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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 415 records · Page 23

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NASA NTRS1992

Propulsion system performance resulting from an Integrated Flight/Propulsion Control design

Propulsion system specific results are presented from the application of the Integrated Methodology for Propulsion and Airframe Control (IMPAC) design approach to Integrated Flight/Propulsion Control design for a STOVL aircraft in transition flight. The IMPAC method is briefly discussed and the propulsion system specifications for the integrated control design are examined. The structure of a linear engine controller that results from partitioning a linear centralized controller is discussed. The details of a nonlinear propulsion control system are presented, including a scheme to protect the engine operational limits: the fan surge margin and the acceleration/deceleration schedule which limits the fuel flow. Also, a simple but effective multivariable integrator windup protection scheme is investigated. Nonlinear closed-loop simulation results are presented for two typical pilot commands for transition flight: acceleration while maintaining flight path angle and a change in flight path angle while maintaining airspeed. The simulation nonlinearities include the airframe/engine coupling, the actuator and sensor dynamics and limits, the protection scheme for the engine operational limits, and the integrator windup protection. Satisfactory performance of the total airframe plus engine system for transition flight, as defined by the specifications, is maintained during the limit operation of the closed-loop engine subsystem.

Mattern, Duane↗

NASA NTRS1992

Propulsion system performance resulting from an integrated flight/propulsion control design

Propulsion-system-specific results are presented from the application of the integrated methodology for propulsion and airframe control (IMPAC) design approach to integrated flight/propulsion control design for a 'short takeoff and vertical landing' (STOVL) aircraft in transition flight. The IMPAC method is briefly discussed and the propulsion system specifications for the integrated control design are examined. The structure of a linear engine controller that results from partitioning a linear centralized controller is discussed. The details of a nonlinear propulsion control system are presented, including a scheme to protect the engine operational limits: the fan surge margin and the acceleration/deceleration schedule that limits the fuel flow. Also, a simple but effective multivariable integrator windup protection scheme is examined. Nonlinear closed-loop simulation results are presented for two typical pilot commands for transition flight: acceleration while maintaining flightpath angle and a change in flightpath angle while maintaining airspeed. The simulation nonlinearities include the airframe/engine coupling, the actuator and sensor dynamics and limits, the protection scheme for the engine operational limits, and the integrator windup protection. Satisfactory performance of the total airframe plus engine system for transition flight, as defined by the specifications, was maintained during the limit operation of the closed-loop engine subsystem.

Mattern, Duane↗

NASA NTRS1980

Near term hybrid passenger vehicle development program, phase 1

Missions for hybrid vehicles that promise to yield high petroleum impact were identified and a preliminary design, was developed that satisfies the mission requirements and performance specifications. Technologies that are critical to successful vehicle design, development and fabrication were determined. Trade-off studies to maximize fuel savings were used to develop initial design specifications of the near term hybrid vehicle. Various designs were "driven" through detailed computer simulations which calculate the petroleum consumption in standard driving cycles, the petroleum and electricity consumptions over the specified missions, and the vehicle's life cycle costs over a 10 year vehicle lifetime. Particular attention was given to the selection of the electric motor, heat engine, drivetrain, battery pack and control system. The preliminary design reflects a modified current compact car powered by a currently available turbocharged diesel engine and a 24 kW (peak) compound dc electric motor.

Source record↗

NASA NTRS2018

MMS Extended Mission Design: Evaluation of a Lunar Gravity Assist Option

This paper will describe a study that was carried out on the design of a set of maneuvers that were considered for the later stages of an extended mission of the Magnetospheric Multiscale (MMS) mission. The goal of these maneuvers was to put MMS into a significantly different orbit from those flown heretofore, so allowing science collection in a different region of the magnetosphere. This study was made feasible by the fact that the rate at which fuel is being consumed to maintain small formations on the MMS high-apogee orbit is less than expected pre-flight: the current consumption rate is only about 2 kg/yr/spacecraft. In addition, the spacecraft finished the prime mission with a significant amount of fuel remaining: this was about 1-sigma above the mean when compared with pre-launch Monte Carlo simulations. The resulting situation is similar to that of a libration orbit mission, where station-keeping requires so little fuel that any margin at all will lead to an extensive mission lifetime. In the case of MMS, the spacecraft could, if desired, perform formation flying in the current orbit for several decades. Alternatively, the spacecraft could use a significant fraction of the remaining fuel to perform major orbit modifications, while still leaving enough to conduct formation flying for on the order of a decade. The extended mission maneuvers studied here are further apogee-raises, with the goal of setting up one or more lunar gravity assists. Geometry dictates that a lunar encounter is only achievable when the MMS apogee vector lies approximately in the lunar orbit plane: this limits the possible dates to mid-2021 or early 2027.

Design↗

NASA NTRS1973

Effect of fuel vapor concentrations on combustor emissions and performance

Effects of fuel vaporization on the exhaust emission levels of oxides of nitrogen, carbon monoxide, total hydrocarbons, and smoke number were obtained in an experimental turbojet combustor segment. Two different fuel injectors were used in which liquid ASTM A-1 jet fuel and vapor propane fuel were independently controlled to simulate varying degrees of vaporization. Tests were conducted over a range of inlet-air temperatures from 478 to 700 K, pressures from 4 to 20 atm, and combustor reference velocities from 15.3 to 27.4 m/sec. Converting from liquid to complete vapor fuel resulted in oxides of nitrogen reductions of as much as 22 percent and smoke number reductions up to 51 percent. Supplement data are also presented on flame emissivity, flame temperature, and primary-zone liner wall temperatures.

Norgren, C. T.↗

NASA NTRS2011

Design and Evaluation of the Terminal Area Precision Scheduling and Spacing System

This paper describes the design, development and results from a high fidelity human-in-the-loop simulation of an integrated set of trajectory-based automation tools providing precision scheduling, sequencing and controller merging and spacing functions. These integrated functions are combined into a system called the Terminal Area Precision Scheduling and Spacing (TAPSS) system. It is a strategic and tactical planning tool that provides Traffic Management Coordinators, En Route and Terminal Radar Approach Control air traffic controllers the ability to efficiently optimize the arrival capacity of a demand-impacted airport while simultaneously enabling fuel-efficient descent procedures. The TAPSS system consists of four-dimensional trajectory prediction, arrival runway balancing, aircraft separation constraint-based scheduling, traffic flow visualization and trajectory-based advisories to assist controllers in efficient metering, sequencing and spacing. The TAPSS system was evaluated and compared to today's ATC operation through extensive series of human-in-the-loop simulations for arrival flows into the Los Angeles International Airport. The test conditions included the variation of aircraft demand from a baseline of today's capacity constrained periods through 5%, 10% and 20% increases. Performance data were collected for engineering and human factor analysis and compared with similar operations both with and without the TAPSS system. The engineering data indicate operations with the TAPSS show up to a 10% increase in airport throughput during capacity constrained periods while maintaining fuel-efficient aircraft descent profiles from cruise to landing.

Swenson, Harry N.↗

NASA NTRS1972

Pollution measurements of a swirl-can combustor

Pollutant levels of oxides of nitrogen, unburned hydrocarbons, and carbon monoxide were measured for an experimental, annular, swirl can combustor. The combustor was 42 inches in diameter, incorporated 120 modules, and was specifically designed for elevated exit temperature performance. Test conditions included combustor inlet temperatures of 600, 900 and 1050 F, inlet pressures of 5 to 6 atmospheres, reference velocities of 69 to 120 feet per second and fuel-air ratios of 0.014 to 0.0695. Tests were also conducted at a simulated engine idle condition. Results demonstrated that swirl can combustors produce oxides of nitrogen levels substantially lower than conventional combustor designs. These reductions are attributed to reduced dwell times resulting from short combustor length, quick mixing of combustion gases with diluent air, and to uniform fuel distributions resulting from the swirl can approach. Radial staging of fuel at idle conditions resulted in increases in combustion efficiencies and corresponding reductions in pollutant levels.

Niedzwiecki, R. W.↗

NASA NTRS1972

Pollution measurements of a swirl-can combustor.

Pollutant levels of oxides of nitrogen, unburned hydrocarbons, and carbon monoxide were measured for an experimental, annular, swirl-can combustor. The combustor was 42 inches in diameter, incorporated 120 modules, and was specifically designed for elevated exit temperature performance. Test conditions included combustor inlet temperatures of 600, 900 and 1050 F, inlet pressures of 5 to 6 atmospheres, reference velocities of 69 to 120 feet per second, and fuel-air ratios of 0.014 to 0.0695. Tests were also conducted at a simulated engine idle condition. Results demonstrated that swirl-can combustors produce oxides of nitrogen levels substantially lower than conventional combustor designs.

Niedzwiecki, R. W.↗

NASA NTRS1981

Combustion characteristics of hydrogen. Carbon monoxide based gaseous fuels

An experimental rig program was conducted with the objective of evaluating the combuston performance of a family of fuel gases based on a mixture of hydrogen and carbon monoxide. These gases, in addition to being members of a family, were also representative of those secondary fuels that could be produced from coal by various gasification schemes. In particular, simulated Winkler, Lurgi, and Blue-water low and medium energy content gases were used as fuels in the experimental combustor rig. The combustor used was originally designed as a low NOx rich-lean system for burning liquid fuels with high bound nitrogen levels. When used with the above gaseous fuels this combustor was operated in a lean-lean mode with ultra long residence times. The Blue-water gas was also operated in a rich-lean mode. The results of these tests indicate the possibility of the existence of an 'optimum' gas turbine hydrogen - carbon monoxide based secondary fuel. Such a fuel would exhibit NOx and high efficiency over the entire engine operating range. It would also have sufficient stability range to allow normal light-off and engine acceleration. Solar Turbines Incorporated would like to emphasize that the results presented here have been obtained with experimental rig combustors. The technologies generated could, however, be utilized in future commercial gas turbines.

Notardonato, J. J.↗

NASA NTRS1974

Effect of diffuser bleed on performance of an annular swirl can combustor

A full-scale annular swirl can combustor equipped with diffuser bleed capability was tested at low fuel-air ratio, burning natural gas fuel. Test results show that, by drawing off a small amount of air on the inner or outer diffuser wall, the radial profile of combustor exit temperature could be changed from a strongly tip biased, to a strongly hub biased shape. Furthermore, combustor total pressure loss could be reduced by about one third by using a total bleed rate equal to 6 percent of diffuser inlet flow. At simulated engine idle conditions combustion efficiency was increased from 64 percent to nearly 100 percent by using 3 percent outer wall bleed. Combustor blowout performance was also improved significantly by using this diffuser bleed scheme.

Juhasz, A. J.↗

NASA NTRS2015

Application of Pinniped Vibrissae to Aeropropulsion

Vibrissae of Phoca Vitulina (Harbor Seal) and Mirounga Angustirostris (Elephant Seal) possessundulations along their length. Harbor Seal Vibrissae were shown to reduce vortex induced vibrations and reduce dragcompared to appropriately scaled cylinders and ellipses. Samples of Harbor Seal vibrissae, Elephant Seal vibrissae andCalifornia Sea Lion vibrissae were collected from the Marine Mammal Center in California. CT scanning, microscopy and3D scanning techniques were utilized to characterize the whiskers. Computational fluid dynamics simulations of thewhiskers were carried out to compare them to an ellipse and a cylinder. Leading edge parameters from the whiskerswere used to create a 3D profile based on a modern power turbine blade. The NASA SW-2 facility was used to performwind tunnel cascade testing on the 'Seal Blades'. Computational Fluid Dynamics simulations were used to studyincidence angles from -37 to +10 degrees on the aerodynamic performance of the Seal Blade. The tests and simulationswere conducted at a Reynolds number of 100,000. The Seal Blades showed consistent performance improvements overthe baseline configuration. It was determined that a fuel burn reduction of approximately 5 could be achieved for a fixedwing aircraft. Noise reduction potential is also explored

vortex induced vibration↗

NASA NTRS2010

The Successful Implementation of NASA Orbital Debris Requirements for the Retirement of TDRS-1

TDRS-1 was decommissioned on October 28th 2009 following more than 26 years of operation. The Grand Old Dame of the TDRSS constellation wa s launched aboard the maiden voyage of the Space Shuttle Challenger ( STS-6) in April 1983. TDRS-1 survived a malfunction of the Inertial Upper Stage eventually utilizing its own propulsion system to success fully reach its assigned station in geosynchronous orbit. The anomalo us beginning of the TDRS-1 mission was not without lasting consequenc es as the primary reaction control system (A-side) was completely di sabled with an apparent propellant leak and the secondary system (B-s ide) suffered damage to its negative roll thruster rendering the thru ster inoperable. Following decommissioning the challenge to completin g a successful TDRS-1 end-of-mission (EOM) was the implementation of the stringent orbital debris requirements of NPR 8715.6 with a parti ally functioning spacecraft not originally designed to meet those req uirements. The TDRS-1 EOM had three key goals: 1) removal of the spac ecraft from geosynchronous orbit; 2) depletion of the remaining prope llant; and 3) passivation of all other sources of energy storage or generation. The TDRS-1 EOM approach was one of minimizing risks while accomplishing the goals above. The orbit raising portion of EOM was accomplished using deltavelocity operations already proven during pre vious stationchanging maneuvers. The propellant depletion approach wa s necessarily more aggressive as over 20 hours of burn time was requ ired to deplete the remaining fuel. A novel approach utilizing a spin ning, thrusting, passively controlled spacecraft configuration was ut ilized to achieve reasonable burn durations that met schedule constra ints. This nonstandard configuration required careful analysis of ele ctrical, thermal, and communication subsystems. The configuration wa s thoroughly simulated prior to the start of operations and carefully characterized during the initial spin period and first burn. Passiva tion was by definition a unique operation not previously performed wi thin the TDRS 1-7 constellation. Use was made of a TDRS spacecraft si mulator to verify the operational procedure to mitigate risks and pro vide crew training. TDRS-1 orbit raising maneuvers commenced on June 5th 2010 and completed on June 14th with an apogee 370 km and a peri gee 352 km above geosynchronous altitude. 127 kg of propellant were e stimated to be remaining in the tanks at the completion of orbit rais ing. TDRS-1 was placed in its spinning orientation on June 16th and the first fuel depletion burn was performed the following day. A seri es of 10 depletion burns were performed ending on June 26th when both propellant tank pressures experienced dramatic drops. Final passivat ion was performed on June 27 th deactivating electronics, removing t he batteries from the bus and solar arrays, and disabling the space-t o-ground communications equipment. The Second TDRS Ground Terminal (S TGT) continued to open loop track TDRS-1 for several days attempting command reacquisition several times a day. All attempts were unsucce ssful confirming passivation was achieved. The TDRS-1 orbit at passiv ation was an orbital debris compliant 36,319 x 36,128 km in height. W hile differing spacecraft designs may preclude mimicking the exact TD RS-1 EOM approach, the TDRS-1 campaign serves to demonstrate that pr e-NPR 8715.6 designs can be made to meet the requirements resulting i n a reduced orbit debris environment for future missions

Mirczak, Walter↗

NASA NTRSDate not supplied

Computational Assessment of Inlet Backflow Effects on Rotating Detonation Engine Performance and Operability

The performance impact of flow reversal at the inlet of an airbreathing rotating detonation engine (RDE) is investigated using 2 and 3-dimensional computational fluid dynamic (CFD) simulations. Flow reversal, or backflow, occurs in RDE inlets in the high-pressure region directly behind the rotating detonation front. This is also where most of the engine thrust or pressure gain is produced. The amount of backflow relative to throughflow depends on the inlet design. For the present work, a simple annular ‘slit’ design is used. The simulations are idealized in several ways, including that fuel and air are premixed, but prevented from reacting when within the inlet region. The results indicate that even with idealizations, the impact of inlet backflow on pressure gain can be substantial. The simulations also reveal an intriguing instability that develops in certain configurations. The mass flow rate into the RDE begins to oscillate at a regular frequency that is substantially less than the detonation rotational frequency. This is accompanied by oscillations in the detonation height. The oscillation amplitude grows over time until the detonation ultimately fails. Both the performance and instability results emphasize the need for carefully designed RDE inlets that provide low loss when flow is in the forward direction, but high resistance when the flow is reversed. Development of such high-diodicity inlets is critical to achieving pressure gain in airbreathing RDE’s.

Detonation↗

NASA NTRSDate not supplied

Computational Assessment of Inlet Backflow Effects on Rotating Detonation Engine Performance and Operability

detonation↗

NASA NTRS1998

Control System and Flexible Satellite Interaction During Orbit Transfer Maneuver

In this paper the interaction between the attitude control system and the flexible structure of an artificial satellite during orbit transfer maneuver has been investigated. The satellite was modeled by a rigid central body with one or more flexible appendages. The dynamics equations were obtained by Lagrangean approach. The flexible appendages were treated as clamped-free beam and its displacement was discretized by assumed- mode method. In order to transfer the satellite, a typical Hohmann transfer and a burn-coast-burn strategy were used and the attitude was controlled by an on-off controller. During transfer procedure a global analysis of satellite has been done, such as: performance of control system, influence of elastic response in control system, thruster firing frequency, fuel consumption and variation of orbital elements. In order to avoid the interaction with structure motion, a control system with bandwidth of one decade bellow the fundamental frequency was used. In the simulations the firing frequency was evaluated in an approximately way but kept below the fundamental frequency of the structure. The control system has kept the attitude below the specifications. As a result, the orbit transfer maneuvering has been done correctly without excessive excitation of flexible appendage.

Adenilson Roberto da Silva↗

NASA NTRS1995

Review of the PDWA Concept for Combustion Enhancement in a Supersonic Air-Breathing Combustor Environment

This paper reviews the design of the Pulsed Detonation Wave Augmentor (PDWA) concept and the preliminary computational fluid dynamics studies that supported it. The PDWA relies on the rapid generation of detonation waves in a small tube, which are then injected into the supersonic stream of the main combustor. The blast waves thus generated are used to stimulate the mixing and combustion inside the main combustor. The mixing enhancement relies on various forms of the baroclinic interaction, where misaligned pressure and density gradients combine to produce vortical flow. By using unsteady shock waves, the concept also uses the Richtmyer-Meshkov effect to further increase the rate of mixing. By carefully designing the respective configurations of the combustor and the detonation tubes, one can also increase the penetration of the fuel into the supersonic air stream. The unsteady shocks produce lower stagnation pressure losses than steady shocks. Combustion enhancement can also be obtained through the transient shock-heating of the fuel-air interface, and the lowering of the ignition delay in these regions. The numerical simulations identify these processes, and show which configurations give the best results. Engineering considerations are also presented, and discuss the feasibility of the concept. Of primary importance are the enhancements in performance, the design simplicity, the minimization of the power, cost, and weight, and the methods to achieve very rapid cycling.

Canbier, Jean-Luc↗

NASA NTRS1991

Ascent performance of an air-breathing horizontal-takeoff launch vehicle

Simulations are conducted to investigate a proposed NASA launch vehicle that is fully reusable, takes off horizontally, and uses airbreathing propulsion in a single stage. The propulsion model is based on a cycle analysis method, and the vehicle is assumed to be a rigid structure with distributed fuel, operating under a range of atmospheric conditions. The program to optimize simulated trajectories (POST) is modified to include a predictor-corrector guidance capability and then used to generate the trajectories. Significant errors are encountered during the unpowered coast phase due to uncertainty in the atmospheric density profile. The amount of ascent propellant needed is shown to be directly related to the thrust-vector angle and the location of the center of gravity of the vehicle because of the importance of aim-drag losses to total ideal velocity.

Powell, Richard W.↗

NASA NTRS1981

Modeling of the aircraft in-trail-following task during profile descent

The cockpit display of traffic information (CDTI) system concepts enable the pilot to observe the surrounding air traffic pattern. The impact of such a system is far reaching in terms of improved safety, pilot and controller workload, and aircraft fuel efficiency. One direct payoff is the ability to distribute the ATC workload to the pilot in such tasks as merging and spacing. The CDTI application of spacing approach aircraft in the terminal area is addressed. In-trail-following/CDTI experiments were performed using realistic cockpit simulators and profile descent approach scenarios. Based on collected experimental simulator data, pilot models were developed which include state estimation, decision making and flight control aspects. These models were coupled with models of aircraft and CDTI equipment to study the dynamic phenomena and stability of strings of aircraft along various approach patterns.

Goka, T.↗