durability of the wheel. The results also showed that if a large difference is to exist between the effective temperature of the exhaust gas and that of the blade material, as must be the case with present turbine materials and the high exhaust-gas temperatures desired (2000F and above), two alternatives are suggested: (a) If metal with a thermal conductivity comparable with copper is used, then the blade temperature can be reduced by strong cooling at both the blade tip and root. The center of the blade will be less than 2000F hotter than the ends; (b) With low conductivity materials some method of direct cooling other than partial admission of cooling air is essential. From this study, it can be deduced that indirect cooling of turbine blades will not make possible large increases in gas temperature.

Description: An analysis is presented of rim cooling of gas-turbine blades; that is, reducing the temperature at the base of the blade (wheel rim), which cools the blade by conduction alone. Formulas for temperature and stress distributions along the blade are derived and, by the use of experimental stress-rupture data for a typical blade alloy, a relation is established between blade life (time for rupture), operating speed, and amount of rim cooling for several gas temperatures. The effect of blade parameter combining the effects of blade dimensions, blade thermal conductivity, and heat-transfer coefficient is determined. The effect of radiation on the results is approximated. The gas temperatures ranged from 1300F to 1900F and the rim temperature, from 0F to 1000F below the gas temperature. This report is concerned only with blades of uniform cross section, but the conclusions drawn are generally applicable to most modern turbine blades. For a typical rim-cooled blade, gas temperature increases are limited to about 200F for 500F of cooling of the blade base below gas temperature, and additional cooling brings progressively smaller increases. In order to obtain large increases in thermal conductivity or very large decreases in heat-transfer coefficient or blade length or necessary. The increases in gas temperature allowable with rim cooling are particularly small for turbines of large dimensions and high specific mass flows. For a given effective gas temperature, substantial increases in blade life, however, are possible with relatively small amounts of rim cooling.

Description: As part of an investigation of the performance and operational characteristics of the TG-100A gas turbine-propeller engine, conducted in the Cleveland altitude wind tunnel, the performance characteristics of the compressor and the turbine were obtained. The data presented were obtained at a compressor-inlet ram-pressure ratio of 1.00 for altitudes from 5000 to 35,000 feet, engine speeds from 8000 to 13,000 rpm, and turbine-inlet temperatures from 1400 to 2100R. The highest compressor pressure ratio was 6.15 at a corrected air flow of 23.7 pounds per second and a corrected turbine-inlet temperature of 2475R. Peak adiabatic compressor efficiencies of about 77 percent were obtained near the value of corrected air flow corresponding to a corrected engine speed of 13,000 rpm. This maximum efficiency may be somewhat low, however, because of dirt accumulations on the compressor blades. A maximum adiabatic turbine efficiency of 81.5 percent was obtained at rated engine speed for all altitudes and turbine-inlet temperatures investigated.

Description: A study has been made of the performance of the induction and the exhaust systems on the XR60 power-plant installation as part of an investigation conducted in the Cleveland altitude wind tunnel. Altitude flight conditions from 5000 to 30,000 feet were simulated for a range of engine powers from 750 to 3000 brake horsepower. Slipstream rotation prevented normal pressure recoveries in the right side of the main duct in the region of the right intercooler cooling-air duct inlet. Total-pressure losses in the charge-air flow between the turbosupercharger and the intercoolers were as high as 2.1 inches of mercury. The total-pressure distribution of the charge air at the intercooler inlets was irregular and varied as much as 1.0 inch of mercury from the average value at extreme conditions, Total-pressure surveys at the carburetor top deck showed a variation from the average value of 0.3 inch of mercury at take-off power and 0.05 inch of mercury at maximum cruising power, The carburetor preheater system increased the temperature of the engine charge air a maximum of about 82 F at an average cowl-inlet air temperature of 9 F, a pressure altitude of 5000 feet, and a brake horsepower of 1240.

Description: An investigation was conducted to determine the coolant-flow distribu tion, the cylinder temperatures, and the heat rejections of the V-165 0-7 engine . The tests were run a t several power levels varying from minimum fuel consumption to war emergency power and at each power l evel the coolant flows corresponded to the extremes of those likely t o be encountered in typical airplane installations, A mixture of 30-p ercent ethylene glycol and 70-percent water was used as the coolant. The temperature of each cylinder was measured between the exhaust val ves, between the intake valves, in the center of the head, on the exh aust-valve guide, at the top of the barrel on the exhaust side, and o n each exhaust spark-plug gasket. For an increase in engine power fro m 628 to approximately 1700 brake horsepower the average temperature for the cylinder heads between the exhaust valves increased from 437 deg to 517 deg F, the engine coolant heat rejection increased from 12 ,600 to 22,700 Btu. per minute, the oil heat rejection increased from 1030 to 4600 Btu per minute, and the aftercooler-coolant heat reject ion increased from 450 to 3500 Btu -per minute.

Description: A flight investigation of an I-16 jet propulsion engine installed in the waist compartment of a B-24M airplane was made to determine the effect of induction-system icing on the performance of the engine. Flights were made at inlet-air temperatures of 15 deg, 20 deg., and 25 F, an indicated airspeed of 180 miles per hour, jet-engine speeds of 13,000 and 15,000 rpm, liquid-water contents of approximately 0.3 to 0.5 gram per cubic meter, and an average water droplet size of approximately 50 microns. Under the most severe icing conditions obtained, ice formed on the screen over the front inlet to the compressor and obstructed about 70 percent of the front-inlet area. The thrust was thereby reduced 13.5 percent, the specific fuel consumption increased 17 percent, and the tail-pipe temperature increased 82 F. No icing of the rear compressor-inlet screen was encountered.

Description: An investigation was conducted on a 12-cylinder V-type liquid-cooled aircraft engine of 1710-cubic-inch displacement to determine the minimum specific fuel consumption at constant cruising engine speed and compression ratios of 6.65, 7.93, and 9.68. At each compression ratio, the effect.of the following variables was investigated at manifold pressures of 28, 34, 40, and 50 inches of mercury absolute: temperature of the inlet-air to the auxiliary-stage supercharger, fuel-air ratio, and spark advance. Standard sea-level atmospheric pressure was maintained at the auxiliary-stage supercharger inlet and the exhaust pressure was atmospheric. Advancing the spark timing from 34 deg and 28 deg B.T.C. (exhaust and intake, respectively) to 42 deg and 36 deg B.T.C. at a compression ratio of 6.65 resulted in a decrease of approximately 3 percent in brake specific fuel consumption. Further decreases in brake specific fuel consumption of 10.5 to 14.1 percent (depending on power level) were observed as the compression ratio was increased from 6.65 to 9.68, maintaining at each compression ratio the spark advance required for maximum torque at a fuel-air ratio of 0.06. This increase in compression ratio with a power output of 0.585 horsepower per cubic inch required a change from . a fuel- lend of 6-percent triptane with 94-percent 68--R fuel at a compression ratio of 6.65 to a fuel blend of 58-percent, triptane with 42-percent 28-R fuel at a compression ratio of 9.68 to provide for knock-free engine operation. As an aid in the evaluation of engine mechanical endurance, peak cylinder pressures were measured on a single-cylinder engine at several operating conditions. Peak cylinder pressures of 1900 pounds per square inch can be expected at a compression ratio of 9.68 and an indicated mean effective pressure of 320 pounds per square inch. The engine durability was considerably reduced at these conditions.

Description: The performance of a 24C-4 combustor was investigated with three different combustor baskets and five modifications of these baskets at conditions simulating static (zero-ram) operation of the 24C jet engine over ranges of altitude and engine speed to determine and improve the altitude operational limits of the 24C combustor. Information was also obtained regarding combustion characteristics, the fuel-flow characteristics of the fuel manifolds, and the combustor total-pressure drop. NACA modifications, which consisted of blocking rows of holes on the baskets, increased the minimum point on the altitude-operational-limit curve, which occurs at low engine speeds, for a narrow-upstream-end basket by 8000 feet (from 23, 000 to 31,000 ft_ and for a wide-upstream-end basket by 21,000 feet (from 12, 000 to 34,000 ft). These improvements were approximately maintained over the entire range of engine speeds investigated.

Description: A brief investigation has been made of the performance of a single combustor of the TG-180 turboJet engine to determine (a) the altitude operational limits of the engine for two fuels (AN-F-32 and AN-F-28), (b) combustion efficiencies at various simulated conditions of altitude and engine speeds, (c) combustion-outlet temperature distribution for several altitudes at constant engine speed, and (d) the combustor total pressure drop The limits with AN-83-F-32 fuel were found to be approximately 60,000 feet for an engine speed of 6000 rpm and approximately 38,000 feet for an engine speed of 1000 rpm. The results indicated that the altitude operational limits with AN-F-32 fuel are higher over the largest part of the engine-speed range than with AN-F-28 fuel, A combination efficiency of 22 percent was obtained at rated engine speed (7600 rpm) and an altitude of 20,000 feet with AN-F-32 fuel. A change in altitude from 20,000 tm 60,000 feet showed a 20-percent decrease in combustion efficiency while the engine was operating at 760G rpm whereas, at an engine speed of 4000 rpm a change of altitude from 10,000 to 40,000 feet showed a 52-percent decrease in combustion efficiency .

Description: An overview is given of an effort that focused on using CFD analysis to complement design and configuration definition of Lean-Direct Injection (LDI) combustion concepts for NASA's Commercial Supersonic Transport (CST) program. The National Combustion Code (OpenNCC) was used to perform non-reacting and two-phase reacting flow computations for second and third generation LDI configurations at CST cruise conditions. All computations were performed with a consistent approach of mesh-generation, spray modeling, ignition and kinetics modeling. Emissions (EINOx) characteristics were predicted for CST cruise conditions, and compared with emissions data from experimental measurements to evaluate the fidelity of the CFD modeling approach to predict emissions changes in response to changes in supersonic cycle conditions.

Description: Some of the challenges associated with developing electric aircraft propulsion systems include developing powertrain components that are both efficient and light-weight. In particular, electric motors must simultaneously achieve high efficiency by minimizing electrical and mechanical losses while also achieving high specific power by increasing the torque and/or speed. Normally increasing torque or speed will increase electrical and mechanical losses. The High Efficiency Megawatt Machine (HEMM) minimizes electrical losses by incorporating a superconductor to enable increased current on the rotor. And the rotor spins in a vacuum to minimize thermal and mechanical losses. Some organizations have been developing superconducting rotors for similar reasons using either cryogenic fluid transfer systems, fully immersed cryogenic cooling, and in a few cases utilized built-in cryogenic cooling on the rotor using a Brayton or Stirling system but the implementation was too large or inefficient for effective motor integration. Instead, a new approach for cryogenically cooling the superconducting rotor coil with an embedded rotating cryocooler is presented that fits completely within the rotating shaft.

Description: An experiment is conducted with hot-wire anemometry to document the exit boundary layer characteristics of two nozzle configurations at jet Mach numbers up to 0.82. Far-field noise and jet plume experimental data from these two configurations have been used in Large Eddy Simulations (LES) of jets by colleagues at other Institutions. The current experiment provides the boundary layer data which have been identified as being critical for validation of the simulations since the initial conditions can significantly affect subsequent jet evolution and its radiated noise. The data exhibit fully turbulent boundary layers for the case with a pipe attached upstream of the nozzle. The case without the pipe involves Blasius-like mean velocity profiles but a highly disturbed laminar state with large turbulence intensities in a range of subsonic Mach numbers.

Description: Considerable progress has, in recent times, been attained in the development of the high-pressure axial blower by well-planned research. The efforts are directed toward improving the efficiencies, which are already high for the axial blower, and in particular the delivery pressure heads. For high pressures multistage arrangements are used. Of fundamental importance is the careful design of all structural parts of the blower that are subject to the effects of the flow. In the present report, several recent results and experiences are reported, which are based on results of German engine research.

Description: Electrified aircraft propulsion (EAP) systems hold potential for the reduction of aircraft fuel burn, emissions, and noise. Currently, NASA and other organizations are actively working to identify and mature technologies necessary to bring EAP designs to reality. This paper specifically focuses on the envisioned control technology challenges associated with EAP designs that include gas turbine technology. Topics discussed include analytical tools for the dynamic modeling and analysis of EAP systems, and control design strategies at the propulsion and component levels. This includes integrated supervisory control facilitating the coordinated operation of turbine and electrical components, control strategies that seek to minimize fuel consumption and lessen the challenges associated with thermal management, and dynamic control to ensure engine operability during system transients. These dynamic control strategies include innovative control approaches that either extract or supply power to engine shafts dependent upon operating phase, which may improve performance and reduced gas turbine engine weight. Finally, a discussion of control architecture design considerations to help alleviate the propulsion/aircraft integration and certification challenges associated with EAP systems is provided.

Description: Electrified aircraft propulsion (EAP) systems hold potential for the reduction of aircraft fuel burn, emissions, and noise. Currently, NASA and other organizations are actively working to identify and mature technologies necessary to bring EAP designs to reality. This paper specifically focuses on the envisioned control technology challenges associated with EAP designs that include gas turbine technology. Topics discussed include analytical tools for the dynamic modeling and analysis of EAP systems, and control design strategies at the propulsion and component levels. This includes integrated supervisory control facilitating the coordinated operation of turbine and electrical components, control strategies that seek to minimize fuel consumption and lessen the challenges associated with thermal management, and dynamic control to ensure engine operability during system transients. These dynamic control strategies include innovative control approaches that either extract or supply power to engine shafts dependent upon operating phase, which may improve performance and reduced gas turbine engine weight. Finally, a discussion of control architecture design considerations to help alleviate the propulsion/aircraft integration and certification challenges associated with EAP systems is provided.

Description: The accuracy of the scale-resolving simulations for practical geometries strongly depends on the inflow boundary conditions. Imposing experimentally observed turbulent inflow profiles for the numerical simulations is a major challenge. Existing methods available in the literature assume self-similar behavior, which is not true for most of the experiments. In the present work, we formulate the turbulent inflow profile generation technique as an optimization problem. An adjoint technique is exploited to evaluate the sensitivities of multiple input parameters for the present problem. The present formulation is then tested to generate a laminar boundary layer profile, turbulent boundary layer profile, and turbulent jet profile.