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A method for predicting transient thrust augmenting ejector characteristics is introduced. The analysis blends classic self-similar turbulent jet descriptions with a mixing region control volume analysis to predict transient effects in a new way. Details of the theoretical foundation, the solution algorithm, and sample calculations are given.
The initial results of a test program conducted on an ejector-augmented-lift STOVL aircraft design in the NASA-Ames 40-by-80-Foot Wind Tunnel has established the efficient performance of this full-scale aircraft realization of the propulsion-system configuration. It has also been noted that the generation of larger lift-augmentation ratios in this form than in component tests. In addition, it is found that configurations possessing thrust/weight ratios and wing loadings representative of current practice could, when outfitted with the present propulsion scheme, produce adequate level-flight acceleration through transition from VTOL hover to wingborne flight.
A method for predicting transient thrust augmenting ejector characteristics is introduced. The analysis blends classic self-similar turbulent jet descriptions with a mixing region control volume analysis to predict transient effects in a new way. Details of the theoretical foundation, the solution algorithm, and sample calculations are given.
Venturi air-jet vacuum ejector pump light in weight, requires no electrical power, does not contribute heat to aircraft, and provides high pumping speeds at moderate suctions. High-pressure motive gas required for this type of pump bled from compressor of aircraft engine with negligible effect on performance of engine. Used as source of vacuum for differential-absorption CO-measurement (DACOM), modified to achieve in situ measurements of CO at frequency response of 10 Hz. Provides improvement in spatial resolution and potentially leads to capability to measure turbulent flux of CO by use of eddy-correlation technique.
New method of analysis of transient flow in thrust-augmenting ejector based on control-volume formulation of governing equations. Considered as potential elements of propulsion subsystems of short-takeoff/vertical-landing airplanes.
A time-iterative full Navier-Stokes code, PARC, is used to analyze the flowfield of a two-dimensional ejector nozzle system. A parametric study was performed for two controlling parameters, duct to nozzle area ratio and nozzle pressure ratio. Results show that there is an optimum area ratio for the efficient pumping of secondary flow. At high area ratios, a freestream flow passes directly through the mixing duct without giving adequate pumping. At low area ratios, the jet boundary blocks the incoming flow. The nozzle pressure ratio variation shows that the pumping rate increases as the pressure ratio increases, provided there is no interaction between the shroud wall and the shock cell structure.
A time-iterative full Navier-Stokes code, PARC, is used to analyze the flowfield of a two-dimensional ejector nozzle system. A parametric study was performed for two controlling parameters, duct to nozzle area ratio and nozzle pressure ratio. Results show that there is an optimum area ratio for the efficient pumping of secondary flow. At high area ratios, a freestream flow passes directly through the mixing duct without giving adequate pumping. At low area ratios, the jet boundary blocks the incoming flow. The nozzle pressure ratio variation shows that the pumping rate increases as the pressure ratio increases, provided there is no interaction between the shroud wall and the shock cell structure.
Propulsion-induced aerodynamic effects are computed by using data obtained from tests of a full-scale model of the E-7A configuration that were run in the NASA Ames 40 x 80 and 80 x 120-foot tunnels. These effects are shown both in aerodynamic coefficients and nondimensionalized by thrust versus equivalent velocity. The test results show that the full-scale ventral nozzle-induced effects are smaller than scale model data, while the ejector-induced effects demonstrate fair agreement. The effect of tunnel test section size is primarily observed at low and static speed, high thrust conditions, as seen in comparisons of data from both tunnels. The 80 x 120-foot tunnel test gave better static thrust calibrations and permitted a greater range of test freestream velocities and angles-of-attack, extending the database provided by the full-scale model.
Documentation of the installation and use of venturi air-jet vacuum ejectors for high-volume atmospheric sampling on aircraft platforms is presented. Information on the types of venturis that are useful for meeting the pumping requirements of atmospheric-sampling experiments is also presented. A description of the configuration and installation of the venturi system vacuum line is included with details on the modifications that were made to adapt a venturi to the NASA Electra aircraft at GSFC, Wallops Flight Facility. Flight test results are given for several venturis with emphasis on applications to the Differential Absorption Carbon Monoxide Measurement (DACOM) system at LaRC. This is a source document for atmospheric scientists interested in using the venturi systems installed on the NASA Electra or adapting the technology to other aircraft.
The factors affecting the performance of ejector suitable for aircraft cooling are investigated theoretically and experimentally. The investigation covers a range of shroud-to-nozzle diameter ratios from 1.1 to 1.6, of shroud lengths from 0.2 to 2.28 nozzle diameters, of secondary-to-primary weight-flow ratios from 0 to 0.12, and of ambient-to-nozzle pressure ratios from 0.7 to 0.06. The results of a simplified theoretical analysis based on each type of flow are in good agreement with those experimentally obtained.
The results indicated increases in auxiliary-inlet pressure recovery with increases in scoop height relative to the boundary-layer thickness. The pressure recovery increased at about the same rate as theoretically predicted for an inlet in a boundary layer having a one-seventh power profile, but was only about 0.68 to 0.75 of the theoretically obtainable values. Under some operating conditions, flow from the primary jet was exhausted through the auxiliary inlet. This phenomenon could be predicted from the ejector pumping characteristics.
An investigation was made at free-stream Mach numbers of 1.62, 1.93, and 2.41 to determine the effects of a primary jet and secondary air flow on the base pressure and pressures acting over the boattailsurface of a body of revolution for two secondary discharge areas. The Mach numbers of the primary nozzles were 1 and 3.23 with the secondary mass flow being varied from 0 to 10 percent of the primary mass flow. The ratio of jet stagnation temperature to tunnel stagnation temperature was about 0.96. The Reynolds number range of the investigation was from 2.1 x 10(6) to 2.9 x 10(6)based on body length. All testing was conducted with a turbulent boundary layer along the model. This report presents results obtained with zero-length ejector and covers jet static-pressure ratios from the jet-off condition to a maximum of about 128 for the sonic nozzle and to a maximum of about 9 for the supersonic nozzle.
A simple lumped parameter based inverse design tool has been developed which provides flow path geometry and entrainment estimates subject to operational, acoustic, and design constraints. These constraints are manifested through specification of primary mass flow rate or ejector thrust, fully-mixed exit velocity, and static pressure matching. Fundamentally, integral forms of the conservation equations coupled with the specified design constraints are combined to yield an easily invertible linear system in terms of the flow path cross-sectional areas. Entrainment is computed by back substitution. Initial comparison with experimental and analogous one-dimensional methods show good agreement. Thus, this simple inverse design code provides an analytically based, preliminary design tool with direct application to High Speed Civil Transport (HSCT) design studies.
A piloted motion simulator evaluation, using the NASA Ames Vertical Motion Simulator, was conducted in support of a NASA Lewis Contractual study of the integration of flight and propulsion systems of a STOVL aircraft. Objectives of the study were to validate the Design Methods for Integrated Control Systems (DMICS) concept, to evaluate the handling qualities, and to assess control power usage. The E-7D ejector-augmentor STOVL fighter design served as the basis for the simulation. Handling-qualities ratings were obtained during precision hover and shipboard landing tasks. Handling-qualities ratings for these tasks ranged from satisfactory to adequate. Further improvement of the design process to fully validate the DMICS concept appears to be warranted.
The Supercharged Ejector Ramjet (SERJ) engine developments of the 1960s, as pursued by The Marquardt Corporation and its associated industry team members, are described. In just three years, engineering work on this combined-cycle powerplant type evolved, from its initial NASA-sponsored reusable space transportation system study status, into a U.S. Air Force/Navy-supported exploratory development program as a candidate 4.5 high-performance military aircraft engine. Bridging a productive transition from the spaceflight to the aviation arena, this case history supports the expectation that fully-integrated airbreathing/rocket propulsion systems hold high promise toward meeting the demanding propulsion requirements of tomorrow's aircraft-like Spaceliner class transportation systems. Lessons to be learned from this "SERJ Story" are offered for consideration by today's advanced space transportation and combined-cycle propulsion researchers and forward-planning communities.
This experimental and theoretical work involved reduction of supersonic jet noise using Mach Wave Elimination (MWE), a method that suppresses noise by means of a gaseous layer that envelops the supersonic jet. Also explored was a new method for mixing enhancement in which an axial, secondary flow enhances mixing in a primary flow. The research is relevant to the advent of future supersonic transports that must adhere to the same take-off and landing restrictions as ordinary subsonic aircraft. To reduce noise, one needs to understand the fundamental fluid mechanics of the jet, namely its turbulent structure and mean-flow characteristics, and to perform high-quality noise measurements. The results generated are applicable to free jets as well as to jets within ejectors.
Contents include the folloving: 1. Motivation. Support NASA's 3d generation launch vehicle technology program. RBCC is promising candidate for 3d generation propulsion system. 2. Approach. Focus on ejector mode p3erformance (Mach 0-3). Perform testing on established flowpath geometry. Use conventional propulsion measurement techniques. Use advanced optical diagnostic techniques to measure local combustion gas properties. 3. Objectives. Gain physical understanding of detailing mixing and combustion phenomena. Establish an experimental data set for CFD code development and validation.
Since 1986, NASA and the U.S. aerospace industry have been assessing the economic viability and environmental acceptability of a second-generation supersonic civil transport, or High Speed Civil Transport (HSCT). Environmental acceptability in terms of airport community noise and economic viability are critical elements in this endeavor. Development of a propulsion system that satisfies strict airport noise regulations (FAR36 Stage III levels), at acceptable performance and weight, is critical to the success of any HSCT program. Two-dimensional mixer-ejector (2DME) exhaust systems are one approach in achieving this goal. In support of HSCT development, GEAE (GE Aircraft Engines), under contract to the NASA Glenn Research Center, conducted this test program at the NASA Langley 16 ft transonic wind tunnel to evaluate the cold aerodynamic performance aspects of the 2DME exhaust system concept. The effects of SAR (SAR, suppressor area ratio, = mixed-flow area/primary nozzle throat area), MAR (MAR = overall exhaust system exit/mixing-plane area), flap length, CER (suppressor chute expansion ratio), chute alignment, and free stream Mach number were investigated on a 1/11th cold aerodynamic scale model of a 2DME exhaust system.