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NASA NTRSDate not supplied

Modeling Turboshaft Engines for the Revolutionary Vertical Lift Technology Project, Expanded

Turboshaft engine performance and weight models were developed to support conceptual propulsion and vehicle mission design and performance under the Revolutionary Vertical Lift Technology (RVLT) Project in 2019 by Snyder in Ref. 1. These models were developed using open data sources, assuming present and future technology levels, and range from 650 to 5,000 output shaft horsepower (485 to 3,730 kW). This paper expands on the previous research, extending the power ranges from 200 to 15000 output shaft horsepower (150 to 11,200 kW) and documenting the methodology, assumptions, and engine performance realizes important benefits for NASA and the aviation community. NASA concept vehicle study efforts using these baseline propulsion models can be more readily shared among the government, industry, and university community to support present and future work. Assessing the benefits of advanced technologies and new configurations can be facilitated using these models, which helps guide technology investment. As the various 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 expand on the previous effort, the turboshaft engine is briefly discussed, highlighting the specific components, and expected performance characteristics over the updated power range and technology levels considered. Additional engine configurations will also be discussed as they vary based on power output and assumed technology level. Engine performance, such as airflow, power output and weight will be updated, noting trends that are important for system studies. Finally, the effect of advanced propulsion technologies on RVLT concept vehicles are also reported along with the tools and software methods used to complete the analysis.

VTOL↗

NASA NTRSDate not supplied

Modeling Turboshaft Engines for the Revolutionary Vertical Lift Technology Project, Expanded

Turboshaft engine performance and weight models were developed to support conceptual propulsion and vehicle mission design and performance under the Revolutionary Vertical Lift Technology (RVLT) Project in 2019 by Snyder in Ref. 1. These models were developed using open data sources, assuming present and future technology levels, and range from 650 to 5,000 output shaft horsepower (485 to 3,730 kW). This paper expands on the previous research, extending the power ranges from 200 to 15000 output shaft horsepower (150 to 11,200 kW) and documenting the methodology, assumptions, and engine performance realizes important benefits for NASA and the aviation community. NASA concept-vehicle study efforts using these baseline propulsion models can be more readily shared among the government, industry, and university community to support present and future work. Assessing the benefits of advanced technologies and new configurations can be facilitated using these models, which helps guide technology investment. As the various 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 expand on the previous effort, the turboshaft engine is briefly discussed, highlighting the specific components, and expected performance characteristics over the updated power range and technology levels considered. Additional engine configurations will also be discussed as they vary based on power output and assumed technology level. Engine performance, such as airflow, power output and weight will be updated, noting trends that are important for system studies. Finally, the effect of advanced propulsion technologies on RVLT concept vehicles are also reported along with the tools and software methods used to complete the analysis

VTOL↗

NASA NTRS2023

Modeling Turboshaft Engines for the Revolutionary Vertical Lift Technology Project, Expanded

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

VTOL↗

NASA NTRS1976

Feasibility study of modern airships, phase 2. Volume 1: Heavy lift airship vehicle. Book 1: Overall study results

A Heavy Lift Airship combining buoyant lift derived from a conventional helium-filled non-rigid airship hull with propulsive lift derived from conventional helicopter rotors was investigated. The buoyant lift essentially offsets the empty weight of the vehicle; thus the rotor thrust is available for useful load and to maneuver and control the vehicle. Such a vehicle is capable of providing a quantum increase in current vertical lifting capability. Certain critical deficiencies of past airships are significantly minimized or eliminated.

Source record↗

NASA NTRS1961

Take-off Distances of a Supersonic Transport Configuration as Affected by Airplane Rotation During the Take-off Run

The take-off distances over a 35-foot obstacle have been determined for a supersonic transport configuration characterized by a low maximum lift coefficient at a high angle of attack and by high drag due to lift. These distances were determined analytically by means of an electronic digital computer. The effects of rotation speed, rotation angle, and rotation time were determined. A few configuration changes were made to determine the effects of thrust-weight ratio, wing loading, maximum lift coefficient, and induced drag on the take-off distance. The required runway lengths based on Special Civil Air Regulation No. SR-422B were determined for various values of rotation speed and compared with those based on full engine power. Increasing or decreasing the rotation speed as much as 5 knots from the value at which the minimum take-off distance occurred increased the distance only slightly more than 1 percent for the configuration studied. Under-rotation by 1 deg to 1.5 deg increased the take-off distance by 9 to 15 percent. Increasing the time required for rotation from 3 to 5 seconds had a rather small effect on the take-off distance when the values of rotation speed were near the values which result in the shortest take-off distance. When the runway length is based on full engine power rather than on SR-422B, the rotation speed which results in the shortest required runway length is 10 knots lower and the runway length is 4.3 percent less.

Hall, Albert W.↗

NASA NTRS1977

Variable pitch fan system for NASA/Navy research and technology aircraft

Preliminary design of a shaft driven, variable-pitch lift fan and lift-cruise fan was conducted for a V/STOL Research and Technology Aircraft. The lift fan and lift-cruise fan employed a common rotor of 157.5 cm diameter, 1.18 pressure ratio variable-pitch fan designed to operate at a rotor-tip speed of 284 mps. Fan performance maps were prepared and detailed aerodynamic characteristics were established. Cost/weight/risk trade studies were conducted for the blade and fan case. Structural sizing was conducted for major components and weights determined for both the lift and lift-cruise fans.

Ryan, W. P.↗

NASA NTRS1974

A simplified flight-test method for determining aircraft takeoff performance that includes effects of pilot technique

A method for evaluating aircraft takeoff performance from brake release to air-phase height that requires fewer tests than conventionally required is evaluated with data for the XB-70 airplane. The method defines the effects of pilot technique on takeoff performance quantitatively, including the decrease in acceleration from drag due to lift. For a given takeoff weight and throttle setting, a single takeoff provides enough data to establish a standardizing relationship for the distance from brake release to any point where velocity is appropriate to rotation. The lower rotation rates penalized takeoff performance in terms of ground roll distance; the lowest observed rotation rate required a ground roll distance that was 19 percent longer than the highest. Rotations at the minimum rate also resulted in lift-off velocities that were approximately 5 knots lower than the highest rotation rate at any given lift-off distance.

Larson, T. J.↗

NASA NTRS2012

Development of Non-Optimum Factors for Launch Vehicle Propellant Tank Bulkhead Weight Estimation

Non-optimum factors are used during aerospace conceptual and preliminary design to account for the increased weights of as-built structures due to future manufacturing and design details. Use of higher-fidelity non-optimum factors in these early stages of vehicle design can result in more accurate predictions of a concept s actual weights and performance. To help achieve this objective, non-optimum factors are calculated for the aluminum-alloy gores that compose the ogive and ellipsoidal bulkheads of the Space Shuttle Super-Lightweight Tank propellant tanks. Minimum values for actual gore skin thicknesses and weld land dimensions are extracted from selected production drawings, and are used to predict reference gore weights. These actual skin thicknesses are also compared to skin thicknesses predicted using classical structural mechanics and tank proof-test pressures. Both coarse and refined weights models are developed for the gores. The coarse model is based on the proof pressure-sized skin thicknesses, and the refined model uses the actual gore skin thicknesses and design detail dimensions. To determine the gore non-optimum factors, these reference weights are then compared to flight hardware weights reported in a mass properties database. When manufacturing tolerance weight estimates are taken into account, the gore non-optimum factors computed using the coarse weights model range from 1.28 to 2.76, with an average non-optimum factor of 1.90. Application of the refined weights model yields non-optimum factors between 1.00 and 1.50, with an average non-optimum factor of 1.14. To demonstrate their use, these calculated non-optimum factors are used to predict heavier, more realistic gore weights for a proposed heavy-lift launch vehicle s propellant tank bulkheads. These results indicate that relatively simple models can be developed to better estimate the actual weights of large structures for future launch vehicles.

Wu, K. Chauncey↗

NASA NTRS1975

Ultra-heavy vertical lift system: The Heli-Stat

A hybrid VTOL airship which is combined with helicopters is evaluated. The static lift of the airship supports approximately the full empty weight of the entire assembly. The helicopter rotors furnish the lift to support the payload as well as the propulsion and control about all axes. Thus existing helicopters, with no new technology required, can be made to lift payloads of ten times the capacity of each one alone, and considerably more than that of any airship built so far. A vehicle is described which has a 75-ton payload, based on four existing CH-53D helicopters and an airship of 3,600,000 cu. ft. The method of interconnection is described along with discussion of control, instrumentation, drive system and critical design conditions. The vertical lift and positioning capabilities of this vehicle far exceed any other means available today, yet can be built with a minimum of risk, development cost and time.

Piasecki, F. N.↗

NASA NTRS1959

Wind-Tunnel Investigation of a Small-Scale Sweptback-Wing Jet-Transport Model Equipped with an External-Flow Jet-Augmented Double Slotted Flap

A wind-tunnel investigation at low speeds has been made to study the aerodynamic characteristics of a small-scale sweptback-wing Jet-transport model equipped with an external-flow jet-augmented double slotted flap. Included in the investigation were tests of the wing alone to study the effects of varying the spanwise extent of blowing on the full-span flap. The results indicated that the double-slotted-flap arrangement of the present investigation was more efficient in terms of lift and drag than were the external-flow single-slotted-flap arrangements previously tested and gave a substantial reduction In the thrust-weight ratio required for a given lift coefficient under trimmed drag conditions. An increase in the spanwise extent of blowing on the full-span flap was also found to increase the efficiency of the model in terms of the lift and drag but, as would be expected on a sweptback-wing configuration, was accompanied by significant increases in negative pitching moment.

Johnson, Joseph L., Jr.↗

NASA NTRS1993

Aerodynamic optimization of an HSCT configuration using variable-complexity modeling

An approach to aerodynamic configuration optimization is presented for the high-speed civil transport (HSCT). A method to parameterize the wing shape, fuselage shape and nacelle placement is described. Variable-complexity design strategies are used to combine conceptual and preliminary-level design approaches, both to preserve interdisciplinary design influences and to reduce computational expense. Conceptual-design-level (approximate) methods are used to estimate aircraft weight, supersonic wave drag and drag due to lift, and landing angle of attack. The drag due to lift, wave drag and landing angle of attack are also evaluated using more detailed, preliminary-design-level techniques. New, approximate methods for estimating supersonic wave drag and drag due to lift are described. The methodology is applied to the minimization of the gross weight of an HSCT that flies at Mach 2.4 with a range of 5500 n.mi. Results are presented for wing planform shape optimization and for combined wing and fuselage optimization with nacelle placement. Case studies include both all-metal wings and advanced composite wings.

Hutchison, M. G.↗

NASA NTRS1976

A rapid method for optimization of the rocket propulsion system for single-stage-to-orbit vehicles

A rapid analytical method for the optimization of rocket propulsion systems is presented for a vertical take-off, horizontal landing, single-stage-to-orbit launch vehicle. This method utilizes trade-offs between propulsion characteristics affecting flight performance and engine system mass. The performance results from a point-mass trajectory optimization program are combined with a linearized sizing program to establish vehicle sizing trends caused by propulsion system variations. The linearized sizing technique was developed for the class of vehicle systems studied herein. The specific examples treated are the optimization of nozzle expansion ratio and lift-off thrust-to-weight ratio to achieve either minimum gross mass or minimum dry mass. Assumed propulsion system characteristics are high chamber pressure, liquid oxygen and liquid hydrogen propellants, conventional bell nozzles, and the same fixed nozzle expansion ratio for all engines on a vehicle.

Eldred, C. H.↗

NASA NTRS1981

A study of high-altitude manned research aircraft employing strut-braced wings of high-aspect-ratio

The effect of increased wing aspect ratio of subsonic aircraft on configurations with and without strut bracing. Results indicate that an optimum cantilever configuration, with a wing aspect ratio of approximately 26, has a 19% improvement in cruise range when compared to a baseline concept with a wing aspect ratio of approximately 10. An optimum strut braced configuration, with a wing aspect ratio of approximately 28, has a 31% improvment in cruise range when compared to the same baseline concept. This improvement is mainly due to the estimated reduction in wing weight resulting from use of lifting struts. All configurations assume the same mission payload and fuel. The drag characteristics of the wings are enhanced with the use of laminar flow airfoils. A method for determining the extent of attainable natural laminar flow, and methods for preliminary structural design and for aerodynamic analysis of wings lifting struts are presented.

Smith, P. M.↗

NASA NTRS1982

A technique for determining powered-lift STOL aircraft performance at sea level from flight data taken at altitude

State-of-the-technology powered-lift STOL aircraft can develop very high lift coefficients. The determination of how this high lift capability can be utilized for landing approaches is accomplished by evaluating aircraft velocity versus flightpath performance. This paper briefly discusses upper-surface-blowing power-lift technology and some of the significant features of the velocity versus flightpath plot that are used in evaluating the performance of powered-lift aircraft. A procedure for deriving the velocity versus flightpath plot for a powered-lift aircraft at any specified weight at sea level, based on flight-test data obtained at various weights and altitudes, is discussed in detail. Actual flight-test data obtained with NASA's Quiet Short-Haul Research Aircraft (QSRA) are used to illustrate this procedure and to evaluate the results.

Stevens, V. C.↗

NASA NTRS1992

Application of Taguchi methods to dual mixture ratio propulsion system optimization for SSTO vehicles

The application of advanced technologies to future launch vehicle designs would allow the introduction of a rocket-powered, single-stage-to-orbit (SSTO) launch system early in the next century. For a selected SSTO concept, a dual mixture ratio, staged combustion cycle engine that employs a number of innovative technologies was selected as the baseline propulsion system. A series of parametric trade studies are presented to optimize both a dual mixture ratio engine and a single mixture ratio engine of similar design and technology level. The effect of varying lift-off thrust-to-weight ratio, engine mode transition Mach number, mixture ratios, area ratios, and chamber pressure values on overall vehicle weight is examined. The sensitivity of the advanced SSTO vehicle to variations in each of these parameters is presented, taking into account the interaction of each of the parameters with each other. This parametric optimization and sensitivity study employs a Taguchi design method. The Taguchi method is an efficient approach for determining near-optimum design parameters using orthogonal matrices from design of experiments (DOE) theory. Using orthogonal matrices significantly reduces the number of experimental configurations to be studied. The effectiveness and limitations of the Taguchi method for propulsion/vehicle optimization studies as compared to traditional single-variable parametric trade studies is also discussed.

Stanley, Douglas O.↗

NASA NTRS1993

Navy GTE seal development activity

Under the auspices of the Integrated High Performance Turbine Engine Technology Initiative, the Naval Air Warfare Center conducts advanced development programs for demonstration in the next generation of air-breathing propulsion systems. Among the target technologies are gas path and lube oil seals. Two development efforts currently being managed by NAWCAD are the High Performance Compressor Discharge Film-Riding Face Seal and the Subsonic Core High Speed Air/Oil Seal. The High Performance Compressor Discharge Film-Riding Face Seal Program aims at reducing parasitic leakage through application of a film-riding face sea concept to the compressor discharge location of a Phase 2 IHPTET engine. An order-of-magnitude leakage reduction relative to current labyrinth seal configurations is expected. Performance goals for these seals are (1) 1200 F air temperature, (2) 800 feet-per-second surface velocity, and (3) 600 SPI differential pressure. The two designs chosen for fabrication and rig test are a spiral groove and a Rayleigh step seal. Rig testing is currently underway. The Subsonic Core High Speed Air/Oil Seal Program is developing shaft-to-ground seals for next-generation propulsion systems that will minimize leakage and provide full life. Significantly higher rotor speeds and temperatures will be experienced. Technologies being exploited include, hydrodynamic lift assist features, ultra light weight designs, and improved cooling schemes. Parametric testing has been completed; a final seal design is entering the endurance test phase.

Grala, Carl P.↗

NASA NTRS2014

Aerostructural Level Set Topology Optimization for a Common Research Model Wing

The purpose of this work is to use level set topology optimization to improve the design of a representative wing box structure for the NASA common research model. The objective is to minimize the total compliance of the structure under aerodynamic and body force loading, where the aerodynamic loading is coupled to the structural deformation. A taxi bump case was also considered, where only body force loads were applied. The trim condition that aerodynamic lift must balance the total weight of the aircraft is enforced by allowing the root angle of attack to change. The level set optimization method is implemented on an unstructured three-dimensional grid, so that the method can optimize a wing box with arbitrary geometry. Fast matching and upwind schemes are developed for an unstructured grid, which make the level set method robust and efficient. The adjoint method is used to obtain the coupled shape sensitivities required to perform aerostructural optimization of the wing box structure.

Dunning, Peter D.↗

NASA NTRS2011

Compact, Controlled Resistance Exercise Device

Spaceflight leads to muscle and bone atrophy. Isoinertial (free-weight) exercises provide a sufficient stimulus to elicit increases in both muscle strength and bone mineral density in Earth-based studies. While exercise equipment is in use on the International Space Station for crewmember health maintenance, current devices are too large to place in a transport vehicle or small spacecraft. Therefore, a portable computer controlled resistance exercise device is being developed that is able to simulate the inertial loading experienced when lifting a mass on Earth. This portable device weighs less than 50 lb and can simulate the resistance of lifting and lowering up to 600 lb of free-weights. The objective is to allow crewmembers to perform resistance exercise with loads capable of maintaining muscle and bone health. The device is reconfigurable and allows for the performance of typical Earth-based free-weight exercises. Forces exerted, volume of work, range of motion, time-under-tension, and speed/ acceleration of movement are recorded and can be remotely monitored to track progress and modify individual protocols based on exercise session data. A performance evaluation will be completed and data will be presented that include ground-reaction force comparisons between the device and free-weight dead-lifts over a spectrum of resistance levels. Movement biomechanics will also be presented.

Paulus, David C.↗