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At least 469 records · Page 26

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NASA NTRS1990

Exhaust Nozzles for Propulsion Systems with Emphasis on Supersonic Cruise Aircraft

This compendium summarizes the contributions of the NASA-Lewis and its contractors to supersonic exhaust nozzle research from 1963 to 1985. Two major research and technology efforts sponsored this nozzle research work; the U.S. Supersonic Transport (SST) Program and the follow-on Supersonic Cruise Research (SCR) Program. They account for two generations of nozzle technology: the first from 1963 to 1971, and the second from 1971 to 1985. First, the equations used to calculate nozzle thrust are introduced. Then the general types of nozzles are presented, followed by a discussion of those types proposed for supersonic aircraft. Next, the first-generation nozzles designed specifically for the Boeing SST and the second-generation nozzles designed under the SCR program are separately reviewed and then compared. A chapter on throttle-dependent afterbody drag is included, since drag has a major effect on the off-design performance of supersonic nozzles. A chapter on the performance of supersonic dash nozzles follows, since these nozzles have similar design problems, Finally, the nozzle test facilities used at NASA-Lewis during this nozzle research effort are identified and discussed. These facilities include static test stands, a transonic wind tunnel, and a flying testbed aircraft. A concluding section points to the future: a third generation of nozzles designed for a new era of high speed civil transports to produce even greater advances in performance, to meet new noise rules, and to ensure the continuity of over two decades of NASA research.

Stitt, Leonard E.↗

NASA NTRS1990

Integrated controls and health monitoring for chemical transfer propulsion

NASA is reviewing various propulsion technologies for exploring space. The requirements are examined for one enabling propulsion technology: Integrated Controls and Health Monitoring (ICHM) for Chemical Transfer Propulsion (CTP). Functional requirements for a CTP-ICHM system are proposed from tentative mission scenarios, vehicle configurations, CTP specifications, and technical feasibility. These CTP-ICHM requirements go beyond traditional reliable operation and emergency shutoff control to include: (1) enhanced mission flexibility; (2) continuously variable throttling; (3) tank-head start control; (4) automated prestart and post-shutoff engine check; (5) monitoring of space exposure degradation; and (6) product evolution flexibility. Technology development plans are also discussed.

Millis, Marc G.↗

NASA NTRS1990

Design of a device to remove lunar dust from space suits for the proposed lunar base

The National Aeronautics and Space Administration plans to begin construction of a lunar base soon after the turn of the century. During the Apollo missions, lunar dust proved to be a problem because the dust adhered to all exposed material surfaces. Since lunar dust will be a problem during the establishment and operation of this base, the need exists for a device to remove the dust from space suits before the astronauts enter clean environments. The physical properties of lunar dust were characterized and energy methods for removing the dust were identified. Eight alternate designs were developed to remove the dust. The final design uses a brush and gas jet to remove the dust. The brush bristles are made from Kevlar fibers and the gas jet uses pressurized carbon dioxide from a portable tank. A throttling valve allows variable gas flow. Also, the tank is insulated with Kapton and electrically heated to prevent condensation of the carbon dioxide when the tank is exposed to the cold (- 240 F) lunar night.

Harrington, David↗

NASA NTRS1990

A simple dynamic engine model for use in a real-time aircraft simulation with thrust vectoring

A simple dynamic engine model was developed for use in thrust vectoring control law development and real-time aircraft simulation. Engine dynamics were simulated using a throttle rate limiter and low-pass filter. This paper includes a description of a method to account for axial thrust loss resulting from thrust vectoring and the development of the simple dynamic engine model and its incorporation into the F-18 high alpha research vehicle (HARV) thrust vectoring simulation. The simple dynamic engine model was evaluated at Mach 0.2, 35,000-ft altitude and at Mach 0.7, 35,000-ft altitude. The simple dynamic engine model is within 3 percent of the steady state response, and within 25 percent of the transient response of the complete nonlinear dynamic engine model.

Johnson, Steven A.↗

NASA NTRS1990

Transient analysis of chilldown in a cryogenic transfer line

A numerical model was developed, with the SINDA'85/FLUINT program, for calculating the thermal and hydrodynamic transients that occur during the chilldown of a cryogenic transfer line, using a well documented test case to validate the modeling process. Using this model, a total of ten cases were analyzed to evaluate the effects of variable inlet valve position, inlet pressures, and the use of an internal flow liner to promote nucleate boiling. It was found that an efficient transfer line cooldown can be achieved if the inlet flow is throttled, to reduce the flow rate and quality, and an internal flow liner such as Teflon is used.

Martin, T.↗

NASA NTRS1990

Integrated controls and health monitoring for chemical transfer propulsion

Millis, Marc G.↗

NASA NTRS1990

Flight control design considerations for STOVL powered-lift flight

Short Takeoff Vertical Landing (STOVL) aircraft rely on the propulsion system for the lift and control functions during slow speed flight. The propulsion system provides the entire lifting force and all of the control power for hovering flight at zero airspeed. STOVL designs such as the General Dynamics E-7D ejector configuration incorporate an integrated flight/propulsion control system to manage the aerodynamic and propulsive-lift control effectors and to reduce the pilot's workload for powered-lift flight. Desired flying qualities characteristics are implemented through the utilization of an explicit model following flight control system. With the model following control system, the pilot commands the desired response (e.g., throttle commands vertical velocity in hover, instead of power lever angle). Design considerations for developing a multivariable model-following flight control system are presented in this paper. When the regulator gains are defined in terms of generalized controls, the design problem becomes how to best transform the generalized controls to aerodynamic control surface, thrust and thrust vectoring commands.

Vincent, James H.↗

NASA NTRS1990

Test plan pressure fed thrust chamber technology

Aerojet is developing the technology for the design of a reliable, low cost, efficient, and lightweight LOX/RP-1 pressure fed engine. This technology program is a direct result of Aerojet's liquid rocket booster (LRB) study and previous NASA studies that identified liquid engines using high bulk density hydrocarbon fuels as very attractive for a space transportation system (STS). Previous large thrust LOX/RP-1 engine development programs were characterized by costly development problems due to combustion instability damage. The combustion stability solution was typically obtained through trial and error methods of minimizing instability damage by degrading engine performance. The approach to this program was to utilize existing and newly developed combustion analysis models and design methodology to create a thrust chamber design with features having the potential of producing reliable and efficient operation. This process resulted in an engine design with a unique high thrust-per-element OFO triplet injector utilizing a low cost modular approach. Cost efficient ablative materials are baselined for the injector face and chamber. Technology demonstration will be accomplished through a hot fire test program using appropriately sized subscale hardware. This subscale testing will provide a data base to supplement the current industry data bank and to anchor and validate the applied analysis models and design methodology. Once anchored and validated, these analysis models and design methodology can be applied with greatly increased confidence to design and characterize a large scale pressure fed LOX/RP-1 thrust chamber. The objective of this test program is to generate a data base that can be used to anchor and validate existing analysis models and design methodologies and to provide early concept demonstration of a low cost, efficient LOX/RP-1 thrust chamber. Test conditions and hardware instrumentation were defined to provide data sufficient to characterize combustion stability, performance, and thermal operation over a wide thrust chamber throttling range.

Dunn, Glenn↗

NASA NTRS1990

A simple dynamic engine model for use in a real-time aircraft simulation with thrust vectoring

A simple dynamic engine model was developed at the NASA Ames Research Center, Dryden Flight Research Facility, for use in thrust vectoring control law development and real-time aircraft simulation. The simple dynamic engine model of the F404-GE-400 engine (General Electric, Lynn, Massachusetts) operates within the aircraft simulator. It was developed using tabular data generated from a complete nonlinear dynamic engine model supplied by the manufacturer. Engine dynamics were simulated using a throttle rate limiter and low-pass filter. Included is a description of a method to account for axial thrust loss resulting from thrust vectoring. In addition, the development of the simple dynamic engine model and its incorporation into the F-18 high alpha research vehicle (HARV) thrust vectoring simulation. The simple dynamic engine model was evaluated at Mach 0.2, 35,000 ft altitude and at Mach 0.7, 35,000 ft altitude. The simple dynamic engine model is within 3 percent of the steady state response, and within 25 percent of the transient response of the complete nonlinear dynamic engine model.

Johnson, Steven A.↗

NASA NTRS1991

Large liquid rocket engine transient performance simulation system

A simulation system, ROCETS, was designed and developed to allow cost-effective computer predictions of liquid rocket engine transient performance. The system allows a user to generate a simulation of any rocket engine configuration using component modules stored in a library through high-level input commands. The system library currently contains 24 component modules, 57 sub-modules and maps, and 33 system routines and utilities. FORTRAN models from other sources can be operated in the system upon inclusion of interface information on comment cards. Operation of the simulation is simplified for the user by run, execution, and output processors. The simulation system makes available steady-state trim balance, transient operation, and linear partial generation. The system utilizes a modern equation solver for efficient operation of the simulations. Transient integration methods include integral and differential forms for the trapezoidal, first order Gear, and second order Gear corrector equations. A detailed technology test bed engine (TTBE) model was generated to be used as the acceptance test of the simulation system. The general level of model detail was that reflected in the Space Shuttle Main Engine DTM. The model successfully obtained steady-state balance in main stage operation and simulated throttle transients, including engine starts and shutdown. A NASA FORTRAN control model was obtained, ROCETS interface installed in comment cards, and operated with the TTBE model in closed-loop transient mode.

Mason, J. R.↗

NASA NTRS1990

STV engine requirements

The Space Exploration Initiative program and the Space Transfer Vehicle studies provide the opportunity to evaluate the engine requirements for the next generation space engine within the context of a credible but technologically demanding program. Some of these requirements are generic, basically independent of the architecture, and others are mission and configuration dependent. Generic requirements include high reliability, space-basing, man-rataing, long life and reusability, and extended health monitoring, including automatic pre-mission checkout. Mission and configuration dependent requirements include vehicle/engine interfaces, engine throttling, performance specifications, engine system architecture, and operational requirements. It is argued that engine technology requirements must be extended from the usual performance-related issues to encompass reliability, space-basing, life, and extended health monitoring.

Redus, J. R.↗

NASA NTRS1990

Digital angular position sensor using wavelength division multiplexing

Future aircraft will use fly-by-light control systems with fiber-linked optical sensors for such measurands as temperature, pressure, and linear and angular position. A digital optical sensor is described which was developed to transmit the angular position of such slowly rotating parts as a throttle of fuel flow control valve on an aircraft. The sensor employs a reflective code plate with ten channels providing a resolution of 0.35 degrees. Two light-emitting diodes with overlapping spectra are used as light sources. A single microoptic multiplexer-demultiplexer composed of a GRIN rod lens and a miniature grating is used to disperse the spectrum and recombine the spectral components from each channel after reflection by the code plate. The results of preliminary environmental tests of this unit are discussed. The sensor has been operated for brief periods of time between -60 C and +125 without adverse effects. Preliminary vibration tests indicate that the unit will work properly at the maximum vibration levels expected in a jet-engine environment.

Fritsch, Klaus↗

NASA NTRS1991

Ascent performance issues of a vertical-takeoff rocket launch vehicle

Advanced manned launch systems studies under way at the NASA Langley Research Center are part of a broader effort that is examining options for the next manned space transportation system to be developed by the United States. One promising concept that uses near-term technologies is a fully reusable, two-stage vertical-takeoff rocket vehicle. This vehicle features parallel thrusting of the booster and orbiter with the booster cross-feeding the propellant to the orbiter until staging. In addition, after staging, the booster glides back unpowered to the launch site. This study concentrated on two issues that could affect the ascent performance of this vehicle. The first is the large gimbal angle range required for pitch trim until staging because of the propellant cross-feed. Results from this analysis show that if control is provided by gimballing of the rocket engines, they must gimbal greater than 20 deg, which is excessive when compared with current vehicles. However, this analysis also showed that this limit could be reduced to 10 deg if gimballing were augmented by throttling the booster engines. The second issue is the potential influence of off-nominal atmospheric conditions (density and winds) on the ascent performance. This study showed that a robust guidance algorithm could be developed that would insure accurate insertion, without prelaunch atmospheric knowledge.

Powell, Richard W.↗

NASA NTRS1991

Performance and stability of a booster class LOX/hydrogen swirl coaxial-element injector

Initial tests were conducted with an axisymmetric subscale version of the Advanced Launch System (ALS) prototype injector, with a pattern of pressure-atomizing LOX-swirled injector elements flowing about 50 percent more propellant per element than the Space Shuttle Main Engine injector element. The swirl coax combustion was statistically stable and quiet with and without combustion stability aids. Artificial perturbations to assess dynamic stability generated overpressures from 2 to 15 percent of chamber pressure, and all combustion oscillations were damped within 3 millisec. Chug-free throttle was demonstrated to 65 percent of the nominal operating chamber pressure. Combustion performance in an ablative-lined chamber was calculated with both specific impulse and characteristic exhaust velocity, and averaged about 97 percent. Combustion performance of the injector element depended upon the momentum angle of the injected propellants rather than the shearing rate of the fuel on the oxidizer.

Hulka, J.↗

NASA NTRS1991

A dual-cooled hydrogen-oxygen rocket engine heat transfer analysis

The potential benefits of simultaneously using hydrogen and oxygen as rocket engine coolants are described. A plug-and-spool rocket engine was examined at heat fluxes ranging from 9290 to 163,500 kW/sq m, using a combined 3D conduction/advection analysis. Both counterflow and parallel flow cooling arrangements were analyzed. The results indicate that a significant amount of heat transfer to the oxygen occurs, reducing both the hot-side wall temperature of the rocket engine and also reducing the exit temperature of the hydrogen coolant. The total heat transferred to the oxygen was found to be largely independent of the oxygen coolant flow direction. The reduction in combustion chamber wall temperatures at throttled conditions is especially desirable since the analysis indicates that double temperature maxima, one at the throat and another in the combustion chamber, occur with a traditional hydrogen-only cooled engine. A dual-cooled engine eliminates any concern for overheating in the combustion chamber.

Kacynski, Kenneth J.↗

NASA NTRS1991

Range optimization for a supersonic aircraft

Range optimal trajectories for an aircraft flying in the vertical plane are obtained from Pontryagin's Minimum Principle. Control variables are load factor n which appears nonlinearly in the equations of motion and throttle setting eta, which appears only linearly. Both controls are subject to fixed bounds, namely eta between values of 0 and 1 and absolute value of n not greater than n(max). Additionally, a dynamic pressure limit is imposed, which represents a first-order state-inequality constraint. For fixed flight time, fixed initial coordinates, and partially fixed final coordinates, the effect of the load factor limit absolute value of n not greater than n(max) is studied. Upon varying n(max), six different switching structures are obtained. All trajectories involve singular control along arcs with active dynamic pressure limit.

Seywald, Hans↗

NASA NTRS1991

Design of an Advanced Expander Test Bed

The Advanced Expander Test Bed (AETB) is the key element for development of technology for future space engines. The AETB will be used to validate the high pressure expander cycle concept, investigate system interactions and conduct investigations of advanced mission focused components and new health monitoring techniques. The AETB will use oxygen/hydrogen propellants and a split expander cycle with nominal operation at a combustion chamber pressure of 1200 psia, a mixture ratio of 6.0, and an equivalent vacuum thrust of 20,000 lbf. It will function over a wide range of conditions including throttling to 5 percent thrust, operation at a mixture ratio of 12.0, and operation in tank head idle and pumped idle modes.

Masters, Arthur I.↗

NASA NTRS1992

Control integration concept for hypersonic cruise-turn maneuvers

Piloting difficulties associated with conducting aircraft maneuvers in hypersonic flight are caused in part by the nonintuitive nature of the aircraft response and the stringent constraints anticipated on allowable angle of attack and dynamic pressure variations. An approach is documented that provides precise, coordinated maneuver control during excursions from a hypersonic cruise flight path and the necessary flight condition constraints. The approach is to achieve specified guidance commands by resolving altitude and cross range errors into a load factor and bank angle command by using a coordinate transformation that acts as an interface between outer and inner loop flight controls. This interface, referred to as a 'resolver', applies constraints on angle of attack and dynamic pressure perturbations while prioritizing altitude regulation over cross range. An unpiloted test simulation, in which the resolver was used to drive inner loop flight controls, produced time histories of responses to guidance commands and atmospheric disturbances at Mach numbers of 6, 10, 15, and 20. Angle of attack and throttle perturbation constraints, combined with high speed flight effects and the desire to maintain constant dynamic pressure, significantly impact the maneuver envelope for a hypersonic vehicle.

Raney, David L.↗