Shape memory alloy actuation for a variable area fan nozzle

NASA Astrophysics Data System (ADS)

Rey, Nancy; Tillman, Gregory; Miller, Robin M.; Wynosky, Thomas; Larkin, Michael J.; Flamm, Jeffrey D.; Bangert, Linda S.

2001-06-01

The ability to control fan nozzle exit area is an enabling technology for next generation high-bypass-ratio turbofan engines. Performance benefits for such designs are estimated at up to 9% in thrust specific fuel consumption (TSFC) relative to current fixed-geometry engines. Conventionally actuated variable area fan nozzle (VAN) concepts tend to be heavy and complicated, with significant aircraft integration, reliability and packaging issues. The goal of this effort was to eliminate these undesirable features and formulate a design that meets or exceeds leakage, durability, reliability, maintenance and manufacturing cost goals. A Shape Memory Alloy (SMA) bundled cable actuator acting to move an array of flaps around the fan nozzle annulus is a concept that meets these requirements. The SMA bundled cable actuator developed by the United Technologies Corporation (Patents Pending) provides significant work output (greater than 2200 in-lb per flap, through the range of motion) in a compact package and minimizes system complexity. Results of a detailed design study indicate substantial engine performance, weight, and range benefits. The SMA- based actuation system is roughly two times lighter than a conventional mechanical system, with significant aircraft direct operating cost savings (2-3%) and range improvements (5-6%) relative to a fixed-geometry nozzle geared turbofan. A full-scale sector model of this VAN system was built and then tested at the Jet Exit Test Facility at NASA Langley to demonstrate the system's ability to achieve 20% area variation of the nozzle under full scale aerodynamic loads. The actuator exceeded requirements, achieving repeated actuation against full-scale loads representative of typical cruise as well as greater than worst-case (ultimate) aerodynamic conditions. Based on these encouraging results, work is continuing with the goal of a flight test on a C-17 transport aircraft.

Experimental aerodynamic and acoustic model testing of the Variable Cycle Engine (VCE) testbed coannular exhaust nozzle system

NASA Technical Reports Server (NTRS)

Nelson, D. P.; Morris, P. M.

1980-01-01

Aerodynamic performance and jet noise characteristics of a one sixth scale model of the variable cycle engine testbed exhaust system were obtained in a series of static tests over a range of simulated engine operating conditions. Model acoustic data were acquired. Data were compared to predictions of coannular model nozzle performance. The model, tested with an without a hardwall ejector, had a total flow area equivalent to a 0.127 meter (5 inch) diameter conical nozzle with a 0.65 fan to primary nozzle area ratio and a 0.82 fan nozzle radius ratio. Fan stream temperatures and velocities were varied from 422 K to 1089 K (760 R to 1960 R) and 434 to 755 meters per second (1423 to 2477 feet per second). Primary stream properties were varied from 589 to 1089 K (1060 R to 1960 R) and 353 to 600 meters per second (1158 to 1968 feet per second). Exhaust plume velocity surveys were conducted at one operating condition with and without the ejector installed. Thirty aerodynamic performance data points were obtained with an unheated air supply. Fan nozzle pressure ratio was varied from 1.8 to 3.2 at a constant primary pressure ratio of 1.6; primary pressure ratio was varied from 1.4 to 2.4 while holding fan pressure ratio constant at 2.4. Operation with the ejector increased nozzle thrust coefficient 0.2 to 0.4 percent.

Performance characteristics of a variable-area vane nozzle for vectoring an ASTOVL exhaust jet up to 45 deg

NASA Technical Reports Server (NTRS)

Mcardle, Jack G.; Esker, Barbara S.

1993-01-01

Many conceptual designs for advanced short-takeoff, vertical landing (ASTOVL) aircraft need exhaust nozzles that can vector the jet to provide forces and moments for controlling the aircraft's movement or attitude in flight near the ground. A type of nozzle that can both vector the jet and vary the jet flow area is called a vane nozzle. Basically, the nozzle consists of parallel, spaced-apart flow passages formed by pairs of vanes (vanesets) that can be rotated on axes perpendicular to the flow. Two important features of this type of nozzle are the abilities to vector the jet rearward up to 45 degrees and to produce less harsh pressure and velocity footprints during vertical landing than does an equivalent single jet. A one-third-scale model of a generic vane nozzle was tested with unheated air at the NASA Lewis Research Center's Powered Lift Facility. The model had three parallel flow passages. Each passage was formed by a vaneset consisting of a long and a short vane. The longer vanes controlled the jet vector angle, and the shorter controlled the flow area. Nozzle performance for three nominal flow areas (basic and plus or minus 21 percent of basic area), each at nominal jet vector angles from -20 deg (forward of vertical) to +45 deg (rearward of vertical) are presented. The tests were made with the nozzle mounted on a model tailpipe with a blind flange on the end to simulate a closed cruise nozzle, at tailpipe-to-ambient pressure ratios from 1.8 to 4.0. Also included are jet wake data, single-vaneset vector performance for long/short and equal-length vane designs, and pumping capability. The pumping capability arises from the subambient pressure developed in the cavities between the vanesets, which could be used to aspirate flow from a source such as the engine compartment. Some of the performance characteristics are compared with characteristics of a single-jet nozzle previously reported.

Free-jet acoustic investigation of high-radius-ratio coannular plug nozzles. Comprehensive data report, volume 1

NASA Technical Reports Server (NTRS)

Knott, P. R.; Janardan, B. A.; Majjigi, R. K.; Shutiani, P. K.; Vogt, P. G.

1981-01-01

Six coannular plug nozzle configurations having inverted velocity and temperature profiles, and a baseline convergent conical nozzle were tested for simulated flight acoustic evaluation in General Electric's Anechoic Free-Jet Acoustic Facility. The nozzles were tested over a range of test conditions that are typical of a Variable Cycle Engine for application to advanced high speed aircraft. The outer stream radius ratio for most of the configurations was 0.853, and the inner-stream-outer-stream area ratio was tested in the range of 0.54. Other variables investigated were the influence of bypass struts, a simple noncontoured convergent-divergent outer stream nozzle for forward quadrant shock noise control, and the effects of varying outer stream radius and inner-stream-to-outer-stream velocity ratios on the flight noise signatures of the nozzles. It was found that in simulated flight, the high-radius-ratio coannular plug nozzles maintain their jet noise and shock noise reduction features previously observed in static testing. The presence of nozzle bypass structs will not significantly effect the acoustic noise reduction features of a General Electric-type nozzle design. A unique coannular plug nozzle flight acoustic spectral prediction method was identified and found to predict the measured results quite well. Special laser velocimeter and acoustic measurements were performed which have given new insight into the jet and shock noise reduction mechanisms of coannular plug nozzles with regard to identifying further beneficial research efforts.

Altitude Performance Characteristics of Turbojet-engine Tail-pipe Burner with Variable-area Exhaust Nozzle Using Several Fuel Systems and Flame Holders

NASA Technical Reports Server (NTRS)

Johnson, Lavern A; Meyer, Carl L

1950-01-01

A tail-pipe burner with a variable-area exhaust nozzle was investigated. From five configurations a fuel-distribution system and a flame holder were selected. The best configuration was investigated over a range of altitudes and flight Mach numbers. For the best configuration, an increase in altitude lowered the augmented thrust ratio, exhaust-gas total temperature, and tail-pipe combustion efficiency, and raised the specific fuel consumption. An increase in flight Mach number raised the augmented thrust ratio but had no apparent effect on exhaust-gas total temperature, tail-pipe combustion efficiency, or specific fuel consumption.

Experimental Determination of Linear Dynamics of Two-Spool Turbojet Engines

NASA Technical Reports Server (NTRS)

Novik, David; Heppler, Herbert

1954-01-01

Transfer functions descriptive of the response of most engine variables were determined from transient data that were obtained from approximate step inputs in fuel flow and in exhaust-nozzle area. The speed responses of both spools to fuel flow and to turbine-inlet temperature appeared as identical first-order lags. Response to exhaust-nozzle area was characterized by a first-order lag response of the outer-spool speed, accompanied by virtually no change in inner-spool speed.

Gen 2.0 Mixer/Ejector Nozzle Test at LSAF June 1995 to July 1996

NASA Technical Reports Server (NTRS)

Arney, L. D.; Sandquist, D. L.; Forsyth, D. W.; Lidstone, G. L.; Long-Davis, Mary Jo (Technical Monitor)

2005-01-01

Testing of the HSCT Generation 2.0 nozzle model hardware was conducted at the Boeing Low Speed Aeroacoustic Facility, LSAF. Concurrent measurements of noise and thrust were made at critical takeoff design conditions for a variety of mixer/ejector model hardware. Design variables such as suppressor area ratio, mixer area ratio, liner type and thickness, ejector length, lobe penetration, and mixer chute shape were tested. Parallel testing was conducted at G.E.'s Cell 41 acoustic free jet facility to augment the LSAF test. The results from the Gen 2.0 testing are being used to help shape the current nozzle baseline configuration and guide the efforts in the upcoming Generation 2.5 and 3.0 nozzle tests. The Gen 2.0 results have been included in the total airplane system studies conducted at MDC and Boeing to provide updated noise and thrust performance estimates.

Analysis and testing of high entrainment single nozzle jet pumps with variable mixing tubes

NASA Technical Reports Server (NTRS)

Hickman, K. E.; Hill, P. G.; Gilbert, G. B.

1972-01-01

An analytical model was developed to predict the performance characteristics of axisymmetric single-nozzle jet pumps with variable area mixing tubes. The primary flow may be subsonic or supersonic. The computer program uses integral techniques to calculate the velocity profiles and the wall static pressures that result from the mixing of the supersonic primary jet and the subsonic secondary flow. An experimental program was conducted to measure mixing tube wall static pressure variations, velocity profiles, and temperature profiles in a variable area mixing tube with a supersonic primary jet. Static pressure variations were measured at four different secondary flow rates. These test results were used to evaluate the analytical model. The analytical results compared well to the experimental data. Therefore, the analysis is believed to be ready for use to relate jet pump performance characteristics to mixing tube design.

Computer code for single-point thermodynamic analysis of hydrogen/oxygen expander-cycle rocket engines

NASA Technical Reports Server (NTRS)

Glassman, Arthur J.; Jones, Scott M.

1991-01-01

This analysis and this computer code apply to full, split, and dual expander cycles. Heat regeneration from the turbine exhaust to the pump exhaust is allowed. The combustion process is modeled as one of chemical equilibrium in an infinite-area or a finite-area combustor. Gas composition in the nozzle may be either equilibrium or frozen during expansion. This report, which serves as a users guide for the computer code, describes the system, the analysis methodology, and the program input and output. Sample calculations are included to show effects of key variables such as nozzle area ratio and oxidizer-to-fuel mass ratio.

Flight effects on the aerodynamic and acoustic characteristics of inverted profile coannular nozzles

NASA Technical Reports Server (NTRS)

Kozlowski, H.; Packman, A. B.

1978-01-01

The effect of forward flight on the jet noise of coannular exhaust nozzles, suitable for Variable Stream Control Engines (VSCE), was investigated in a series of wind tunnel tests. The primary stream properties were maintained constant at 300 mps and 394 K. A total of 230 acoustic data points was obtained. Force measurement tests using an unheated air supply covered the same range of tunnel speeds and nozzle pressure ratios on each of the nozzle configurations. A total of 80 points was taken. The coannular nozzle OASPL and PNL noise reductions observed statically relative to synthesized values were basically retained under simulated flight conditions. The effect of fan to primary stream area ratio on flight effects was minor. At take-off speed, the peak jet noise for a VSCE was estimated to be over 6 PNdB lower than the static noise level. High static thrust coefficients were obtained for the basic coannular nozzles, with a decay of 0.75 percent at take-off speeds.

Stepped nozzle

DOEpatents

Sutton, George P.

1998-01-01

An insert which allows a supersonic nozzle of a rocket propulsion system to operate at two or more different nozzle area ratios. This provides an improved vehicle flight performance or increased payload. The insert has significant advantages over existing devices for increasing nozzle area ratios. The insert is temporarily fastened by a simple retaining mechanism to the aft end of the diverging segment of the nozzle and provides for a multi-step variation of nozzle area ratio. When mounted in place, the insert provides the nozzle with a low nozzle area ratio. During flight, the retaining mechanism is released and the insert ejected thereby providing a high nozzle area ratio in the diverging nozzle segment.

Noise characteristics of upper surface blown configurations. Experimental program and results

NASA Technical Reports Server (NTRS)

Brown, W. H.; Searle, N.; Blakney, D. F.; Pennock, A. P.; Gibson, J. S.

1977-01-01

An experimental data base was developed from the model upper surface blowing (USB) propulsive lift system hardware. While the emphasis was on far field noise data, a considerable amount of relevant flow field data were also obtained. The data were derived from experiments in four different facilities resulting in: (1) small scale static flow field data; (2) small scale static noise data; (3) small scale simulated forward speed noise and load data; and (4) limited larger-scale static noise flow field and load data. All of the small scale tests used the same USB flap parts. Operational and geometrical variables covered in the test program included jet velocity, nozzle shape, nozzle area, nozzle impingement angle, nozzle vertical and horizontal location, flap length, flap deflection angle, and flap radius of curvature.

Stepped nozzle

DOEpatents

Sutton, G.P.

1998-07-14

An insert is described which allows a supersonic nozzle of a rocket propulsion system to operate at two or more different nozzle area ratios. This provides an improved vehicle flight performance or increased payload. The insert has significant advantages over existing devices for increasing nozzle area ratios. The insert is temporarily fastened by a simple retaining mechanism to the aft end of the diverging segment of the nozzle and provides for a multi-step variation of nozzle area ratio. When mounted in place, the insert provides the nozzle with a low nozzle area ratio. During flight, the retaining mechanism is released and the insert ejected thereby providing a high nozzle area ratio in the diverging nozzle segment. 5 figs.

Computational Fluid Dynamics Modeling of a Supersonic Nozzle and Integration into a Variable Cycle Engine Model

NASA Technical Reports Server (NTRS)

Connolly, Joseph W.; Friedlander, David; Kopasakis, George

2015-01-01

This paper covers the development of an integrated nonlinear dynamic simulation for a variable cycle turbofan engine and nozzle that can be integrated with an overall vehicle Aero-Propulso-Servo-Elastic (APSE) model. A previously developed variable cycle turbofan engine model is used for this study and is enhanced here to include variable guide vanes allowing for operation across the supersonic flight regime. The primary focus of this study is to improve the fidelity of the model's thrust response by replacing the simple choked flow equation convergent-divergent nozzle model with a MacCormack method based quasi-1D model. The dynamic response of the nozzle model using the MacCormack method is verified by comparing it against a model of the nozzle using the conservation element/solution element method. A methodology is also presented for the integration of the MacCormack nozzle model with the variable cycle engine.

Computational Fluid Dynamics Modeling of a Supersonic Nozzle and Integration into a Variable Cycle Engine Model

NASA Technical Reports Server (NTRS)

Connolly, Joseph W.; Friedlander, David; Kopasakis, George

2014-01-01

This paper covers the development of an integrated nonlinear dynamic simulation for a variable cycle turbofan engine and nozzle that can be integrated with an overall vehicle Aero-Propulso-Servo-Elastic (APSE) model. A previously developed variable cycle turbofan engine model is used for this study and is enhanced here to include variable guide vanes allowing for operation across the supersonic flight regime. The primary focus of this study is to improve the fidelity of the model's thrust response by replacing the simple choked flow equation convergent-divergent nozzle model with a MacCormack method based quasi-1D model. The dynamic response of the nozzle model using the MacCormack method is verified by comparing it against a model of the nozzle using the conservation element/solution element method. A methodology is also presented for the integration of the MacCormack nozzle model with the variable cycle engine.

Prototype Morphing Fan Nozzle Demonstrated

NASA Technical Reports Server (NTRS)

Lee, Ho-Jun; Song, Gang-Bing

2004-01-01

Ongoing research in NASA Glenn Research Center's Structural Mechanics and Dynamics Branch to develop smart materials technologies for aeropropulsion structural components has resulted in the design of the prototype morphing fan nozzle shown in the photograph. This prototype exploits the potential of smart materials to significantly improve the performance of existing aircraft engines by introducing new inherent capabilities for shape control, vibration damping, noise reduction, health monitoring, and flow manipulation. The novel design employs two different smart materials, a shape-memory alloy and magnetorheological fluids, to reduce the nozzle area by up to 30 percent. The prototype of the variable-area fan nozzle implements an overlapping spring leaf assembly to simplify the initial design and to provide ease of structural control. A single bundle of shape memory alloy wire actuators is used to reduce the nozzle geometry. The nozzle is subsequently held in the reduced-area configuration by using magnetorheological fluid brakes. This prototype uses the inherent advantages of shape memory alloys in providing large induced strains and of magnetorheological fluids in generating large resistive forces. In addition, the spring leaf design also functions as a return spring, once the magnetorheological fluid brakes are released, to help force the shape memory alloy wires to return to their original position. A computerized real-time control system uses the derivative-gain and proportional-gain algorithms to operate the system. This design represents a novel approach to the active control of high-bypass-ratio turbofan engines. Researchers have estimated that such engines will reduce thrust specific fuel consumption by 9 percent over that of fixed-geometry fan nozzles. This research was conducted under a cooperative agreement (NCC3-839) at the University of Akron.

Cooled variable nozzle radial turbine for rotor craft applications

NASA Technical Reports Server (NTRS)

Rogo, C.

1981-01-01

An advanced, small 2.27 kb/sec (5 lbs/sec), high temperature, variable area radial turbine was studied for a rotor craft application. Variable capacity cycles including single-shaft and free-turbine engine configurations were analyzed to define an optimum engine design configuration. Parametric optimizations were made on cooled and uncooled rotor configurations. A detailed structural and heat transfer analysis was conducted to provide a 4000-hour life HP turbine with material properties of the 1988 time frame. A pivoted vane and a moveable sidewall geometry were analyzed. Cooling and variable geometry penalties were included in the cycle analysis. A variable geometry free-turbine engine configuration with a design 1477K (2200 F) inlet temperature and a compressor pressure ratio of 16:1 was selected. An uncooled HP radial turbine rotor with a moveable sidewall nozzle showed the highest performance potential for a time weighted duty cycle.

Nozzle design study for a quasi-axisymmetric scramjet-powered vehicle at Mach 7.9 flight conditions

NASA Astrophysics Data System (ADS)

Tanimizu, Katsuyoshi; Mee, David J.; Stalker, Raymond J.; Jacobs, Peter A.

2013-09-01

A nozzle shape optimization study for a quasi-axisymmetric scramjet has been performed for a Mach 7.9 operating condition with hydrogen fuel, aiming at the application of a hypersonic airbreathing vehicle. In this study, the nozzle geometry which is parameterized by a set of design variables, is optimized for the single objective of maximum net thrust using an in-house CFD solver for inviscid flowfields with a simple force prediction methodology. The combustion is modelled using a simple chemical reaction code. The effects of the nozzle design on the overall vehicle performance are discussed. For the present geometry, net thrust is achieved for the optimized vehicle design. The results of the nozzle-optimization study show that performance is limited by the nozzle area ratio that can be incorporated into the vehicle without leading to too large a base diameter of the vehicle and increasing the external drag of the vehicle. This study indicates that it is very difficult to achieve positive thrust at Mach 7.9 using the basic geometry investigated.

Thrust distribution for attitude control in a variable thrust propulsion system with four ACS nozzles

NASA Astrophysics Data System (ADS)

Lim, Yeerang; Lee, Wonsuk; Bang, Hyochoong; Lee, Hosung

2017-04-01

A thrust distribution approach is proposed in this paper for a variable thrust solid propulsion system with an attitude control system (ACS) that uses a reduced number of nozzles for a three-axis attitude maneuver. Although a conventional variable thrust solid propulsion system needs six ACS nozzles, this paper proposes a thrust system with four ACS nozzles to reduce the complexity and mass of the system. The performance of the new system was analyzed with numerical simulations, and the results show that the performance of the system with four ACS nozzles was similar to the original system while the mass of the whole system was simultaneously reduced. Moreover, a feasibility analysis was performed to determine whether a thrust system with three ACS nozzles is possible.

Variable volume combustor with pre-nozzle fuel injection system

DOE Office of Scientific and Technical Information (OSTI.GOV)

Keener, Christopher Paul; Johnson, Thomas Edward; McConnaughhay, Johnie Franklin

The present application provides a combustor for use with a gas turbine engine. The combustor may include a number of fuel nozzles, a pre-nozzle fuel injection system supporting the fuel nozzles, and a linear actuator to maneuver the fuel nozzles and the pre-nozzle fuel injection system.

Free jet feasibility study of a thermal acoustic shield concept for AST/VCE application-dual flow. Comprehensive data report. Volume 1: Test nozzles and acoustic data

NASA Technical Reports Server (NTRS)

Janardan, B. A.; Brausch, J. F.; Price, A. O.

1984-01-01

Acoustic and diagnostic data that were obtained to determine the influence of selected geometric and aerodynamic flow variables of coannular nozzles with thermal acoustic shields are summarized in this comprehensive data report. A total of 136 static and simulated flight acoustic test points were conducted with 9 scale-model nozzles The tested nozzles included baseline (unshielded), 180 deg shielded, and 360 deg shielded dual flow coannular plug configurations. The baseline configurations include a high radius ratio unsuppressed coannular plug nozzle and a coanuular plug nozzle and a coannular plug nozzle with a 20-chute outer stream suppressor. The tests were conducted at nozzle temperatures and pressure typical of operating conditions of variable cycle engine.

Progress with variable cycle engines

NASA Technical Reports Server (NTRS)

Westmoreland, J. S.

1980-01-01

The evaluation of components of an advanced propulsion system for a future supersonic cruise vehicle is discussed. These components, a high performance duct burner for thrust augmentation and a low jet noise coannular exhaust nozzle, are part of the variable stream control engine. An experimental test program involving both isolated component and complete engine tests was conducted for the high performance, low emissions duct burner with excellent results. Nozzle model tests were completed which substantiate the inherent jet noise benefit associated with the unique velocity profile possible of a coannular exhaust nozzle system on a variable stream control engine. Additional nozzle model performance tests have established high thrust efficiency levels at takeoff and supersonic cruise for this nozzle system. Large scale testing of these two critical components is conducted using an F100 engine as the testbed for simulating the variable stream control engine.

Turbocharger with variable nozzle having vane sealing surfaces

DOE Office of Scientific and Technical Information (OSTI.GOV)

Arnold, Philippe; Petitjean, Dominique; Ruquart, Anthony

2011-11-15

A variable nozzle for a turbocharger includes a plurality of vanes rotatably mounted on a nozzle ring and disposed in a nozzle flow path defined between the nozzle ring and an opposite nozzle wall. Either or both of the faces of the nozzle ring and nozzle wall include(s) at least one step that defines sealing surfaces positioned to be substantially abutted by airfoil surfaces of the vanes in the closed position of the vanes and to be spaced from the airfoil surfaces in positions other than the closed position. This substantial abutment between the airfoil surfaces and the sealing surfacesmore » serves to substantially prevent exhaust gas from leaking past the ends of the airfoil portions. At the same time, clearances between the nozzle ring face and the end faces of the airfoil portions can be sufficiently large to prevent binding of the vanes under all operating conditions.« less

Computational Study of an Axisymmetric Dual Throat Fluidic Thrust Vectoring Nozzle for a Supersonic Aircraft Application

NASA Technical Reports Server (NTRS)

Deere, Karen A.; Flamm, Jeffrey D.; Berrier, Bobby L.; Johnson, Stuart K.

2007-01-01

A computational investigation of an axisymmetric Dual Throat Nozzle concept has been conducted. This fluidic thrust-vectoring nozzle was designed with a recessed cavity to enhance the throat shifting technique for improved thrust vectoring. The structured-grid, unsteady Reynolds- Averaged Navier-Stokes flow solver PAB3D was used to guide the nozzle design and analyze performance. Nozzle design variables included extent of circumferential injection, cavity divergence angle, cavity length, and cavity convergence angle. Internal nozzle performance (wind-off conditions) and thrust vector angles were computed for several configurations over a range of nozzle pressure ratios from 1.89 to 10, with the fluidic injection flow rate equal to zero and up to 4 percent of the primary flow rate. The effect of a variable expansion ratio on nozzle performance over a range of freestream Mach numbers up to 2 was investigated. Results indicated that a 60 circumferential injection was a good compromise between large thrust vector angles and efficient internal nozzle performance. A cavity divergence angle greater than 10 was detrimental to thrust vector angle. Shortening the cavity length improved internal nozzle performance with a small penalty to thrust vector angle. Contrary to expectations, a variable expansion ratio did not improve thrust efficiency at the flight conditions investigated.

Variable Geometry Aircraft Pylon Structure and Related Operation Techniques

NASA Technical Reports Server (NTRS)

Shah, Parthiv N. (Inventor)

2014-01-01

An aircraft control structure can be utilized for purposes of drag management, noise control, or aircraft flight maneuvering. The control structure includes a high pressure engine nozzle, such as a bypass nozzle or a core nozzle of a turbofan engine. The nozzle exhausts a high pressure fluid stream, which can be swirled using a deployable swirl vane architecture. The control structure also includes a variable geometry pylon configured to be coupled between the nozzle and the aircraft. The variable geometry pylon has a moveable pylon section that can be deployed into a deflected state to maintain or alter a swirling fluid stream (when the swirl vane architecture is deployed) for drag management purposes, or to assist in the performance of aircraft flight maneuvers.

VCE early acoustic test results of General Electric's high-radius ratio coannular plug nozzle

NASA Technical Reports Server (NTRS)

Knott, P. R.; Brausch, J. F.; Bhutiani, P. K.; Majjigi, R. K.; Doyle, V. L.

1980-01-01

Results of variable cycle engine (VCE) early acoustic engine and model scale tests are presented. A summary of an extensive series of far field acoustic, advanced acoustic, and exhaust plume velocity measurements with a laser velocimeter of inverted velocity and temperature profile, high radius ratio coannular plug nozzles on a YJ101 VCE static engine test vehicle are reviewed. Select model scale simulated flight acoustic measurements for an unsuppressed and a mechanical suppressed coannular plug nozzle are also discussed. The engine acoustic nozzle tests verify previous model scale noise reduction measurements. The engine measurements show 4 to 6 PNdB aft quadrant jet noise reduction and up to 7 PNdB forward quadrant shock noise reduction relative to a fully mixed conical nozzle at the same specific thrust and mixed pressure ratio. The influences of outer nozzle radius ratio, inner stream velocity ratio, and area ratio are discussed. Also, laser velocimeter measurements of mean velocity and turbulent velocity of the YJ101 engine are illustrated. Select model scale static and simulated flight acoustic measurements are shown which corroborate that coannular suppression is maintained in forward speed.

Arcjet nozzle area ratio effects

NASA Technical Reports Server (NTRS)

Curran, Francis M.; Sarmiento, Charles J.; Birkner, Bjorn W.; Kwasny, James

1990-01-01

An experimental investigation was conducted to determine the effect of nozzle area ratio on the operating characteristics and performance of a low power dc arcjet thruster. Conical thoriated tungsten nozzle inserts were tested in a modular laboratory arcjet thruster run on hydrogen/nitrogen mixtures simulating the decomposition products of hydrazine. The converging and diverging sides of the inserts had half angles of 30 and 20 degrees, respectively, similar to a flight type unit currently under development. The length of the diverging side was varied to change the area ratio. The nozzle inserts were run over a wide range of specific power. Current, voltage, mass flow rate, and thrust were monitored to provide accurate comparisons between tests. While small differences in performance were observed between the two nozzle inserts, it was determined that for each nozzle insert, arcjet performance improved with increasing nozzle area ratio to the highest area ratio tested and that the losses become very pronounced for area ratios below 50. These trends are somewhat different than those obtained in previous experimental and analytical studies of low Re number nozzles. It appears that arcjet performance can be enhanced via area ratio optimization.

Arcjet Nozzle Area Ratio Effects

NASA Technical Reports Server (NTRS)

Curran, Francis M.; Sarmiento, Charles J.; Birkner, Bjorn W.; Kwasny, James

1990-01-01

An experimental investigation was conducted to determine the effect of nozzle area ratio on the operating characteristics and performance of a low power dc arcjet thruster. Conical thoriated tungsten nozzle inserts were tested in a modular laboratory arcjet thruster run on hydrogen/nitrogen mixtures simulating the decomposition products of hydrazine. The converging and diverging sides of the inserts had half angles of 30 and 20 degrees, respectively, similar to a flight type unit currently under development. The length of the diverging side was varied to change the area ratio. The nozzle inserts were run over a wide range of specific power. Current, voltage, mass flow rate, and thrust were monitored to provide accurate comparisons between tests. While small differences in performance were observed between the two nozzle inserts, it was determined that for each nozzle insert, arcjet performance improved with increasing nozzle area ratio to the highest area ratio tested and that the losses become very pronounced for area ratios below 50. These trends are somewhat different than those obtained in previous experimental and analytical studies of low Re number nozzles. It appears that arcjet performance can be enhanced via area ratio optimization.

Static internal performance evaluation of several thrust reversing concepts for 2D-CD nozzles

NASA Technical Reports Server (NTRS)

Rowe, R. K.; Duss, D. J.; Leavitt, L. D.

1984-01-01

Recent performance testing of the two-dimensional convergent-divergent (2D-CD) nozzle has established the concept as a viable alternative to the axisymmetric nozzle for advanced technology aircraft. This type of exhaust system also offers potential integration and performance advantages in the areas of thrust reversing and vectoring over axi-symmetric nozzles. These advantages include the practical integration of thrust reversers which operate not only to reduce landing roll but also operate in-flight for enhanced maneuvering and thrust spoiling. To date there is a very limited data base available from which criteria can be developed for the design and evaluation of this type of thrust reverser system. For this reason, a static scale model test was conducted in which five different thrust reverser designs were evaluated. Each of the five models had varying performance/integration requirements which dictated the five different designs. Some of the parameters investigated in this test included; variable angle external cascade vanes, fixed angle internal cascade vanes, variable position inner doors, external slider doors and internal slider valves. In addition, normal force and yawing moment generation was investigated using the thrust reverser system. Selected results from this test will be presented and discussed in this paper.

Progress on Variable Cycle Engines

NASA Technical Reports Server (NTRS)

Westmoreland, J. S.; Howlett, R. A.; Lohmann, R. P.

1979-01-01

Progress in the development and future requirements of the Variable Stream Control Engine (VSCE) are presented. The two most critical components of this advanced system for future supersonic transports, the high performance duct burner for thrust augmentation, and the low jet coannular nozzle were studied. Nozzle model tests substantiated the jet noise benefit associated with the unique velocity profile possible with a coannular nozzle system on a VSCE. Additional nozzle model performance tests have established high thrust efficiency levels only at takeoff and supersonic cruise for this nozzle system. An experimental program involving both isolated component and complete engine tests has been conducted for the high performance, low emissions duct burner with good results and large scale testing of these two components is being conducted using a F100 engine as the testbed for simulating the VSCE. Future work includes application of computer programs for supersonic flow fields to coannular nozzle geometries, further experimental testing with the duct burner segment rig, and the use of the Variable Cycle Engine (VCE) Testbed Program for evaluating the VSCE duct burner and coannular nozzle technologies.

IR signature study of aircraft engine for variation in nozzle exit area

NASA Astrophysics Data System (ADS)

Baranwal, Nidhi; Mahulikar, Shripad P.

2016-01-01

In general, jet engines operate with choked nozzle during take-off, climb and cruise, whereas unchoking occurs while landing and taxiing (when engine is not running at full power). Appropriate thrust in an aircraft in all stages of the flight, i.e., take-off, climb, cruise, descent and landing is achieved through variation in the nozzle exit area. This paper describes the effect on thrust and IR radiance of a turbojet engine due to variation in the exit area of a just choked converging nozzle (Me = 1). The variations in the nozzle exit area result in either choking or unchoking of a just choked converging nozzle. Results for the change in nozzle exit area are analyzed in terms of thrust, mass flow rate and specific fuel consumption. The solid angle subtended (Ω) by the exhaust system is estimated analytically, for the variation in nozzle exit area (Ane), as it affects the visibility of the hot engine parts from the rear aspect. For constant design point thrust, IR radiance is studied from the boresight (ϕ = 0°, directly from the rear side) for various percentage changes in nozzle exit area (%ΔAne), in the 1.9-2.9 μm and 3-5 μm bands.

Experimental thrust performance of a high-area-ratio rocket nozzle

NASA Technical Reports Server (NTRS)

Pavli, Albert J.; Kacynski, Kenneth J.; Smith, Tamara A.

1987-01-01

An experimental investigation was conducted to determine the thrust performance attainable from high-area-ratio rocket nozzles. A modified Rao-contoured nozzle with an expansion area of 1030 was test fired with hydrogen-oxygen propellants at altitude conditions. The nozzle was also tested as a truncated nozzle, at an expansion area ratio of 428. Thrust coefficient and thrust coefficient efficiency values are presented for each configuration at various propellant mixture ratios (oxygen/fuel). Several procedural techniques were developed permitting improved measurement of nozzle performance. The more significant of these were correcting the thrust for the aneroid effects, determining the effective chamber pressure, and referencing differential pressure transducers to a vacuum reference tank.

Experimental thrust performance of a high area-ratio rocket nozzle

NASA Technical Reports Server (NTRS)

Pavli, A. J.; Kacynski, K. J.; Smith, T. A.

1986-01-01

An experimental investigation was conducted to determine the thrust performance attainable from high-area-ratio rocket nozzles. A modified Rao-contoured nozzle with an expansion area of 1030 was test fired with hydrogen-oxygen propellants at altitude conditions. The nozzle was also tested as a truncated nozzle, at an expansion area ratio of 428. Thrust coefficient and thrust coefficient efficiency values are presented for each configuration at various propellant mixture ratios (oxygen/fuel). Several procedural techniques were developed permitting improved measurement of nozzle performance. The more significant of these were correcting the thrust for the aneroid effects, determining the effective chamber pressure, and referencing differential pressure transducers to a vacuum reference tank.

Experimental performance of a high-area-ratio rocket nozzle at high combustion chamber pressure

NASA Technical Reports Server (NTRS)

Jankovsky, Robert S.; Kazaroff, John M.; Pavli, Albert J.

1996-01-01

An experimental investigation was conducted to determine the thrust coefficient of a high-area-ratio rocket nozzle at combustion chamber pressures of 12.4 to 16.5 MPa (1800 to 2400 psia). A nozzle with a modified Rao contour and an expansion area ratio of 1025:1 was tested with hydrogen and oxygen at altitude conditions. The same nozzle, truncated to an area ratio of 440:1, was also tested. Values of thrust coefficient are presented along with characteristic exhaust velocity efficiencies, nozzle wall temperatures, and overall thruster specific impulse.

Effect of Installation of Mixer/Ejector Nozzles on the Core Flow Exhaust of High-Bypass-Ratio Turbofan Engines

NASA Technical Reports Server (NTRS)

Harrington, Douglas E.

1998-01-01

The aerospace industry is currently investigating the effect of installing mixer/ejector nozzles on the core flow exhaust of high-bypass-ratio turbofan engines. This effort includes both full-scale engine tests at sea level conditions and subscale tests in static test facilities. Subscale model tests are to be conducted prior to full-scale testing. With this approach, model results can be analyzed and compared with analytical predications. Problem areas can then be identified and design changes made and verified in subscale prior to committing to any final design configurations for engine ground tests. One of the subscale model test programs for the integrated mixer/ejector development was a joint test conducted by the NASA Lewis Research Center and Pratt & Whitney Aircraft. This test was conducted to study various mixer/ejector nozzle configurations installed on the core flow exhaust of advanced, high-bypass-ratio turbofan engines for subsonic, commercial applications. The mixer/ejector concept involves the introduction of largescale, low-loss, streamwise vortices that entrain large amounts of secondary air and rapidly mix it with the primary stream. This results in increased ejector pumping relative to conventional ejectors and in more complete mixing within the ejector shroud. The latter improves thrust performance through the efficient energy exchange between the primary and secondary streams. This experimental program was completed in April 1997 in Lewis' CE-22 static test facility. Variables tested included the nozzle area ratio (A9/A8), which ranged from 1.6 to 3.0. This ratio was varied by increasing or decreasing the nozzle throat area, A8. Primary nozzles tested included both lobed mixers and conical primaries. These configurations were tested with and without an outer shroud, and the shroud position was varied by inserting spacers in it. In addition, data were acquired with and without secondary flow.

Variable volume combustor

DOEpatents

Ostebee, Heath Michael; Ziminsky, Willy Steve; Johnson, Thomas Edward; Keener, Christopher Paul

2017-01-17

The present application provides a variable volume combustor for use with a gas turbine engine. The variable volume combustor may include a liner, a number of micro-mixer fuel nozzles positioned within the liner, and a linear actuator so as to maneuver the micro-mixer fuel nozzles axially along the liner.

Exhaust-nozzle characterisitcs for a twin-jet variable-wing-sweep fighter airplane model at Mach numbers to 2.2

NASA Technical Reports Server (NTRS)

Reubush, D. E.; Mercer, C. E.

1974-01-01

A wind-tunnel investigation has been conducted to determine the exhaust-nozzle aerodynamic and propulsive characteristics for a twin-jet variable-wing-sweep fighter airplane model. The powered model was tested in the Langley 16-foot transonic tunnel and in the Langley 4-foot supersonic pressure tunnel at Mach numbers to 2.2 and at angles of attack from about minus 2 to 6 deg. Compressed air was used to simulate the nozzle exhaust flow at values of jet total-pressure ratio from approximately 1 (jet off) to about 21. Effects of configuration variables such as speed-brake deflection, store installation, and boundary-layer thickness on the the nozzle characteristics were also investigated.

Exhaust Nozzles for Supersonic Flight with Turbojet Engines

NASA Technical Reports Server (NTRS)

Shillito, Thomas B.; Hearth, Donald P.; Cortright, Edgar M.

1956-01-01

Good internal performance over a wide range of flight conditions can be obtained with either a plug nozzle or a variable ejector nozzle that can provide a divergent shroud at high pressure ratios. For both the ejector and the plug nozzle, external flow can sometimes cause serious drag losses and, for some plug-nozzle installations, external flow can cause serious internal performance losses. Plug-nozzle cooling and design of the secondary-air-flow systems for ejectors were also considered .

Jet Nozzle Having Centerbody for Enhanced Exit Area Mixing

NASA Technical Reports Server (NTRS)

Seiner, John M. (Inventor); Gilinsky, Mikhail M. (Inventor)

1999-01-01

A nozzle arrangement includes a nozzle and a centerbody. The longitudinal axis of the centerbody is coaxially aligned with the nozzle. The centerbody has a free end portion shaped to create vortices in exhaust exiting the exit area. The vortices enhance mixing action in the exhaust and reduce exhaust noise while augmenting thrust.

Acoustic and aerodynamic performance investigation of inverted velocity profile coannular plug nozzles. [variable cycle engines

NASA Technical Reports Server (NTRS)

Knott, P. R.; Blozy, J. T.; Staid, P. S.

1981-01-01

The results of model scale parametric static and wind tunnel aerodynamic performance tests on unsuppressed coannular plug nozzle configurations with inverted velocity profile are discussed. The nozzle configurations are high-radius-ratio coannular plug nozzles applicable to dual-stream exhaust systems typical of a variable cycle engine for Advanced Supersonic Transport application. In all, seven acoustic models and eight aerodynamic performance models were tested. The nozzle geometric variables included outer stream radius ratio, inner stream to outer stream ratio, and inner stream plug shape. When compared to a conical nozzle at the same specific thrust, the results of the static acoustic tests with the coannular nozzles showed noise reductions of up to 7 PNdB. Extensive data analysis showed that the overall acoustic results can be well correlated using the mixed stream velocity and the mixed stream density. Results also showed that suppression levels are geometry and flow regulation dependent with the outer stream radius ratio, inner stream-to-outer stream velocity ratio and inner stream velocity ratio and inner stream plug shape, as the primary suppression parameters. In addition, high-radius ratio coannular plug nozzles were found to yield shock associated noise level reductions relative to a conical nozzle. The wind tunnel aerodynamic tests showed that static and simulated flight thrust coefficient at typical takeoff conditions are quite good - up to 0.98 at static conditions and 0.974 at a takeoff Mach number of 0.36. At low inner stream flow conditions significant thrust loss was observed. Using an inner stream conical plug resulted in 1% to 2% higher performance levels than nozzle geometries using a bent inner plug.

Sliding vane geometry turbines

DOEpatents

Sun, Harold Huimin; Zhang, Jizhong; Hu, Liangjun; Hanna, Dave R

2014-12-30

Various systems and methods are described for a variable geometry turbine. In one example, a turbine nozzle comprises a central axis and a nozzle vane. The nozzle vane includes a stationary vane and a sliding vane. The sliding vane is positioned to slide in a direction substantially tangent to an inner circumference of the turbine nozzle and in contact with the stationary vane.

Static internal performance of single-expansion-ramp nozzles with various combinations of internal geometric parameters

NASA Technical Reports Server (NTRS)

Re, R. J.; Leavitt, L. D.

1984-01-01

The effects of five geometric design parameters on the internal performance of single-expansion-ramp nozzles were investigated at nozzle pressure ratios up to 10 in the static-test facility of the Langley 16-Foot Transonic Tunnel. The geometric variables on the expansion-ramp surface of the upper flap consisted of ramp chordal angle, ramp length, and initial ramp angle. On the lower flap, the geometric variables consisted of flap angle and flap length. Both internal performance and static-pressure distributions on the centerlines of the upper and lower flaps were obtained for all 43 nozzle configurations tested.

Design and analysis report for the flight weight 20-inch Columbium secondary nozzle for the RL10 engine

NASA Technical Reports Server (NTRS)

Castro, J. H.

1989-01-01

Pratt & Whitney (P and W) is currently under contract to NASA-LeRC for a multi-year program to evaluate the feasibility of the RL10-IIB/IIC engine models and the various improvements which broaden the engine capabilities and range of applications. The features being evaluated include the operation of the RL10 engine at low thrust levels and/or high mixture ratio levels and the addition of a high area ratio (250:1) translating nozzle to the engine to increase its specific impulse while shortening the installed engine length. The translating nozzle for the RL10-IIB/IIC engine is approximately 55 inches long with an exit plane diameter of 71 inches and an inlet plane diameter of 40 inches. This report documents the design and analysis work done investigating a small subscale Columbium nozzle which could be built and tested to provide findings which then could be incorporated into the high area ratio nozzle final design for the RL10-IIB/IIC engine. This report documents the design and analysis work done investigating a small subscale Columbium nozzle which could be built and tested to provide findings which then could be incorporated into the high area ratio nozzle final design for the RL10-IIB/IIC engine. The length of the subscale nozzle is 20 in.; its exit diameter is 46 in. With the nozzle in the stowed position, an RL10A-3-3A engine system is 70 inches long (Area Ratio = 61:1); with the nozzle deployed the engine length and area ratio are increased to 90 inches and 83:1 respectively. The increase in area ratio provides a calculated increase of 7 + or - 1 second of specific impulse.

Aeroacoustic Analysis of Fan Noise Reduction With Increased Bypass Nozzle Area

NASA Technical Reports Server (NTRS)

Woodward, Richard P.; Hughes, Christopher E.; Podboy, Gary G.

2005-01-01

An advanced model turbofan was tested in the NASA Glenn 9-by 15-Foot Low Speed Wind Tunnel (9x15 LSWT) to explore far field acoustic effects of increased bypass nozzle area. This fan stage test was part of the NASA Glenn Fan Broadband Source Diagnostic Test, second entry (SDT2) which acquired aeroacoustic results over a range of test conditions. The baseline nozzle was sized to produce maximum stage performance at cruise condition. However, the wind tunnel testing is conducted near sea level condition. Therefore, in order to simulate and obtain performance at other operating conditions, two additional nozzles were designed and tested one with +5 percent increase in weight flow (+5.4 percent increase in nozzle area compared with the baseline nozzle), sized to simulate the performance at the stage design point (takeoff) condition, and the other with a +7.5 percent increase in weight flow (+10.9 percent increase in nozzle area) sized for maximum weight flow with a fixed nozzle at sea level condition. Measured acoustic benefits with increased nozzle area were very encouraging, showing overall sound power level (OAPWL) reductions of 2 or more dB while the stage thrust actually increased by 2 to 3 percent except for the most open nozzle at takeoff rotor speed where stage performance decreased. Effective perceived noise levels for a 1500 ft engine flyover and 3.35 scale factor showed a similar noise reduction of 2 or more EPNdB. Noise reductions, principally in the level of broadband noise, were observed everywhere in the far field. Laser Doppler Velocimetry measurements taken downstream of the rotor showed that the total turbulent velocity decreased with increasing nozzle flow, which may explain the reduced rotor broadband noise levels.

Operating Characteristics of the Multiple Critical Venturi System and Secondary Calibration Nozzles Used for Weight-Flow Measurements in the Langley 16-Foot Transonic Tunnel

NASA Technical Reports Server (NTRS)

Berrier, B. L.; Leavitt, L. D.; Bangert, L. S.

1985-01-01

An investigation has been conducted in the Langley 16 Foot Transonic Tunnel to determine the weight flow measurement characteristics of a multiple critical Venturi system and the nozzle discharge coefficient characteristics of a series of convergent calibration nozzles. The effects on model discharge coefficient of nozzle throat area, model choke plate open area, nozzle pressure ratio, jet total temperature, and number and combination of operating Venturis were investigated. Tests were conducted at static conditions (tunnel wind off) at nozzle pressure ratios from 1.3 to 7.0.

Numerical investigation of the variable nozzle effect on the mixed flow turbine performance characteristics

NASA Astrophysics Data System (ADS)

Meziri, B.; Hamel, M.; Hireche, O.; Hamidou, K.

2016-09-01

There are various matching ways between turbocharger and engine, the variable nozzle turbine is the most significant method. The turbine design must be economic with high efficiency and large capacity over a wide range of operational conditions. These design intents are used in order to decrease thermal load and improve thermal efficiency of the engine. This paper presents an original design method of a variable nozzle vane for mixed flow turbines developed from previous experimental and numerical studies. The new device is evaluated with a numerical simulation over a wide range of rotational speeds, pressure ratios, and different vane angles. The compressible turbulent steady flow is solved using the ANSYS CFX software. The numerical results agree well with experimental data in the nozzleless configuration. In the variable nozzle case, the results show that the turbine performance characteristics are well accepted in different open positions and improved significantly in low speed regime and at low pressure ratio.

High-Area-Ratio Rocket Nozzle at High Combustion Chamber Pressure: Experimental and Analytical Validation

NASA Technical Reports Server (NTRS)

Jankovsky, Robert S.; Smith, Timothy D.; Pavli, Albert J.

1999-01-01

Experimental data were obtained on an optimally contoured nozzle with an area ratio of 1025:1 and on a truncated version of this nozzle with an area ratio of 440:1. The nozzles were tested with gaseous hydrogen and liquid oxygen propellants at combustion chamber pressures of 1800 to 2400 psia and mixture ratios of 3.89 to 6.15. This report compares the experimental performance, heat transfer, and boundary layer total pressure measurements with theoretical predictions of the current Joint Army, Navy, NASA, Air Force (JANNAF) developed methodology. This methodology makes use of the Two-Dimensional Kinetics (TDK) nozzle performance code. Comparisons of the TDK-predicted performance to experimentally attained thrust performance indicated that both the vacuum thrust coefficient and the vacuum specific impulse values were approximately 2.0-percent higher than the turbulent prediction for the 1025:1 configurations, and approximately 0.25-percent higher than the turbulent prediction for the 440:1 configuration. Nozzle wall temperatures were measured on the outside of a thin-walled heat sink nozzle during the test fittings. Nozzle heat fluxes were calculated front the time histories of these temperatures and compared with predictions made with the TDK code. The heat flux values were overpredicted for all cases. The results range from nearly 100 percent at an area ratio of 50 to only approximately 3 percent at an area ratio of 975. Values of the integral of the heat flux as a function of nozzle surface area were also calculated. Comparisons of the experiment with analyses of the heat flux and the heat rate per axial length also show that the experimental values were lower than the predicted value. Three boundary layer rakes mounted on the nozzle exit were used for boundary layer measurements. This arrangement allowed total pressure measurements to be obtained at 14 different distances from the nozzle wall. A comparison of boundary layer total pressure profiles and analytical predictions show good agreement for the first 0.5 in. from the nozzle wall; but the further into the core flow that measurements were taken, the more that TDK overpredicted the boundary layer thickness.

An electronic flow control system for a variable-rate tree sprayer

USDA-ARS?s Scientific Manuscript database

Precise modulation of nozzle flow rates is a critical measure to achieve variable-rate spray applications. An electronic flow rate control system accommodating with microprocessors and pulse width modulation (PWM) controlled solenoid valves was designed to manipulate the output of spray nozzles inde...

Nozzle Aerodynamic Stability During a Throat Shift

NASA Technical Reports Server (NTRS)

Kawecki, Edwin J.; Ribeiro, Gregg L.

2005-01-01

An experimental investigation was conducted on the internal aerodynamic stability of a family of two-dimensional (2-D) High Speed Civil Transport (HSCT) nozzle concepts. These nozzles function during takeoff as mixer-ejectors to meet acoustic requirements, and then convert to conventional high-performance convergent-divergent (CD) nozzles at cruise. The transition between takeoff mode and cruise mode results in the aerodynamic throat and the minimum cross-sectional area that controls the engine backpressure shifting location within the nozzle. The stability and steadiness of the nozzle aerodynamics during this so called throat shift process can directly affect the engine aerodynamic stability, and the mechanical design of the nozzle. The objective of the study was to determine if pressure spikes or other perturbations occurred during the throat shift process and, if so, identify the caused mechanisms for the perturbations. The two nozzle concepts modeled in the test program were the fixed chute (FC) and downstream mixer (DSM). These 2-D nozzles differ principally in that the FC has a large over-area between the forward throat and aft throat locations, while the DSM has an over-area of only about 10 percent. The conclusions were that engine mass flow and backpressure can be held constant simultaneously during nozzle throat shifts on this class of nozzles, and mode shifts can be accomplished at a constant mass flow and engine backpressure without upstream pressure perturbations.

Supersonics Project - Airport Noise Tech Challenge

NASA Technical Reports Server (NTRS)

Bridges, James

2010-01-01

The Airport Noise Tech Challenge research effort under the Supersonics Project is reviewed. While the goal of "Improved supersonic jet noise models validated on innovative nozzle concepts" remains the same, the success of the research effort has caused the thrust of the research to be modified going forward in time. The main activities from FY06-10 focused on development and validation of jet noise prediction codes. This required innovative diagnostic techniques to be developed and deployed, extensive jet noise and flow databases to be created, and computational tools to be developed and validated. Furthermore, in FY09-10 systems studies commissioned by the Supersonics Project showed that viable supersonic aircraft were within reach using variable cycle engine architectures if exhaust nozzle technology could provide 3-5dB of suppression. The Project then began to focus on integrating the technologies being developed in its Tech Challenge areas to bring about successful system designs. Consequently, the Airport Noise Tech Challenge area has shifted efforts from developing jet noise prediction codes to using them to develop low-noise nozzle concepts for integration into supersonic aircraft. The new plan of research is briefly presented by technology and timelines.

Advanced Engine Cycles Analyzed for Turbofans With Variable-Area Fan Nozzles Actuated by a Shape Memory Alloy

NASA Technical Reports Server (NTRS)

Berton, Jeffrey J.

2002-01-01

Advanced, large commercial turbofan engines using low-fan-pressure-ratio, very high bypass ratio thermodynamic cycles can offer significant fuel savings over engines currently in operation. Several technological challenges must be addressed, however, before these engines can be designed. To name a few, the high-diameter fans associated with these engines pose a significant packaging and aircraft installation challenge, and a large, heavy gearbox is often necessary to address the differences in ideal operating speeds between the fan and the low-pressure turbine. Also, the large nacelles contribute aerodynamic drag penalties and require long, heavy landing gear when mounted on conventional, low wing aircraft. Nevertheless, the reduced fuel consumption rates of these engines are a compelling economic incentive, and fans designed with low pressure ratios and low tip speeds offer attractive noise-reduction benefits. Another complication associated with low-pressure-ratio fans is their need for variable flow-path geometry. As the design fan pressure ratio is reduced below about 1.4, an operational disparity is set up in the fan between high and low flight speeds. In other words, between takeoff and cruise there is too large a swing in several key fan parameters-- such as speed, flow, and pressure--for a fan to accommodate. One solution to this problem is to make use of a variable-area fan nozzle (VAFN). However, conventional, hydraulically actuated variable nozzles have weight, cost, maintenance, and reliability issues that discourage their use with low-fan-pressure-ratio engine cycles. United Technologies Research, in cooperation with NASA, is developing a revolutionary, lightweight, and reliable shape memory alloy actuator system that can change the on-demand nozzle exit area by up to 20 percent. This "smart material" actuation technology, being studied under NASA's Ultra-Efficient Engine Technology (UEET) Program and Revolutionary Concepts in Aeronautics (RevCon) Program, has the potential to enable the next generation of efficient, quiet, very high bypass ratio turbofans. NASA Glenn Research Center's Propulsion Systems Analysis Office, along with NASA Langley Research Center's Systems Analysis Branch, conducted an independent analytical assessment of this new technology to provide strategic guidance to UEET and RevCon. A 2010-technology-level high-spool engine core was designed for this evaluation. Two families of low-spool components, one with and one without VAFN's, were designed to operate with the core. This "constant core" approach was used to hold most design parameters constant so that any performance differences between the VAFN and fixed nozzle cycles could be attributed to the VAFN technology alone. In this manner, the cycle design regimes that offer a performance payoff when VAFN's are used could be identified. The NASA analytical model of a performance-optimized VAFN turbofan with a fan pressure ratio of 1.28 is shown. Mission analyses of the engines were conducted using the notional, long-haul, advanced commercial twinjet shown. A high wing design was used to accommodate the large high-bypassratio engines. The mission fuel reduction benefit of very high bypass shape-memory-alloy VAFN aircraft was calculated to be 8.3 percent lower than a moderate bypass cycle using a conventional fixed nozzle. Shape-memory-alloy VAFN technology is currently under development in NASA's UEET and RevCon Programs.

Spray outputs from a variable-rate sprayer manipulated with PWM solenoid valves

USDA-ARS?s Scientific Manuscript database

Pressure fluctuations during variable-rate spray applications can affect nozzle flow rate fluctuations, resulting in spray outputs that do not coincide with the prescribed canopy structure volume. Variations in total flow rate discharged from 40 nozzles, each coupled with a pulse-width-modulated (PW...

Dynamic Analysis for a Geared Turbofan Engine with Variable Area Fan Nozzle

NASA Technical Reports Server (NTRS)

Csank, Jeffrey T.; Thomas, George L.

2017-01-01

Aggressive design goals have been set for future aero-propulsion systems with regards to fuel economy, noise, and emissions. To meet these challenging goals, advanced propulsion concepts are being explored and current operating margins are being re-evaluated to find additional concessions that can be made. One advanced propulsion concept being evaluated is a geared turbofan with a variable area fan nozzle (VAFN), developed by NASA. This engine features a small core, a fan driven by the low pressure turbine through a reduction gearbox, and a shape memory alloy (SMA)-actuated VAFN. The VAFN is designed to allow both a small exit area for efficient operation at cruise, while being able to open wider at high power conditions to reduce backpressure on the fan and ensure a safe level of stall margin is maintained. The VAFN is actuated via a SMA-based system instead of a conventional system to decrease overall weight of the system, however, SMA-based actuators respond relatively slowly, which introduces dynamic issues that are investigated in this work. This paper describes both a control system designed specifically for issues associated with SMAs, and dynamic analysis of the geared turbofan VAFN with the SMA actuators. Also, some future recommendations are provided for this type of propulsion system.

Model aerodynamic test results for two variable cycle engine coannular exhaust systems at simulated takeoff and cruise conditions. Comprehensive data report. Volume 2: Tabulated aerodynamic data book 2

NASA Technical Reports Server (NTRS)

Nelson, D. P.

1981-01-01

Tabulated aerodynamic data from coannular nozzle performance tests are given for test runs 26 through 37. The data include nozzle thrust coefficient parameters, nozzle discharge coefficients, and static pressure tap measurements.

Effect of several geometric parameters on the static internal performance of three nonaxisymmetric nozzle concepts

NASA Technical Reports Server (NTRS)

Berrier, B. L.; Re, R. J.

1979-01-01

Effects of several geometric parameters on the internal performance of nonaxisymmetric convergent-divergent, single-ramp expansion, and wedge nozzles were investigated at nozzle pressure ratios up to approximately 10. In addition, two different thrust-vectoring schemes were investigated with the wedge nozzle. The results indicated that as with conventional round nozzles, peak nonaxisymmetric nozzle, internal performance occurred near the nozzle pressure ratio required for fully expanded exhaust flow. Nozzle sidewall length or area generally had little effect on the internal performance of the nozzles investigated.

Unconventional nozzle tradeoff study. [space tug propulsion

NASA Technical Reports Server (NTRS)

Obrien, C. J.

1979-01-01

Plug cluster engine design, performance, weight, envelope, operational characteristics, development cost, and payload capability, were evaluated and comparisons were made with other space tug engine candidates using oxygen/hydrogen propellants. Parametric performance data were generated for existing developed or high technology thrust chambers clustered around a plug nozzle of very large diameter. The uncertainties in the performance prediction of plug cluster engines with large gaps between the modules (thrust chambers) were evaluated. The major uncertainty involves, the aerodynamics of the flow from discrete nozzles, and the lack of this flow to achieve the pressure ratio corresponding to the defined area ratio for a plug cluster. This uncertainty was reduced through a cluster design that consists of a plug contour that is formed from the cluster of high area ratio bell nozzles that have been scarfed. Light-weight, high area ratio, bell nozzles were achieved through the use of AGCarb (carbon-carbon cloth) nozzle extensions.

Experimental evaluation of a translating nozzle sidewall radial turbine

NASA Technical Reports Server (NTRS)

Roelke, Richard J.; Rogo, Casimir

1987-01-01

Studies have shown that reduced specific fuel consumption of rotorcraft engines can be achieved with a variable capacity engine. A key component in such an engine in a high-work, high-temperature variable geometry gas generator turbine. An optimization study indicated that a radial turbine with a translating nozzle sidewall could produce high efficiency over a wide range of engine flows but substantiating data were not available. An experimental program with Teledyne CAE, Toledo, Ohio was undertaken to evaluate the moving sidewall concept. A variety of translating nozzle sidewall turbine configurations were evaluated. The effects of nozzle leakage and coolant flows were also investigated. Testing was done in warm air (121 C). The results of the contractual program were summarized.

Undulated Nozzle for Enhanced Exit Area Mixing

NASA Technical Reports Server (NTRS)

Seiner, John M. (Inventor); Gilinsky, Mikhail M. (Inventor)

2000-01-01

A nozzle having an undulating surface for enhancing the mixing of a primary flow with a secondary flow or ambient air, without requiring an ejector. The nozzle includes a nozzle structure and design for introducing counter-rotating vorticity into the primary flow either through (i) internal surface corrugations where an axisymmetric line through each corrugation is coincident with an axisymmetric line through the center of the flow passageway or (ii) through one or more sets of alternating convexities and cavities in the internal surface of the nozzle where an axisymmetric line through each convexity and cavity is coincident with an axisymmetric line through the center of the flow passageway, and where the convexities contract from the entrance end towards the exit end. Exit area mixing is also enhanced by one or more chevrons attached to the exit edge of the nozzle. The nozzle is ideally suited for application as a jet engine nozzle. When used as a jet engine nozzle, noise suppression with simultaneous thrust augmentation/minimal thrust loss is achieved.

Variable volume combustor with aerodynamic fuel flanges for nozzle mounting

DOE Office of Scientific and Technical Information (OSTI.GOV)

McConnaughhay, Johnie Franklin; Keener, Christopher Paul; Johnson, Thomas Edward

2016-09-20

The present application provides a combustor for use with a gas turbine engine. The combustor may include a number of micro-mixer fuel nozzles and a fuel injection system for providing a flow of fuel to the micro-mixer fuel nozzles. The fuel injection system may include a number of support struts supporting the fuel nozzles and for providing the flow of fuel therethrough. The fuel injection system also may include a number of aerodynamic fuel flanges connecting the micro-mixer fuel nozzles and the support struts.

AeroPropulsoServoElasticity: Dynamic Modeling of the Variable Cycle Propulsion System

NASA Technical Reports Server (NTRS)

Kopasakis, George

2012-01-01

This presentation was made at the 2012 Fundamental Aeronautics Program Technical Conference and it covers research work for the Dynamic Modeling of the Variable cycle Propulsion System that was done under the Supersonics Project, in the area of AeroPropulsoServoElasticity. The presentation covers the objective for the propulsion system dynamic modeling work, followed by the work that has been done so far to model the variable Cycle Engine, modeling of the inlet, the nozzle, the modeling that has been done to model the affects of flow distortion, and finally presenting some concluding remarks and future plans.

Static internal performance of a single-engine onaxisymmetric-nozzle vaned-thrust-reverser design with thrust modulation capabilities

NASA Technical Reports Server (NTRS)

Leavitt, L. D.; Burley, J. R., II

1985-01-01

An investigation has been conducted at wind-off conditions in the stati-test facility of the Langley 16-Foot Transonic Tunnel. The tests were conducted on a single-engine reverser configuration with partial and full reverse-thrust modulation capabilities. The reverser design had four ports with equal areas. These ports were angled outboard 30 deg from the vertical impart of a splay angle to the reverse exhaust flow. This splaying of reverser flow was intended to prevent impingement of exhaust flow on empennage surfaces and to help avoid inlet reingestion of exhaust gas when the reverser is integrated into an actual airplane configuration. External vane boxes were located directly over each of the four ports to provide variation of reverser efflux angle from 140 deg to 26 deg (measured forward from the horizontal reference axis). The reverser model was tested with both a butterfly-type inner door and an internal slider door to provide area control for each individual port. In addition, main nozzle throat area and vector angle were varied to examine various methods of modulating thrust levels. Other model variables included vane box configuration (four or six vanes per box), orientation of external vane boxes with respect to internal port walls (splay angle shims), and vane box sideplates. Nozzle pressure ratio was varied from 2.0 approximately 7.0.

Static internal performance of ventral and rear nozzle concepts for short-takeoff and vertical-landing aircraft

NASA Technical Reports Server (NTRS)

Re, Richard J.; Carson, George T., Jr.

1991-01-01

The internal performance of two exhaust system concepts applicable to single-engine short-take-off and vertical-landing tactical fighter configurations was investigated. These concepts involved blocking (or partially blocking) tailpipe flow to the rear (cruise) nozzle and diverting it through an opening to a ventral nozzle exit for vertical thrust. A set of variable angle vanes at the ventral nozzle exit were used to vary ventral nozzle thrust angle between 45 and 110 deg relative to the positive axial force direction. In the vertical flight mode the rear nozzle (or tailpipe flow to it) was completely blocked. In the transition flight mode flow in the tailpipe was split between the rear and ventral nozzles and the flow was vectored at both exits for aircraft control purposes through this flight regime. In the cruise flight mode the ventral nozzle was sealed and all flow exited through the rear nozzle.

Computational Fluid Dynamics Simulation of Dual Bell Nozzle Film Cooling

NASA Technical Reports Server (NTRS)

Braman, Kalen; Garcia, Christian; Ruf, Joseph; Bui, Trong

2015-01-01

Marshall Space Flight Center (MSFC) and Armstrong Flight Research Center (AFRC) are working together to advance the technology readiness level (TRL) of the dual bell nozzle concept. Dual bell nozzles are a form of altitude compensating nozzle that consists of two connecting bell contours. At low altitude the nozzle flows fully in the first, relatively lower area ratio, nozzle. The nozzle flow separates from the wall at the inflection point which joins the two bell contours. This relatively low expansion results in higher nozzle efficiency during the low altitude portion of the launch. As ambient pressure decreases with increasing altitude, the nozzle flow will expand to fill the relatively large area ratio second nozzle. The larger area ratio of the second bell enables higher Isp during the high altitude and vacuum portions of the launch. Despite a long history of theoretical consideration and promise towards improving rocket performance, dual bell nozzles have yet to be developed for practical use and have seen only limited testing. One barrier to use of dual bell nozzles is the lack of control over the nozzle flow transition from the first bell to the second bell during operation. A method that this team is pursuing to enhance the controllability of the nozzle flow transition is manipulation of the film coolant that is injected near the inflection between the two bell contours. Computational fluid dynamics (CFD) analysis is being run to assess the degree of control over nozzle flow transition generated via manipulation of the film injection. A cold flow dual bell nozzle, without film coolant, was tested over a range of simulated altitudes in 2004 in MSFC's nozzle test facility. Both NASA centers have performed a series of simulations of that dual bell to validate their computational models. Those CFD results are compared to the experimental results within this paper. MSFC then proceeded to add film injection to the CFD grid of the dual bell nozzle. A series of nozzle pressure ratios and film coolant flow rates are investigated to determine the effect of the film injection on the nozzle flow transition behavior. The results of this CFD study of a dual bell with film injection are presented in this paper.

How to design low-noise burners

DOE Office of Scientific and Technical Information (OSTI.GOV)

Sams, G.; Jordan, J.

1996-12-01

Frequently, natural draft burner designs used in indirect heaters fail to meet the low noise standard of 85 to 88 dBA three feet from the flame arrestor. Noise encountered with indirect burner designs has been shown to be related to nozzle and firetube gas velocities. Testing shows that when the nozzle velocity is sufficiently greater than the firetube velocity, the low-frequency rumble that accompanies current designs ceases. Data obtained from field testing was used to construct a relationship between burner noise level and gas volume expansion ratio, burner air-to-fuel ratio, mixture flowrate, orifice velocity, burner area, and the number ofmore » burners. The noise from a burner can be predicted if the above easily calculable variables are known.« less

Design of a variable area diffuser for a 15-inch Mach 6 open-jet tunnel

NASA Technical Reports Server (NTRS)

Loney, Norman W.

1994-01-01

The Langley 15-inch Mach 6 High Temperature Tunnel was recently converted from a Mach 10 Hypersonic Flow Apparatus. This conversion was effected to improve the capability of testing in Mach 6 air at relatively high reservoir temperatures not previously possible at Langley. Elevated temperatures allow the matching of the Mach numbers, Reynolds numbers, and ratio of wall-to-adiabatic-wall temperatures (TW/Taw) between this and the Langley 20-inch Mach 6 CF4 Tunnel. This ratio is also matched for Langley's 31-inch Mach 10 Tunnel and is an important parameter useful in the simulation of slender bodies such as National Aerospace Plane (NASP) configurations currently being studied. Having established the nozzle's operating characteristics, the decision was made to install another test section to provide model injection capability. This test section is an open-jet type, with an injection system capable of injecting a model from retracted position to nozzle centerline between 0.5 and 2 seconds. Preliminary calibrations with the new test section resulted in Tunnel blockage. This blockage phenomenon was eliminated when the conical center body in the diffuser was replaced. The issue then, is to provide a new and more efficient variable area diffuser configuration with the capability to withstand testing of larger models without sending the Tunnel into an unstart condition. Use of the 1-dimensional steady flow equation with due regard to friction and heat transfer was employed to estimate the required area ratios (exit area / throat area) in a variable area diffuser. Correlations between diffuser exit Mach number and area ratios, relative to the stagnation pressure ratios and diffuser inlet Mach number were derived. From these correlations, one can set upper and lower operating pressures and temperatures for a given diffuser throat area. In addition, they will provide appropriate input conditions for the full 3-dimensional computational fluid dynamics (CFD) code for further simulation studies.

Uncertainty analysis for the steady-state flows in a dual throat nozzle

DOE Office of Scientific and Technical Information (OSTI.GOV)

Chen, Q.-Y.; Gottlieb, David; Hesthaven, Jan S.

2005-03-20

It is well known that the steady state of an isentropic flow in a dual-throat nozzle with equal throat areas is not unique. In particular there is a possibility that the flow contains a shock wave, whose location is determined solely by the initial condition. In this paper, we consider cases with uncertainty in this initial condition and use generalized polynomial chaos methods to study the steady-state solutions for stochastic initial conditions. Special interest is given to the statistics of the shock location. The polynomial chaos (PC) expansion modes are shown to be smooth functions of the spatial variable x,more » although each solution realization is discontinuous in the spatial variable x. When the variance of the initial condition is small, the probability density function of the shock location is computed with high accuracy. Otherwise, many terms are needed in the PC expansion to produce reasonable results due to the slow convergence of the PC expansion, caused by non-smoothness in random space.« less

Computer Graphic Design Using Auto-CAD and Plug Nozzle Research

NASA Technical Reports Server (NTRS)

Rogers, Rayna C.

2004-01-01

The purpose of creating computer generated images varies widely. They can be use for computational fluid dynamics (CFD), or as a blueprint for designing parts. The schematic that I will be working on the summer will be used to create nozzles that are a part of a larger system. At this phase in the project, the nozzles needed for the systems have been fabricated. One part of my mission is to create both three dimensional and two dimensional models on Auto-CAD 2002 of the nozzles. The research on plug nozzles will allow me to have a better understanding of how they assist in the thrust need for a missile to take off. NASA and the United States military are working together to develop a new design concept. On most missiles a convergent-divergent nozzle is used to create thrust. However, the two are looking into different concepts for the nozzle. The standard convergent-divergent nozzle forces a mixture of combustible fluids and air through a smaller area in comparison to where the combination was mixed. Once it passes through the smaller area known as A8 it comes out the end of the nozzle which is larger the first or area A9. This creates enough thrust for the mechanism whether it is an F-18 fighter jet or a missile. The A9 section of the convergent-divergent nozzle has a mechanism that controls how large A9 can be. This is needed because the pressure of the air coming out nozzle must be equal to that of the ambient pressure other wise there will be a loss of performance in the machine. The plug nozzle however does not need to have an A9 that can vary. When the air flow comes out it can automatically sense what the ambient pressure is and will adjust accordingly. The objective of this design is to create a plug nozzle that is not as complicated mechanically as it counterpart the convergent-divergent nozzle.

Static investigation of several yaw vectoring concepts on nonaxisymmetric nozzles

NASA Technical Reports Server (NTRS)

Mason, M. L.; Berrier, B. L.

1985-01-01

A test has been conducted in the static test facility of the Langley 16-Foot Transonic Tunnel to determine the flow-turning capability and the effects on nozzle internal performance of several yaw vectoring concepts. Nonaxisymmetric convergent-divergent nozzles with throat areas simulating dry and afterburning power settings and single expansion ramp nozzles with a throat area simulating a dry power setting were modified for yaw thrust vectoring. Forward-thrust and pitch-vectored nozzle configurations were tested with each yaw vectoring concept. Four basic yaw vectoring concepts were investigated on the nonaxisymmetric convergent-divergent nozzles: (1) translating sidewall; (2) downstream (of throat) flaps; (3) upstream (of throat) port/flap; and (4) powered rudder. Selected combinations of the rudder with downstream flaps or upstream port/flap were also tested. A single yaw vectoring concept, post-exit flaps, was investigated on the single expansion ramp nozzles. All testing was conducted at static (no external flow) conditions and nozzle pressure ratios varied from 2.0 up to 10.0.

Performance of high area ratio nozzles for a small rocket thruster

NASA Technical Reports Server (NTRS)

Kushida, R. O.; Hermel, J.; Apfel, S.; Zydowicz, M.

1986-01-01

Theoretical estimates of supersonic nozzle performance have been compared to experimental test data for nozzles with an area ratio of 100:1 conical and 300:1 optimum contour, and 300:1 nozzles cut off at 200:1 and 100:1. These tests were done on a Hughes Aircraft Company 5 lbf monopropellant hydrazine thruster with chamber pressures ranging from 25 to 135 psia. The analytic method used is the conventional inviscid method of characteristic with correction for laminar boundary layer displacement and drag. Replacing the 100:1 conical nozzle with the 300:1 contoured nozzle resulted in an improvement in thrust performance of 0.74 percent at chamber pressure of 25 psia to 2.14 percent at chamber pressure of 135 psia. The data is significant because it is experimental verification that conventional nozzle design techniques are applicable even where the boundary layer is laminar and displaces as much as 35 percent of the flow at the nozzle exit plane.

Numerical Investigation of Flow in an Over-Expanded Nozzle with Porous Surfaces

NASA Technical Reports Server (NTRS)

Elmiligui, Alaa; Abdol-Hamid, K. S.; Hunter, Craig A.

2005-01-01

A new porous condition has been implemented in the PAB3D solver for simulating the flow over porous surfaces. The newly-added boundary condition is utilized to compute the flow field of a non-axisymmetric, convergent-divergent nozzle incorporating porous cavities for shock-boundary layer interaction control. The nozzle has an expansion ratio (exit area/throat area) of 1.797 and a design nozzle pressure ratio of 8.78. The flow fields for a baseline nozzle (no porosity) and for a nozzle with porous surfaces (10% porosity ratio) are computed for NPR varying from 2.01 to 9.54. Computational model results indicate that the over-expanded nozzle flow was dominated by shock-induced boundary-layer separation. Porous configurations were capable of controlling off-design separation in the nozzle by encouraging stable separation of the exhaust flow. Computational simulation results, wall centerline pressure, mach contours, and thrust efficiency ratio are presented and discussed. Computed results are in excellent agreement with experimental data.

Numerical Investigation of Flow in an Over-expanded Nozzle with Porous Surfaces

NASA Technical Reports Server (NTRS)

Abdol-Hamid, Khaled S.; Elmilingui, Alaa A.; Hunter, Craig A.

2006-01-01

A new porous condition has been implemented in the PAB3D solver for simulating the flow over porous surfaces. The newly-added boundary condition is utilized to compute the flow field of a non-axisymmetric, convergent-divergent nozzle incorporating porous cavities for shock-boundary layer interaction control. The nozzle has an expansion ratio (exit area/throat area) of 1.797 and a design nozzle pressure ratio of 8.78. The flow fields for a baseline nozzle (no porosity) and for a nozzle with porous surfaces (10% porosity ratio) are computed for NPR varying from 2.01 to 9.54. Computational model results indicate that the over-expanded nozzle flow is dominated by shock-induced boundary-layer separation. Porous configurations are capable of controlling off-design separation in the nozzle by encouraging stable separation of the exhaust flow. Computational simulation results, wall centerline pressure, mach contours, and thrust efficiency ratio are presented and discussed. Computed results are in excellent agreement with experimental data.

Variable volume combustor with nested fuel manifold system

DOEpatents

McConnaughhay, Johnie Franklin; Keener, Christopher Paul; Johnson, Thomas Edward; Ostebee, Heath Michael

2016-09-13

The present application provides a combustor for use with a gas turbine engine. The combustor may include a number of micro-mixer fuel nozzles, a fuel manifold system in communication with the micro-mixer fuel nozzles to deliver a flow of fuel thereto, and a linear actuator to maneuver the micro-mixer fuel nozzles and the fuel manifold system.

Multi-Nozzle Base Flow Model in the 10- by 10-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1964-02-21

Researchers check the setup of a multi-nozzle base flow model in the 10- by 10-Foot Supersonic Wind Tunnel at the National Aeronautics and Space Administration (NASA) Lewis Research Center. NASA researchers were struggling to understand the complex flow phenomena resulting from the use of multiple rocket engines. Robert Wasko and Theodore Cover of the Advanced Development and Evaluation Division’s analysis and operations sections conducted a set of tests in the 10- by 10 tunnel to further understand the flow issues. The Lewis researchers studied four and five-nozzle configurations in the 10- by 10 at simulated altitudes from 60,000 to 200,000 feet. The nozzles were gimbaled during some of the test runs to simulate steering. The flow field for the four-nozzle clusters was surveyed in the center and the lateral areas between the nozzles, whereas the five-nozzle cluster was surveyed in the lateral area only.

Development and Assessment of Altitude Adjustable Convergent Divergent Nozzles Using Passive Flow Control

NASA Astrophysics Data System (ADS)

Mandour Eldeeb, Mohamed

The backward facing steps nozzle (BFSN) is a new developed flow adjustable exit area nozzle. It consists of two parts, the first is a base nozzle with small area ratio and the second part is a nozzle extension with surface consists of backward facing steps. The steps number and heights are carefully chosen to produce controlled flow separation at steps edges that adjust the nozzle exit area at all altitudes (pressure ratios). The BFSN performance parameters are assessed numerically in terms of thrust and side loads against the dual-bell nozzle with the same pressure ratios and cross sectional areas. Cold flow inside the planar BFSN and planar DBN are simulated using three-dimensional turbulent Navier-Stoke equations solver at different pressure ratios. The pressure distribution over the upper and the lower nozzles walls show symmetrical flow separation location inside the BFSN and an asymmetrical flow separation location inside the DBN at same vertical plane. The side loads are calculated by integrate the pressure over the nozzles walls at different pressure ratios for both nozzles. Time dependent solution for the DBN and the BFSN are obtained by solving two-dimensional turbulent flow. The side loads over the upper and lower nozzles walls are plotted against the flow time. The BFSN side loads history shows a small values of fluctuated side loads compared with the DBN which shows a high values with high fluctuations. Hot flow 3-D numerical solutions inside the axi-symmetric BFSN and DBN are obtained at different pressure ratios and compared to assess the BFSN performance against the DBN. Pressure distributions over the nozzles walls at different circumferential angels are plotted for both nozzles. The results show that the flow separation location is axi-symmetric inside the BFSN with symmetrical pressure distributions over the nozzle circumference at different pressure ratios. While the DBN results show an asymmetrical flow separation locations over the nozzle circumference at all pressure ratios.The results show that the side loads in the BFSN is 0.01%-0.6% of its value in the DBN for same pressure ratio. For further confirmation of the axi-symmetric nature of the flow in the BFSN, 2-D axi-symmetric solutions are obtained at same pressure ratios and boundary conditions. The flow parameters at the nozzle exit are calculated the 3-D and the 2-D solutions and compared to each other. The maximum difference between the 3-D and the 2-D solutions is less than 1%. Parametric studies are carried out with number of the backward facing steps varied from two to forty. The results show that as the number of backward facing steps increase, the nozzle performance in terms of thrust approach the DBN performance. The BFSN with two and six steps are simulated for pressure ratios range from 148 to 1500 and compared with the DBN and a conventional bell nozzle. Expandable BFSN study is carried out on the BFSN with two steps where the nozzle operation is divided into three modes related to the operating altitude (PR). Backward facing steps concept is applied to a full scale conventional bell nozzle by adding two backward facing steps at the end of the nozzle increasing its expansion area results in 1.8% increasing in its performance in terms of thrust coefficient at high altitudes.

Simulated flight acoustic investigation of treated ejector effectiveness on advanced mechanical suppresors for high velocity jet noise reduction

NASA Technical Reports Server (NTRS)

Brausch, J. F.; Motsinger, R. E.; Hoerst, D. J.

1986-01-01

Ten scale-model nozzles were tested in an anechoic free-jet facility to evaluate the acoustic characteristics of a mechanically suppressed inverted-velocity-profile coannular nozzle with an accoustically treated ejector system. The nozzle system used was developed from aerodynamic flow lines evolved in a previous contract, defined to incorporate the restraints imposed by the aerodynamic performance requirements of an Advanced Supersonic Technology/Variable Cycle Engine system through all its mission phases. Accoustic data of 188 test points were obtained, 87 under static and 101 under simulated flight conditions. The tests investigated variables of hardwall ejector application to a coannular nozzle with 20-chute outer annular suppressor, ejector axial positioning, treatment application to ejector and plug surfaces, and treatment design. Laser velocimeter, shadowgraph photograph, aerodynamic static pressure, and temperature measurement were acquired on select models to yield diagnositc information regarding the flow field and aerodynamic performance characteristics of the nozzles.

Calibration for Thrust and Airflow Measurements in the CE-22 Advanced Nozzle Test Facility

NASA Technical Reports Server (NTRS)

Werner, Roger A.; Wolter, John D.

2010-01-01

CE-22 facility procedures and measurements for thrust and airflow calibration obtained with choked-flow ASME nozzles are presented. Six calibration nozzles are used at an inlet total pressure from 20 to 48 psia. Throat areas are from 9.9986 to 39.986 sq. in.. Throat Reynolds number varies from 1.8 to 7.9 million. Nozzle gross thrust coefficient (CFG) uncertainty is 0.25 to 0.75 percent, with smaller uncertainly generally for larger nozzles and higher inlet total pressure. Nozzle discharge coefficient (CDN) uncertainty is 0.15 percent or less for all the data. ASME nozzle calibrations need to be done before and after research model testing to achieve these uncertainties. In addition, facility capability in terms of nozzle pressure ratio (NPR) and nozzle airflow are determined. Nozzle pressure ratio of 50 or more is obtainable at 40 psia for throat areas between 20 and 30 sq. in.. Also presented are results for two of the ASME nozzles vectored at 10deg, a dead-weight check of the vertical (perpendicular to the jet axis) force measurement, a calibration of load cell forces for the effects of facility tank deflection with tank pressure, and the calibration of the metric-break labyrinth seal.

Performance of Several Conical Convergent-Divergent Rocket-Type Exhaust Nozzles

NASA Technical Reports Server (NTRS)

Campbell, C. E.; Farley, J. M.

1960-01-01

An investigation was conducted to obtain nozzle performance data with relatively large-scale models at pressure ratios as high as 120. Conical convergent-divergent nozzles with divergence angles alpha of 15, 25, and 29 deg. were each tested at area ratios of approximately 10, 25, and 40. Heated air (1200 F) was supplied at the nozzle inlet at pressures up to 145 pounds per square inch absolute and was exhausted into quiescent air at pressures as low as 1.2 pounds per square inch absolute. Thrust ratios for all nozzle configurations are presented over the range of pressure ratios attainable and were extrapolated when possible to design pressure ratio and beyond. Design thrust ratios decreased with increasing nozzle divergence angle according to the trend predicted by the (1 + cos alpha)/2 parameter. Decreasing the nozzle divergence angle resulted in sizable increases in thrust ratio for a given surface-area ratio (nozzle weight), particularly at low nozzle pressure ratios. Correlations of the nozzle static pressure at separation and of the average static pressure downstream of separation with various nozzle parameters permitted the calculation of thrust in the separated-flow region from unseparated static-pressure distributions. Thrust ratios calculated by this method agreed with measured values within about 1 percent.

Variable volume combustor with aerodynamic support struts

DOEpatents

Ostebee, Heath Michael; Johnson, Thomas Edward; Stewart, Jason Thurman; Keener, Christopher Paul

2017-03-07

The present application provides a combustor for use with a gas turbine engine. The combustor may include a number of micro-mixer fuel nozzles and a fuel injection system for providing a flow of fuel to the micro-mixer fuel nozzles. The fuel injection system may include a number of support struts supporting the fuel nozzles and providing the flow of fuel therethrough. The support struts may include an aerodynamic contoured shape so as to distribute evenly a flow of air to the micro-mixer fuel nozzles.

Static internal performance of a single expansion ramp nozzle with multiaxis thrust vectoring capability

NASA Technical Reports Server (NTRS)

Capone, Francis J.; Schirmer, Alberto W.

1993-01-01

An investigation was conducted at static conditions in order to determine the internal performance characteristics of a multiaxis thrust vectoring single expansion ramp nozzle. Yaw vectoring was achieved by deflecting yaw flaps in the nozzle sidewall into the nozzle exhaust flow. In order to eliminate any physical interference between the variable angle yaw flap deflected into the exhaust flow and the nozzle upper ramp and lower flap which were deflected for pitch vectoring, the downstream corners of both the nozzle ramp and lower flap were cut off to allow for up to 30 deg of yaw vectoring. The effects of nozzle upper ramp and lower flap cutout, yaw flap hinge line location and hinge inclination angle, sidewall containment, geometric pitch vector angle, and geometric yaw vector angle were studied. This investigation was conducted in the static-test facility of the Langley 16-Foot Transonic Tunnel at nozzle pressure ratios up to 8.0.

A Parametric Investigation of Nozzle Planform and Internal/External Geometry at Transonic Speeds

NASA Technical Reports Server (NTRS)

Cler, Daniel L.

1995-01-01

An experimental investigation of multidisciplinary (scarfed trailing edge) nozzle divergent flap geometry was conducted at transonic speeds in the NASA Langley 16-Foot Transonic Tunnel. The geometric parameters investigated include nozzle planform, nozzle contouring location (internal and/or external), and nozzle area ratio (area ratio 1.2 and 2.0). Data were acquired over a range of Mach Numbers from 0.6 to 1.2, angle-of-attack from 0.0 degrees to 9.6 degrees and nozzle pressure ratios from 1.0 to 20.0. Results showed that increasing the rate of change internal divergence angle across the width of the nozzle or increasing internal contouring will decrease static, aeropropulsive and thrust removed drag performance regardless of the speed regime. Also, increasing the rate of change in boattail angle across the width of the nozzle or increasing external contouring will provide the lowest thrust removed drag. Scarfing of the nozzle trailing edges reduces the aeropropulsive performance for the most part and adversely affects the nozzle plume shape at higher nozzle pressure ratios thus increasing the thrust removed drag. The effects of contouring were primary in nature and the effects of planform were secondary in nature. Larger losses occur supersonically than subsonically when scarfing of nozzle trailing edges occurs. The single sawtooth nozzle almost always provided lower thrust removed drag than the double sawtooth nozzles regardless the speed regime. If internal contouring is required, the double sawtooth nozzle planform provides better static and aeropropulsive performance than the single sawtooth nozzle and if no internal contouring is required the single sawtooth provides the highest static and aeropropulsive performance.

Free jet feasibility study of a thermal acoustic shield concept for AST/VCE application: Single stream nozzles

NASA Technical Reports Server (NTRS)

Majjigi, R. K.; Brausch, J. F.; Janardan, B. A.; Balsa, T. F.; Knott, P. R.; Pickup, N.

1984-01-01

A technology base for the thermal acoustic shield concept as a noise suppression device for single stream exhaust nozzles was developed. Acoustic data for 314 test points for 9 scale model nozzle configurations were obtained. Five of these configurations employed an unsuppressed annular plug core jet and the remaining four nozzles employed a 32 chute suppressor core nozzle. Influence of simulated flight and selected geometric and aerodynamic flow variables on the acoustic behavior of the thermal acoustic shield was determined. Laser velocimeter and aerodynamic measurements were employed to yield valuable diagnostic information regarding the flow field characteristics of these nozzles. An existing theoretical aeroacoustic prediction method was modified to predict the acoustic characteristics of partial thermal acoustic shields.

Design and Analyses of High Aspect Ratio Nozzles for Distributed Propulsion Acoustic Measurements

NASA Technical Reports Server (NTRS)

Dippold, Vance F., III

2016-01-01

A series of three convergent round-to-rectangular high-aspect ratio nozzles were designed for acoustics measurements. The nozzles have exit area aspect ratios of 8:1, 12:1, and 16:1. With septa inserts, these nozzles will mimic an array of distributed propulsion system nozzles, as found on hybrid wing-body aircraft concepts. Analyses were performed for the three nozzle designs and showed that the flow through the nozzles was free of separated flow and shocks. The exit flow was mostly uniform with the exception of a pair of vortices at each span-wise end of the nozzle.

Model aerodynamic test results for two variable cycle engine coannular exhaust systems at simulated takeoff and cruise conditions. Comprehensive data report. Volume 3: Graphical data book 1

NASA Technical Reports Server (NTRS)

Nelson, D. P.

1981-01-01

A graphical presentation of the aerodynamic data acquired during coannular nozzle performance wind tunnel tests is given. The graphical data consist of plots of nozzle gross thrust coefficient, fan nozzle discharge coefficient, and primary nozzle discharge coefficient. Normalized model component static pressure distributions are presented as a function of primary total pressure, fan total pressure, and ambient static pressure for selected operating conditions. In addition, the supersonic cruise configuration data include plots of nozzle efficiency and secondary-to-fan total pressure pumping characteristics. Supersonic and subsonic cruise data are given.

Preburner of Staged Combustion Rocket Engine

NASA Technical Reports Server (NTRS)

Yost, M. C.

1978-01-01

A regeneratively cooled LOX/hydrogen staged combustion assembly system with a 400:1 expansion area ratio nozzle utilizing an 89,000 Newton (20,000 pound) thrust regeneratively cooled thrust chamber and 175:1 tubular nozzle was analyzed, assembled, and tested. The components for this assembly include two spark/torch oxygen-hydrogen igniters, two servo-controlled LOX valves, a preburner injector, a preburner combustor, a main propellant injector, a regeneratively cooled combustion chamber, a regeneratively cooled tubular nozzle with an expansion area ratio of 175:1, an uncooled heavy-wall steel nozzle with an expansion area ratio of 400:1, and interconnecting ducting. The analytical effort was performed to optimize the thermal and structural characteristics of each of the new components and the ducting, and to reverify the capabilities of the previously fabricated components. The testing effort provided a demonstration of the preburner/combustor chamber operation, chamber combustion efficiency and stability, and chamber and nozzle heat transfer.

Annular Internal-External-Expansion Rocket Nozzles for Large Booster Applications

NASA Technical Reports Server (NTRS)

Connors, James F.; Cubbison, Robert W.; Mitchell, Glenn A.

1961-01-01

For large-thrust booster applications, annular rocket nozzles employing both internal and external expansion are investigated. In these nozzles, free-stream air flows through the center as well as around the outside of the exiting jet. Flaps for deflecting the rocket exhaust are incorporated on the external-expansion surface for thrust-vector control. In order to define nozzle off-design performance, thrust vectoring effectiveness, and external stream effects, an experimental investigation was conducted on two annular nozzles with area ratios of 15 and 25 at Mach 0, 2, and 3 in the Lewis 10- by 10-foot wind tunnel. Air, pressurized to 600 pounds per square inch absolute, was used to simulate the exhaust flow. For a nozzle-pressure-ratio range of 40 to 1000, the ratio of actual to ideal thrust was essentially constant at 0.98 for both nozzles. Compared with conventional convergent-divergent configurations on hypothetical boost missions, the performance gains of the annular nozzle could yield significant orbital payload increases (possibly 8 to 17 percent). A single flap on the external-expansion surface of the area-ratio-25 annular nozzle produced a side force equal to 4 percent of the axial force with no measurable loss in axial thrust.

Free-jet acoustic investigation of high-radius-ratio coannular plug nozzles

NASA Technical Reports Server (NTRS)

Knott, P. R.; Janardan, B. A.; Majjigi, R. K.; Bhutiani, P. K.; Vogt, P. G.

1984-01-01

The experimental and analytical results of a scale model simulated flight acoustic exploratory investigation of high radius ratio coannular plug nozzles with inverted velocity and temperature profiles are summarized. Six coannular plug nozzle configurations and a baseline convergent conical nozzle were tested for simulated flight acoustic evaluation. The nozzles were tested over a range of test conditions that are typical of a Variable Cycle Engine for application to advanced high speed aircraft. It was found that in simulate flight, the high radius ratio coannular plug nozzles maintain their jet noise and shock noise reduction features previously observed in static testing. The presence of nozzle bypass struts will not significantly affect the acousticn noise reduction features of a General Electric type nozzle design. A unique coannular plug nozzle flight acoustic spectral prediction method was identified and found to predict the measured results quite well. Special laser velocimeter and acoustic measurements were performed which have given new insights into the jet and shock noise reduction mechanisms of coannular plug nozzles with regard to identifying further benificial research efforts.

Reduction of acoustic disturbances in the test section of supersonic wind tunnels by laminarizing their nozzle and test section wall boundary layers by means of suction

NASA Technical Reports Server (NTRS)

Pfenninger, W.; Syberg, J.

1974-01-01

The feasibility of quiet, suction laminarized, high Reynolds number (Re) supersonic wind tunnel nozzles was studied. According to nozzle wall boundary layer development and stability studies, relatively weak area suction can prevent amplified nozzle wall TS (Tollmien-Schlichting) boundary layer oscillations. Stronger suction is needed in and shortly upstream of the supersonic concave curvature nozzle area to avoid transition due to amplified TG (Taylor-Goertler) vortices. To control TG instability, moderately rapid and slow expansion nozzles require smaller total suction rates than rapid expansion nozzles, at the cost of larger nozzle length Re and increased TS disturbances. Test section mean flow irregularities can be minimized with suction through longitudinal or highly swept slots (swept behind local Mach cone) as well as finely perforated surfaces. Longitudinal slot suction is optimized when the suction-induced crossflow velocity increases linearly with surface distance from the slot attachment line toward the slot (through suitable slot geometry). Suction in supersonic blowdown tunnels may be operated by one or several individual vacuum spheres.

Variable volume combustor with an air bypass system

DOE Office of Scientific and Technical Information (OSTI.GOV)

Johnson, Thomas Edward; Ziminsky, Willy Steve; Ostebee, Heath Michael

The present application provides a combustor for use with flow of fuel and a flow of air in a gas turbine engine. The combustor may include a number of micro-mixer fuel nozzles positioned within a liner and an air bypass system position about the liner. The air bypass system variably allows a bypass portion of the flow of air to bypass the micro-mixer fuel nozzles.

The Prediction of Nozzle Performance and Heat Transfer in Hydrogen/Oxygen Rocket Engines with Transpiration Cooling, Film Cooling, and High Area Ratios

NASA Technical Reports Server (NTRS)

Kacynski, Kenneth J.; Hoffman, Joe D.

1994-01-01

An advanced engineering computational model has been developed to aid in the analysis of chemical rocket engines. The complete multispecies, chemically reacting and diffusing Navier-Stokes equations are modelled, including the Soret thermal diffusion and Dufour energy transfer terms. Demonstration cases are presented for a 1030:1 area ratio nozzle, a 25 lbf film-cooled nozzle, and a transpiration-cooled plug-and-spool rocket engine. The results indicate that the thrust coefficient predictions of the 1030:1 nozzle and the film-cooled nozzle are within 0.2 to 0.5 percent, respectively, of experimental measurements. Further, the model's predictions agree very well with the heat transfer measurements made in all of the nozzle test cases. It is demonstrated that thermal diffusion has a significant effect on the predicted mass fraction of hydrogen along the wall of the nozzle and was shown to represent a significant fraction of the diffusion fluxes occurring in the transpiration-cooled rocket engine.

The effect of nozzle-exit-channel shape on resultant fiber diameter in melt-electrospinning

NASA Astrophysics Data System (ADS)

Esmaeilirad, Ahmad; Ko, Junghyuk; Rukosuyev, Maxym V.; Lee, Jason K.; Lee, Patrick C.; Jun, Martin B. G.

2017-01-01

In recent decades, electrospinning using a molten poly (ε-caprolactone) resin has gained attention for creating fibrous tissue scaffolds. The topography and diameter control of such electrospun microfibers is an important issue for their different applications in tissue engineering. Charge density, initial nozzle-exit-channel cross-sectional area, nozzle to collector distance, viscosity, and processing temperature are the most important input parameters that affect the final electrospun fiber diameters. In this paper we will show that the effect of nozzle-exit-channel shape is as important as the other effective parameters in a resultant fiber diameter. However, to the best of our knowledge, the effect of nozzle-exit-channel shapes on a resultant fiber diameter have not been studied before. Comparing rectangular and circular nozzles with almost the same exit-channel cross-sectional areas in a similar processing condition showed that using a rectangular nozzle resulted in decreasing final fiber diameter up to 50%. Furthermore, the effect of processing temperature on the final fiber topography was investigated.

Variable volume combustor with center hub fuel staging

DOE Office of Scientific and Technical Information (OSTI.GOV)

Ostebee, Heath Michael; McConnaughhay, Johnie Franklin; Stewart, Jason Thurman

The present application and the resultant patent provide a combustor for use with a gas turbine engine. The combustor may include a number of micro-mixer fuel nozzles and a fuel injection system for providing a flow of fuel to the micro-mixer fuel nozzles. The fuel injection system may include a center hub for providing the flow of fuel therethrough. The center hub may include a first supply circuit for a first micro-mixer fuel nozzle and a second supply circuit for a second micro-mixer fuel nozzle.

Critical Propulsion and Noise reduction Technologies for Future Commercial Subsonic Engines. Area of Interest 14.3: Separate Flow Exhaust System Noise

NASA Technical Reports Server (NTRS)

Janardan, B. A.; Hoff, G. E.; Barter, J. W.; Brausch, J. F.; Gliebe, P. R.; Coffin, R. S.; Martens, S.; Delaney, B. R.; Dalton, W. N.; Mengle, V. G.

2000-01-01

This presentation discusses: Project Objectives, Approach and Goal; Baseline Nozzles and Test Cycle Definition; Repeatability and Baseline Nozzle Results; Noise Reduction Concepts; Noise Reduction Tests Configurations of BPR=5 Internal Plug Nozzle adn Acoustic Results; Noise Reduction Test Configurations of BPR=5 External Plug Nozzle and Acoustic Results; and Noise Reduction Tests Configurations of BPR=8 External Plug Nozzle and Acoustic Results.

Full-Authority Fault-Tolerant Electronic Engine Control Systems for Variable Cycle Engines.

DTIC Science & Technology

1981-12-01

Geometry or Fuel Flow Scheduled as a Function of Engine State, i.e. FIGV = f( N1 C2 ) Closed Loop - Geometry or Fuel Flow Modulated To Maintain an Engine...Low Pressure Turbine Inlet Area (A41) Closed Loop (Integral) N2, T22 Core Stream Exhaust Nozzle Area (AJE) Closed Loop (Integral) N1 , T2 Duct Stream...to remain at the breakpoint value while low rotor speed reference ( N1 reference) is scheduled to decrease as a function of power lever angle (PLA), to

Tables of critical-flow functions and thermodynamic properties for methane and computational procedures for both methane and natural gas

NASA Technical Reports Server (NTRS)

Johnson, R. C.

1972-01-01

Procedures for calculating the mass flow rate of methane and natural gas through nozzles are given, along with the FORTRAN 4 subroutines used to make these calculations. Three sets of independent variables are permitted in these routines. In addition to the plenum pressure and temperature, the third independent variable is either nozzle exit pressure, Mach number, or temperature. A critical-flow factor that becomes a convenient means for determining the mass flow rate of methane through critical-flow nozzles is tabulated. Other tables are included for nozzle throat velocity and critical pressure, density, and temperature ratios, along with some thermodynamic properties of methane, including compressibility factor, enthalpy, entropy, specific heat, specific-heat ratio, and speed of sound. These tabulations cover a temperature range from 120 to 600 K and pressures to 3 million N/sq m.

Computational Study of Axisymmetric Off-Design Nozzle Flows

NASA Technical Reports Server (NTRS)

DalBello, Teryn; Georgiadis, Nicholas; Yoder, Dennis; Keith, Theo

2003-01-01

Computational Fluid Dynamics (CFD) analyses of axisymmetric circular-arc boattail nozzles operating off-design at transonic Mach numbers have been completed. These computations span the very difficult transonic flight regime with shock-induced separations and strong adverse pressure gradients. External afterbody and internal nozzle pressure distributions computed with the Wind code are compared with experimental data. A range of turbulence models were examined, including the Explicit Algebraic Stress model. Computations have been completed at freestream Mach numbers of 0.9 and 1.2, and nozzle pressure ratios (NPR) of 4 and 6. Calculations completed with variable time-stepping (steady-state) did not converge to a true steady-state solution. Calculations obtained using constant timestepping (timeaccurate) indicate less variations in flow properties compared with steady-state solutions. This failure to converge to a steady-state solution was the result of using variable time-stepping with large-scale separations present in the flow. Nevertheless, time-averaged boattail surface pressure coefficient and internal nozzle pressures show reasonable agreement with experimental data. The SST turbulence model demonstrates the best overall agreement with experimental data.

Flow in a planar convergent-divergent nozzle

NASA Astrophysics Data System (ADS)

Kotteda, V. M. K.; Mittal, S.

2017-05-01

Flow in a convergent-divergent nozzle is studied for pressure ratios (NPR) of 1-11 and exit-to-throat area ratios of 1.2 to 2.0. The unsteady compressible Navier-Stokes equations along with the Spalart-Allmaras turbulence model are solved using a stabilized finite element method in two dimensions. Asymmetric flow is observed at moderate NPR. The side loads due to the flow asymmetry increase with increases in NPR and area ratio. Various flow regimes that are possible in the entire parameter space are identified. The introduction of boundary layer bleed results in steady and symmetric flow conditions at all NPR. Consequently, the nozzle does not experience a lateral force for any NPR. Application of bleed leads to a significant downstream shift in the shock location at low to moderate NPR. Compared to no-bleed, the nozzle experiences a loss of thrust in this regime. The thrust performance for {NPR} > 6 is, however, unaffected by bleed. The effect of nozzle geometry on the flow at various NPR is studied. Four different geometries with the same area ratio and nozzle length are considered. These geometries differ from each other in terms of the nozzle surface profile, including the discontinuity in slope of the surface. Barring some minor differences at low to moderate NPR, the flow is similar for all the geometries considered.

Distributed Exhaust Nozzles for Jet Noise Reduction

NASA Technical Reports Server (NTRS)

Ahuja, K. K.; Gaeta, R. J.; Hellman, B.; Schein, D. B.; Solomon, W. D., Jr.; Huff, Dennis (Technical Monitor)

2001-01-01

The main objective of this study is to validate the jet noise reduction potential of a concept associated with distributed exhaust nozzles. Under this concept the propulsive thrust is generated by a larger number of discrete plumes issuing from an array of small or mini-nozzles. The potential of noise reduction of this concept stems from the fact that a large number of small jets will produce very high frequency noise and also, if spaced suitably, they will coalesce at a smaller velocity to produce low amplitude, low frequency noise. This is accomplished through detailed acoustic and fluid measurements along with a Computational Fluidic Dynamic (CFD) solution of the mean (DE) Distributed Exhaust nozzle flowfield performed by Northrop-Grumman. The acoustic performance is quantified in an anechoic chamber. Farfield acoustic data is acquired for a DE nozzle as well as a round nozzle of the same area. Both these types of nozzles are assessed numerically using Computational Fluid Dynamic (CFD) techniques. The CFD analysis ensures that both nozzles issued the same amount of airflow for a given nozzle pressure ratio. Data at a variety of nozzle pressure ratios are acquired at a range of polar and azimuthal angles. Flow visualization of the DE nozzle is used to assess the fluid dynamics of the small jet interactions. Results show that at high subsonic jet velocities, the DE nozzle shifts its frequency of peak amplitude to a higher frequency relative to a round nozzle of equivalent area (from a S(sub tD) = 0.24 to 1. 3). Furthermore, the DE nozzle shows reduced sound pressure levels (as much as 4 - 8 dB) in the low frequency part of the spectrum (less than S(sub tD) = 0.24 ) compared to the round nozzle. At supersonic jet velocities, the DE nozzle does not exhibit the jet screech and the shock-associated broadband noise is reduced by as much as 12 dB.

Computations of Internal and External Axisymmetric Nozzle Aerodynamics at Transonic Speeds

NASA Technical Reports Server (NTRS)

Dalbello, Teryn; Georgiadis, Nicholas; Yoder, Dennis; Keith, Theo

2003-01-01

Computational Fluid Dynamics (CFD) analyses of axisymmetric circular-arc boattail nozzles have been completed in support of NASA's Next Generation Launch Technology Program to investigate the effects of high-speed nozzle geometries on the nozzle internal flow and the surrounding boattail regions. These computations span the very difficult transonic flight regime, with shock-induced separations and strong adverse pressure gradients. External afterbody and internal nozzle pressure distributions computed with the Wind code are compared with experimental data. A range of turbulence models were examined in Wind, including an Explicit Algebraic Stress model (EASM). Computations on two nozzle geometries have been completed at freestream Mach numbers ranging from 0.6 to 0.9, driven by nozzle pressure ratios (NPR) ranging from 2.9 to 5. Results obtained on converging-only geometry indicate reasonable agreement to experimental data, with the EASM and Shear Stress Transport (SST) turbulence models providing the best agreement. Calculations completed on a converging-diverging geometry involving large-scale internal flow separation did not converge to a true steady-state solution when run with variable timestepping (steady-state). Calculations obtained using constant timestepping (time-accurate) indicate less variations in flow properties compared with steady-state solutions. This failure to converge to a steady-state solution was found to be the result of difficulties in using variable time-stepping with large-scale separations present in the flow. Nevertheless, time-averaged boattail surface pressure coefficient and internal nozzle pressures show fairly good agreement with experimental data. The SST turbulence model demonstrates the best over-all agreement with experimental data.

Compressed air noise reductions from using advanced air gun nozzles in research and development environments.

PubMed

Prieve, Kurt; Rice, Amanda; Raynor, Peter C

2017-08-01

The aims of this study were to evaluate sound levels produced by compressed air guns in research and development (R&D) environments, replace conventional air gun models with advanced noise-reducing air nozzles, and measure changes in sound levels to assess the effectiveness of the advanced nozzles as engineering controls for noise. Ten different R&D manufacturing areas that used compressed air guns were identified and included in the study. A-weighted sound level and Z-weighted octave band measurements were taken simultaneously using a single instrument. In each area, three sets of measurements, each lasting for 20 sec, were taken 1 m away and perpendicular to the air stream of the conventional air gun while a worker simulated typical air gun work use. Two different advanced noise-reducing air nozzles were then installed. Sound level and octave band data were collected for each of these nozzles using the same methods as for the original air guns. Both of the advanced nozzles provided sound level reductions of about 7 dBA, on average. The highest noise reductions measured were 17.2 dBA for one model and 17.7 dBA for the other. In two areas, the advanced nozzles yielded no sound level reduction, or they produced small increases in sound level. The octave band data showed strong similarities in sound level among all air gun nozzles within the 10-1,000 Hz frequency range. However, the advanced air nozzles generally had lower noise contributions in the 1,000-20,000 Hz range. The observed decreases at these higher frequencies caused the overall sound level reductions that were measured. Installing new advanced noise-reducing air nozzles can provide large sound level reductions in comparison to existing conventional nozzles, which has direct benefit for hearing conservation efforts.

Investigation of ramp injectors for supersonic mixing enhancement

NASA Technical Reports Server (NTRS)

Haimovitch, Y.; Gartenberg, E.; Roberts, A. S., Jr.

1994-01-01

A comparative study of wall mounted swept ramp injectors fitted with injector nozzles of different shape has been conducted in a constant area duct to explore mixing enhancement techniques for scramjet combustors. Six different injector nozzle inserts, all having equal exit and throat areas, were tested to explore the interaction between the preconditioned fuel jet and the vortical flowfield produced by the ramp: circular nozzle (baseline), nozzle with three downstream facing steps, nozzle with four vortex generators, elliptical nozzle, tapered-slot nozzle, and trapezoidal nozzle. The main flow was air at Mach 2, and the fuel was simulated by air injected at Mach 1.63 or by helium injected at Mach 1.7. Pressure and temperature surveys, combined with Mie and Rayleigh scattering visualization, were used to investigate the flow field. The experiments were compared with three dimensional Navier-Stokes computations. The results indicate that the mixing process is dominated by the streamwise vorticity generated by the ramp, the injectors' inner geometry having a minor effect. It was also found that the injectant/air mixing in the far-field is nearly independent of the injector geometry, molecular weight of the injectant, and the initial convective Mach number.

Improvement on Main/backup Controller Switching Device of the Nozzle Throat Area Control System for a Turbofan Aero Engine

NASA Astrophysics Data System (ADS)

Li, Jie; Duan, Minghu; Yan, Maode; Li, Gang; Li, Xiaohui

2014-06-01

A full authority digital electronic controller (FADEC) equipped with a full authority hydro-mechanical backup controller (FAHMBC) is adopted as the nozzle throat area control system (NTACS) of a turbofan aero engine. In order to ensure the switching reliability of the main/backup controller, the nozzle throat area control switching valve was improved from three-way convex desktop slide valve to six-way convex desktop slide valve. Simulation results show that, if malfunctions of FAEDC occur and abnormal signals are outputted from FADEC, NTACS will be seriously influenced by the main/backup controller switching in several working states, while NTACS will not be influenced by using the improved nozzle throat area control switching valve, thus the controller switching process will become safer and smoother and the working reliability of this turbofan aero engine is improved by the controller switching device improvement.

Advanced Space Propulsion System Flowfield Modeling

NASA Technical Reports Server (NTRS)

Smith, Sheldon

1998-01-01

Solar thermal upper stage propulsion systems currently under development utilize small low chamber pressure/high area ratio nozzles. Consequently, the resulting flow in the nozzle is highly viscous, with the boundary layer flow comprising a significant fraction of the total nozzle flow area. Conventional uncoupled flow methods which treat the nozzle boundary layer and inviscid flowfield separately by combining the two calculations via the influence of the boundary layer displacement thickness on the inviscid flowfield are not accurate enough to adequately treat highly viscous nozzles. Navier Stokes models such as VNAP2 can treat these flowfields but cannot perform a vacuum plume expansion for applications where the exhaust plume produces induced environments on adjacent structures. This study is built upon recently developed artificial intelligence methods and user interface methodologies to couple the VNAP2 model for treating viscous nozzle flowfields with a vacuum plume flowfield model (RAMP2) that is currently a part of the Plume Environment Prediction (PEP) Model. This study integrated the VNAP2 code into the PEP model to produce an accurate, practical and user friendly tool for calculating highly viscous nozzle and exhaust plume flowfields.

Turbofan forced mixer-nozzle internal flowfield. Volume 1: A benchmark experimental study

NASA Technical Reports Server (NTRS)

Paterson, R. W.

1982-01-01

An experimental investigation of the flow field within a model turbofan forced mixer nozzle is described. Velocity and thermodynamic state variable data for use in assessing the accuracy and assisting the further development of computational procedures for predicting the flow field within mixer nozzles are provided. Velocity and temperature data suggested that the nozzle mixing process was dominated by circulations (secondary flows) of a length scale on the order the lobe dimensions which were associated with strong radial velocities observed near the lobe exit plane. The 'benchmark' model mixer experiment conducted for code assessment purposes is discussed.

Experimental and Computational Investigation of a Translating-Throat Single-Expansion-Ramp Nozzle

NASA Technical Reports Server (NTRS)

Deere, Karen A.; Asbury, Scott C.

1999-01-01

An experimental and computational study was conducted on a high-speed, single-expansion-ramp nozzle (SERN) concept designed for efficient off-design performance. The translating-throat SERN concept adjusts the axial location of the throat to provide a variable expansion ratio and allow a more optimum jet exhaust expansion at various flight conditions in an effort to maximize nozzle performance. Three design points (throat locations) were investigated to simulate the operation of this concept at subsonic-transonic, low supersonic, and high supersonic flight conditions. The experimental study was conducted in the jet exit test facility at the Langley Research Center. Internal nozzle performance was obtained at nozzle pressure ratios (NPR's) up to 13 for six nozzles with design nozzle pressure ratios near 9, 42, and 102. Two expansion-ramp surfaces, one concave and one convex, were tested for each design point. Paint-oil flow and focusing schlieren flow visualization techniques were utilized to acquire additional flow data at selected NPR'S. The Navier-Stokes code, PAB3D, was used with a two-equation k-e turbulence model for the computational study. Nozzle performance characteristics were predicted at nozzle pressure ratios of 5, 9, and 13 for the concave ramp, low Mach number nozzle and at 10, 13, and 102 for the concave ramp, high Mach number nozzle.

A static investigation of several STOVL exhaust system concepts

NASA Technical Reports Server (NTRS)

Romine, B. M., Jr.; Meyer, B. E.; Re, R. J.

1989-01-01

A static cold flow scale model test was performed in order to determine the internal performance characteristics of various STOVL exhaust systems. All of the concepts considered included a vectorable cruise nozzle and a separate vectorable vertical thrust ventral nozzle mounted on the tailpipe. The two ventral nozzle configurations tested featured vectorable constant thickness cascade vanes for area control and improved performance during transition and vertical lift flight. The best transition performance was achieved using a butterfly door type ventral nozzle and a pitch vectoring 2DCD or axisymmetric cruise nozzle. The clamshell blocker type of ventral nozzle had reduced transition performance due to the choking of the tailpipe flow upstream of the cruise nozzle.

Vanishing Viscosity Approach to the Compressible Euler Equations for Transonic Nozzle and Spherically Symmetric Flows

NASA Astrophysics Data System (ADS)

Chen, Gui-Qiang G.; Schrecker, Matthew R. I.

2018-04-01

We are concerned with globally defined entropy solutions to the Euler equations for compressible fluid flows in transonic nozzles with general cross-sectional areas. Such nozzles include the de Laval nozzles and other more general nozzles whose cross-sectional area functions are allowed at the nozzle ends to be either zero (closed ends) or infinity (unbounded ends). To achieve this, in this paper, we develop a vanishing viscosity method to construct globally defined approximate solutions and then establish essential uniform estimates in weighted L p norms for the whole range of physical adiabatic exponents γ\\in (1, ∞) , so that the viscosity approximate solutions satisfy the general L p compensated compactness framework. The viscosity method is designed to incorporate artificial viscosity terms with the natural Dirichlet boundary conditions to ensure the uniform estimates. Then such estimates lead to both the convergence of the approximate solutions and the existence theory of globally defined finite-energy entropy solutions to the Euler equations for transonic flows that may have different end-states in the class of nozzles with general cross-sectional areas for all γ\\in (1, ∞) . The approach and techniques developed here apply to other problems with similar difficulties. In particular, we successfully apply them to construct globally defined spherically symmetric entropy solutions to the Euler equations for all γ\\in (1, ∞).

Noise tests of a model engine-over-the-wing STOL configuration using a multijet nozzle with deflector

NASA Technical Reports Server (NTRS)

Olsen, W. A.; Friedman, R.

1973-01-01

Noise data were obtained with a small scale model stationary STOL configuration that used an eight lobe mixer nozzle with deflector mounted above a 32-cm-chord wing section. The factors varied to determine their effect upon the noise were wing flap angle, nozzle shape, nozzle location, deflector configuration, and jet velocity. The noise from the mixer nozzle model was compared to the noise from a model using a circular nozzle of the same area. The mixer nozzle model was quieter at the low to middle frequencies, while the circular nozzle was quieter at high frequencies. The perceived noise level (PNL) was calculated for an aircraft 10 times larger than the model. The PNL at 500 feet for the mixer nozzle turned out to be within 1 db of the PNL for the circular nozzle. For some configurations at highly directional broadband noise, which could be eliminated by changes in nozzle and/or deflector location, occurred below the wing.

Three-Dimensional Computational Model for Flow in an Over-Expanded Nozzle With Porous Surfaces

NASA Technical Reports Server (NTRS)

Abdol-Hamid, K. S.; Elmiligui, Alaa; Hunter, Craig A.; Massey, Steven J.

2006-01-01

A three-Dimensional computational model is used to simulate flow in a non-axisymmetric, convergent-divergent nozzle incorporating porous cavities for shock-boundary layer interaction control. The nozzle has an expansion ratio (exit area/throat area) of 1.797 and a design nozzle pressure ratio of 8.78. Flow fields for the baseline nozzle (no porosity) and for the nozzle with porous surfaces of 10% openness are computed for Nozzle Pressure Ratio (NPR) varying from 1.29 to 9.54. The three dimensional computational results indicate that baseline (no porosity) nozzle performance is dominated by unstable, shock-induced, boundary-layer separation at over-expanded conditions. For NPR less than or equal to 1.8, the separation is three dimensional, somewhat unsteady, and confined to a bubble (with partial reattachment over the nozzle flap). For NPR greater than or equal to 2.0, separation is steady and fully detached, and becomes more two dimensional as NPR increased. Numerical simulation of porous configurations indicates that a porous patch is capable of controlling off design separation in the nozzle by either alleviating separation or by encouraging stable separation of the exhaust flow. In the present paper, computational simulation results, wall centerline pressure, mach contours, and thrust efficiency ratio are presented, discussed and compared with experimental data. Results indicate that comparisons are in good agreement with experimental data. The three-dimensional simulation improves the comparisons for over-expanded flow conditions as compared with two-dimensional assumptions.

Nozzle airfoil having movable nozzle ribs

DOEpatents

Yu, Yufeng Phillip; Itzel, Gary Michael

2002-01-01

A nozzle vane or airfoil structure is provided in which the nozzle ribs are connected to the side walls of the vane or airfoil in such a way that the ribs provide the requisite mechanical support between the concave side and convex side of the airfoil but are not locked in the radial direction of the assembly, longitudinally of the airfoil. The ribs may be bi-cast onto a preformed airfoil side wall structure or fastened to the airfoil by an interlocking slide connection and/or welding. By attaching the nozzle ribs to the nozzle airfoil metal in such a way that allows play longitudinally of the airfoil, the temperature difference induced radial thermal stresses at the nozzle airfoil/rib joint area are reduced while maintaining proper mechanical support of the nozzle side walls.

Highly Variable Cycle Exhaust Model Test (HVC10)

NASA Technical Reports Server (NTRS)

Henderson, Brenda; Wernet, Mark; Podboy, Gary; Bozak, Rick

2010-01-01

Results from acoustic and flow-field studies using the Highly Variable Cycle Exhaust (HVC) model were presented. The model consisted of a lobed mixer on the core stream, an elliptic nozzle on the fan stream, and an ejector. For baseline comparisons, the fan nozzle was replaced with a round nozzle and the ejector doors were removed from the model. Acoustic studies showed far-field noise levels were higher for the HVC model with the ejector than for the baseline configuration. Results from Particle Image Velocimetry (PIV) studies indicated that large flow separation regions occurred along the ejector doors, thus restricting flow through the ejector. Phased array measurements showed noise sources located near the ejector doors for operating conditions where tones were present in the acoustic spectra.

The Effect of Nozzle Trailing Edge Thickness on Jet Noise

NASA Technical Reports Server (NTRS)

Henderson, Brenda; Kinzie, Kevin; Haskin, Henry

2004-01-01

The effect of nozzle trailing edge thickness on broadband acoustic radiation and the production of tones is investigated for coannular nozzles. Experiments were performed for a core nozzle trailing edge thickness between 0.38 mm and 3.17 mm. The on-set of discrete tones was found to be predominantly affected by the velocity ratio, the ratio of the fan velocity to the core velocity, although some dependency on trailing edge thickness was also noted. For a core nozzle trailing edge thickness greater than or equal to 0.89 mm, tones were produced for velocity ratios between 0.91 and 1.61. For a constant nozzle trailing edge thickness, the frequency varied almost linearly with the core velocity. The Strouhal number based on the core velocity changed with nozzle trailing edge thickness and varied between 0.16 and 0.2 for the core nozzles used in the experiments. Increases in broadband noise with increasing trailing edge thickness were observed for tone producing and non-tone producing conditions. A variable thickness trailing edge (crenellated) nozzle resulted in no tonal production and a reduction of the broadband trailing edge noise relative to that of the corresponding constant thickness trailing edge.

The effects of finite rate chemical processes on high enthalpy nozzle performance - A comparison between SPARK and SEAGULL

NASA Technical Reports Server (NTRS)

Carpenter, M. H.

1988-01-01

The generalized chemistry version of the computer code SPARK is extended to include two higher-order numerical schemes, yielding fourth-order spatial accuracy for the inviscid terms. The new and old formulations are used to study the influences of finite rate chemical processes on nozzle performance. A determination is made of the computationally optimum reaction scheme for use in high-enthalpy nozzles. Finite rate calculations are compared with the frozen and equilibrium limits to assess the validity of each formulation. In addition, the finite rate SPARK results are compared with the constant ratio of specific heats (gamma) SEAGULL code, to determine its accuracy in variable gamma flow situations. Finally, the higher-order SPARK code is used to calculate nozzle flows having species stratification. Flame quenching occurs at low nozzle pressures, while for high pressures, significant burning continues in the nozzle.

Internal performance of a hybrid axisymmetric/nonaxisymmetric convergent-divergent nozzle

NASA Technical Reports Server (NTRS)

Taylor, John G.

1991-01-01

An investigation was conducted in the static test facility of the Langley 16-foot transonic tunnel to determine the internal performance of a hybrid axisymmetric/nonaxisymmetric nozzle in forward-thrust mode. Nozzle cross-sections in the spherical convergent section were axisymmetric whereas cross-sections in the divergent flap area nonaxisymmetric (two-dimensional). Nozzle concepts simulating dry and afterburning power settings were investigated. Both subsonic cruise and supersonic cruise expansion ratios were tested for the dry power nozzle concepts. Afterburning power configurations were tested at an expansion ratio typical for subsonic acceleration. The spherical convergent flaps were designed in such a way that the transition from axisymmetric to nonaxisymmetric cross-section occurred in the region of the nozzle throat. Three different nozzle throat geometries were tested for each nozzle power setting. High-pressure air was used to simulate jet exhaust at nozzle pressure ratios up to 12.0.

Acoustic and aerodynamic testing of a scale model variable pitch fan

NASA Technical Reports Server (NTRS)

Jutras, R. R.; Kazin, S. B.

1974-01-01

A fully reversible pitch scale model fan with variable pitch rotor blades was tested to determine its aerodynamic and acoustic characteristics. The single-stage fan has a design tip speed of 1160 ft/sec (353.568 m/sec) at a bypass pressure ratio of 1.5. Three operating lines were investigated. Test results show that the blade pitch for minimum noise also resulted in the highest efficiency for all three operating lines at all thrust levels. The minimum perceived noise on a 200-ft (60.96 m) sideline was obtained with the nominal nozzle. At 44% of takeoff thrust, the PNL reduction between blade pitch and minimum noise blade pitch is 1.8 PNdB for the nominal nozzle and decreases with increasing thrust. The small nozzle (6% undersized) has the highest efficiency at all part thrust conditions for the minimum noise blade pitch setting; although, the noise is about 1.0 PNdB higher for the small nozzle at the minimum noise blade pitch position.

The prediction of nozzle performance and heat transfer in hydrogen/oxygen rocket engines with transpiration cooling, film cooling, and high area ratios

NASA Technical Reports Server (NTRS)

Kacynski, Kenneth J.; Hoffman, Joe D.

1993-01-01

An advanced engineering computational model has been developed to aid in the analysis and design of hydrogen/oxygen chemical rocket engines. The complete multi-species, chemically reacting and diffusing Navier-Stokes equations are modelled, finite difference approach that is tailored to be conservative in an axisymmetric coordinate system for both the inviscid and viscous terms. Demonstration cases are presented for a 1030:1 area ratio nozzle, a 25 lbf film cooled nozzle, and transpiration cooled plug-and-spool rocket engine. The results indicate that the thrust coefficient predictions of the 1030:1 nozzle and the film cooled nozzle are within 0.2 to 0.5 percent, respectively, of experimental measurements when all of the chemical reaction and diffusion terms are considered. Further, the model's predictions agree very well with the heat transfer measurements made in all of the nozzle test cases. The Soret thermal diffusion term is demonstrated to have a significant effect on the predicted mass fraction of hydrogen along the wall of the nozzle in both the laminar flow 1030:1 nozzle and the turbulent plug-and-spool rocket engine analysis cases performed. Further, the Soret term was shown to represent a significant fraction of the diffusion fluxes occurring in the transpiration cooled rocket engine.

Aerolization During Boron Nanoparticle Multi-Component Fuel Group Burning Studies

DTIC Science & Technology

2014-02-03

Anderson, University of Utah). …………………… 14 Figure 2. Photograph of group burning facility showing benchtop flat flame burner unit with injector nozzle ...and (B) aerosol generator. 16 Figure 6. Diagram of benchtop flat flame burner unit showing injector nozzle assembly with VOAG orifice, fuel and...translation stage, variable fuel and gas supply rates, and injector nozzles that can be configured to investigate diffusion and premixed flames (Fig. 2 & 3

Preliminary Internal Performance Data for a Variable-Ejector Assembly on the XJ79-GE-1 Turbojet Engine. II; Afterburning Configurations

NASA Technical Reports Server (NTRS)

Bloomer, Harry E.; Groesbeck, Donald E.

1957-01-01

Internal performance of an XJ79-GE-1 variable ejector was experimentally determined with the primary nozzle in two representative after-burning positions. Jet-thrust and air-handling data were obtained in quiescent air for 4 selected ejector configurations over a wide range of secondary to primary airflow ratios and primary-nozzle pressure ratios. The experimental ejector data are presented in both graphical and tabulated form.

Flight effects on the aero/acoustic characteristics of inverted profile coannular nozzles

NASA Technical Reports Server (NTRS)

Kozlowski, H.; Packman, A. B.

1978-01-01

The effect of simulated flight speed on the acoustic and aerodynamic characteristics of coannular nozzles is examined. The noise and aerodynamic performance of the coannular nozzle exhaust systems over a large range of operating flight conditions is presented. The jet noise levels of the coannular nozzles are discussed. The impact of fan to primary nozzle area ratio and the presence of an ejector on flight effects are investigated. The impact of flight speed on the individual components of the coannular jet noise was ascertained.

Erosion resistant nozzles for laser plasma extreme ultraviolet (EUV) sources

DOEpatents

Kubiak, Glenn D.; Bernardez, II, Luis J.

2000-01-04

A gas nozzle having an increased resistance to erosion from energetic plasma particles generated by laser plasma sources. By reducing the area of the plasma-facing portion of the nozzle below a critical dimension and fabricating the nozzle from a material that has a high EUV transmission as well as a low sputtering coefficient such as Be, C, or Si, it has been shown that a significant reduction in reflectance loss of nearby optical components can be achieved even after exposing the nozzle to at least 10.sup.7 Xe plasma pulses.

Internal performance characteristics of thrust-vectored axisymmetric ejector nozzles

NASA Technical Reports Server (NTRS)

Lamb, Milton

1995-01-01

A series of thrust-vectored axisymmetric ejector nozzles were designed and experimentally tested for internal performance and pumping characteristics at the Langley research center. This study indicated that discontinuities in the performance occurred at low primary nozzle pressure ratios and that these discontinuities were mitigated by decreasing expansion area ratio. The addition of secondary flow increased the performance of the nozzles. The mid-to-high range of secondary flow provided the most overall improvements, and the greatest improvements were seen for the largest ejector area ratio. Thrust vectoring the ejector nozzles caused a reduction in performance and discharge coefficient. With or without secondary flow, the vectored ejector nozzles produced thrust vector angles that were equivalent to or greater than the geometric turning angle. With or without secondary flow, spacing ratio (ejector passage symmetry) had little effect on performance (gross thrust ratio), discharge coefficient, or thrust vector angle. For the unvectored ejectors, a small amount of secondary flow was sufficient to reduce the pressure levels on the shroud to provide cooling, but for the vectored ejector nozzles, a larger amount of secondary air was required to reduce the pressure levels to provide cooling.

Effect of nozzle and vertical-tail variables on the performance of a 3-surface F-15 model at transonic Mach numbers. [Langley 16 foot transonic tunnel

NASA Technical Reports Server (NTRS)

Pendergraft, O. C., Jr.; Bare, E. A.

1982-01-01

An investigation was conducted in the Langley 16 foot transonic tunnel to determine the longitudinal aerodynamic characteristics of twin two dimensional nozzles and twin baseline axisymmetric nozzles installed on a fully metric 0.047 scale model of the F-15 three surface configuration (canards, wing, horizontal tails). The effects on performance of two dimensional nozzle in flight thrust reversing, locations and orientation of the vertical tails, and deflections of the horizontal tails were also determined. Test data were obtained at static conditions and at Mach numbers from 0.60 to 1.20 over an angle of attack range from -2 deg to 15 deg. Nozzle pressure ratio was varied from jet off to about 6.5.

Upper-surface-blowing flow-turning performance

NASA Technical Reports Server (NTRS)

Sleeman, W. C., Jr.; Phelps, A. E., III

1976-01-01

Jet exhaust flow-turning characteristics were determined for systematic variations in upper-surface blowing exhaust nozzles and trailing-edge flap configuration variables from experimental wind-off (static) flow studies. For conditions with parallel flow exhausting from the nozzle, jet height (as indicated by nozzle exit height) and flap radius were found to be the most important parameters relating to flow turning. Nonparallel flow from the nozzle, as obtained from an internal roof angle and/or side spread angle, had a large favorable effect on flow turning. Comparisons made between static turning results and wind tunnel aerodynamic studies of identical configurations indicated that static flow-turning results can be indicative of wind-on powered lift performance for both good and poor nozzle-flap combinations but, for marginal designs, can lead to overly optimistic assessment of powered lift potential.

The Effect of Bypass Nozzle Exit Area on Fan Aerodynamic Performance and Noise in a Model Turbofan Simulator

NASA Technical Reports Server (NTRS)

Hughes, Christopher E.; Podboy, Gary, G.; Woodward, Richard P.; Jeracki, Robert, J.

2013-01-01

The design of effective new technologies to reduce aircraft propulsion noise is dependent on identifying and understanding the noise sources and noise generation mechanisms in the modern turbofan engine, as well as determining their contribution to the overall aircraft noise signature. Therefore, a comprehensive aeroacoustic wind tunnel test program was conducted called the Fan Broadband Source Diagnostic Test as part of the NASA Quiet Aircraft Technology program. The test was performed in the anechoic NASA Glenn 9- by 15-Foot Low Speed Wind Tunnel using a 1/5 scale model turbofan simulator which represented a current generation, medium pressure ratio, high bypass turbofan aircraft engine. The investigation focused on simulating in model scale only the bypass section of the turbofan engine. The test objectives were to: identify the noise sources within the model and determine their noise level; investigate several component design technologies by determining their impact on the aerodynamic and acoustic performance of the fan stage; and conduct detailed flow diagnostics within the fan flow field to characterize the physics of the noise generation mechanisms in a turbofan model. This report discusses results obtained for one aspect of the Source Diagnostic Test that investigated the effect of the bypass or fan nozzle exit area on the bypass stage aerodynamic performance, specifically the fan and outlet guide vanes or stators, as well as the farfield acoustic noise level. The aerodynamic performance, farfield acoustics, and Laser Doppler Velocimeter flow diagnostic results are presented for the fan and four different fixed-area bypass nozzle configurations. The nozzles simulated fixed engine operating lines and encompassed the fan stage operating envelope from near stall to cruise. One nozzle was selected as a baseline reference, representing the nozzle area which would achieve the design point operating conditions and fan stage performance. The total area change from the smallest to the largest nozzle was 12.9 percent of the baseline nozzle area. The results will show that there are significant changes in aerodynamic performance and farfield acoustics as the fan nozzle area is increased. The weight flow through the fan model increased between 7 and 9 percent, the fan and stage pressure dropped between 8 and 10 percent, and the adiabatic efficiency increased between 2 and 3 percent--the magnitude of the change dependent on the fan speed. Results from force balance measurements of fan and outlet guide vane thrust will show that as the nozzle exit area is increased the combined thrust of the fan and outlet guide vanes together also increases, between 2 and 3.5 percent, mainly due to the increase in lift from the outlet guide vanes. In terms of farfield acoustics, the overall sound power level produced by the fan stage dropped nearly linearly between 1 dB at takeoff condition and 3.5 dB at approach condition, mainly due to a decrease in the broadband noise levels. Finally, fan swirl angle survey and Laser Doppler Velocimeter mean velocity and turbulence data obtained in the fan wake will show that the swirl angles and turbulence levels within the wake decrease as the fan nozzle area increases, which helps to explain the drop in the fan broadband noise at all fan speeds.

Analysis, design and testing of high pressure waterjet nozzles

NASA Technical Reports Server (NTRS)

Mazzoleni, Andre P.

1996-01-01

The Hydroblast Research Cell at MSFC is both a research and a processing facility. The cell is used to investigate fundamental phenomena associated with waterjets as well as to clean hardware for various NASA and contractor projects. In the area of research, investigations are made regarding the use of high pressure waterjets to strip paint, grease, adhesive and thermal spray coatings from various substrates. Current industrial methods of cleaning often use ozone depleting chemicals (ODC) such as chlorinated solvents, and high pressure waterjet cleaning has proven to be a viable alternative. Standard methods of waterjet cleaning use hand held or robotically controlled nozzles. The nozzles used can be single-stream or multijet nozzles, and the multijet nozzles may be mounted in a rotating head or arranged in a fan-type shape. We consider in this paper the use of a rotating, multijet, high pressure water nozzle which is robotically controlled. This method enables rapid cleaning of a large area, but problems such as incomplete coverage (e.g. the formation of 'islands' of material not cleaned) and damage to the substrate from the waterjet have been observed. In addition, current stripping operations require the nozzle to be placed at a standoff distance of approximately 2 inches in order to achieve adequate performance. This close proximity of the nozzle to the target to be cleaned poses risks to the nozzle and the target in the event of robot error or the striking of unanticipated extrusions on the target surface as the nozzle sweeps past. Two key motivations of this research are to eliminate the formation of 'coating islands' and to increase the allowable standoff distance of the nozzle.

Liquid rocket engine nozzles

NASA Technical Reports Server (NTRS)

1976-01-01

The nozzle is a major component of a rocket engine, having a significant influence on the overall engine performance and representing a large fraction of the engine structure. The design of the nozzle consists of solving simultaneously two different problems: the definition of the shape of the wall that forms the expansion surface, and the delineation of the nozzle structure and hydraulic system. This monography addresses both of these problems. The shape of the wall is considered from immediately upstream of the throat to the nozzle exit for both bell and annular (or plug) nozzles. Important aspects of the methods used to generate nozzle wall shapes are covered for maximum-performance shapes and for nozzle contours based on criteria other than performance. The discussion of structure and hydraulics covers problem areas of regeneratively cooled tube-wall nozzles and extensions; it treats also nozzle extensions cooled by turbine exhaust gas, ablation-cooled extensions, and radiation-cooled extensions. The techniques that best enable the designer to develop the nozzle structure with as little difficulty as possible and at the lowest cost consistent with minimum weight and specified performance are described.

Comparison of NASA and contractor results from aeroacoustic tests of QCSEE OTW engine

NASA Technical Reports Server (NTRS)

Bloomer, H. E.; Loeffler, I. J.; Kreim, W. J.; Coats, J. W.

1981-01-01

The aerodynamics and acoustics of the over-the-wing (OTW) Quiet, Clean, Short Haul Experimental Engine (QCSEE) were tested. A boilerplate (nonflight weight), high-throat Mach number, acoustically treated inlet and a D-shaped OTW exhaust nozzle with variable position side doors were used. Some acoustic directivity results for the type "D" nozzle and acoustic effects of variations in the nozzle side door positions are included. It was found that the results are in agreement with those previously obtained.

Exhaust-stack nozzle area and shape for individual cylinder exhaust-gas jet-propulsion system

NASA Technical Reports Server (NTRS)

Pinkel, Benjamin; Turner, Richard; Voss, Fred; Humble, Leroy V

1943-01-01

This report presents the results of an investigation conducted on the effect of exhaust-stack nozzle area, shape, and length on engine power, jet thrust, and gain in net thrust (engine propeller plus jet). Single-cylinder engine data were obtained using three straight stacks 25, 44, and 108 inches in length; an S-shaped stack, a 90 degree bend, a 180 degree bend, and a short straight stack having a closed branch faired into it. Each stack was fitted with nozzles varying in exit area from 0.91 square inch to the unrestricted area of the stack of 4.20 square inches. The engine was generally operated over a range of engine speeds from 1300 to 2100 r.p.m, inlet-manifold pressures from 22 to 30 inches of mercury absolute, and a fuel-air ratio of 0.08. The loss in engine power, the jet thrust, and the gain in net thrust are correlated in terms of several simple parameters. An example is given for determining the optimum nozzle area and the overall net thrust.

Acoustic, Flow Related, and Performance Related Experimental Results for Generation 1.5 High Speed Civil Transport (HSCT) 2-Dimensional Exhaust Nozzles

NASA Technical Reports Server (NTRS)

Salikuddin, M.; Wisler, S.; Majjigi, R.

2004-01-01

The principle objectives of the current program were to experimentally investigate the repeatability of acoustic and aerodynamic characteristics of 2D-CD mixer-ejector nozzles and the effects on the acoustic and aerodynamic characteristics of 2D mixer-ejectors due to (1) the configurational variations, which include mixers with aligned CD chutes, aligned convergent chutes, and staggered CD chutes and aerodynamic cycle variables, (2) treatment variations by using different treatment materials, treating the ejector with varying area, location, and treatment thickness for a mixer-ejector configuration, and (3) secondary inlet shape (i.e., a more realistic inlet) and the blockage across the inlet (a possible fin-like structure needed for installation purpose) by modifying one of the inlet of a mixer-ejector configuration. The objectives also included the measurement dynamic pressures internal to the ejector for a few selected configuration to examine the internal noise characteristics.

Analysis of turbojet-engine controls for afterburning starting

NASA Technical Reports Server (NTRS)

Phillips, W E , Jr

1956-01-01

A simulation procedure is developed for studying the effects of an afterburner start on a controlled turbojet engine. The afterburner start is represented by introducing a step decrease in the effective exhaust-nozzle area, after which the control returns the controlled engine variables to their initial values. The degree and speed with which the control acts are a measure of the effectiveness of the particular control system. Data are presented from five systems investigated using an electronic analog computer and the developed simulation procedure. These systems are compared with respect to steady-state errors, speed of response, and transient deviations of the system variables.

Design and Analyses of High Aspect Ratio Nozzles for Distributed Propulsion Acoustic Measurements

NASA Technical Reports Server (NTRS)

Dippold, Vance F., III

2016-01-01

A series of three convergent, round-to-rectangular high aspect ratio (HAR) nozzles were designed for acoustic testing at the NASA Glenn Research Center Nozzle Acoustic Test Rig (NATR). The HAR nozzles had exit area aspect ratios of 8:1, 12:1, and 16:1. The nozzles were designed to mimic a distributed propulsion system array with a slot nozzle. The nozzle designs were screened using Reynolds-Averaged Navier-Stokes (RANS) simulations. In addition to meeting the geometric constraints required for testing in the NATR, the HAR nozzles were designed to be free of flow features that would produce unwanted noise (e.g., flow separations) and to have uniform flow at the nozzle exit. Multiple methods were used to generate HAR nozzle designs. The final HAR nozzle designs were generated in segments using a computer code that parameterized each segment. RANS screening simulations showed that intermediate nozzle designs suffered flow separation, a normal shockwave at the nozzle exit (caused by an aerodynamic throat produced by boundary layer growth), and non-uniform flow at the nozzle exit. The RANS simulations showed that the final HAR nozzle designs were free of flow separations, but were not entirely successful at producing a fully uniform flow at the nozzle exit. The final designs suffered a pair of counter-rotating vortices along the outboard walls of the nozzle. The 16:1 aspect ratio HAR nozzle had the least uniform flow at the exit plane; the 8:1 aspect ratio HAR nozzles had a fairly uniform flow at the nozzle exit plane.

Flight effects on the aerodynamic and acoustic characteristics of inverted profile coannular nozzles, volume 2. [supersonic cruise aircraft research wind tunnel tests

NASA Technical Reports Server (NTRS)

Kozlowski, H.; Packman, A. B.

1978-01-01

Data from the acoustic tests of the convergent reference nozzle and the 0.75 area ratio coannular nozzle are presented in tables. Data processing routines used to scale the acoustic data and to correct the data for atmospheric attenuation are included.

Internal performance of a nonaxisymmetric nozzle with a rotating upper flap and a center-pivoted lower flap

NASA Technical Reports Server (NTRS)

Wing, David J.; Leavitt, Laurence D.; Re, Richard J.

1993-01-01

An investigation was conducted at wind-off conditions in the static-test facility of the Langley 16-Foot Transonic Tunnel to determine the internal performance characteristics of a single expansion-ramp nozzle with thrust-vectoring capability to 105 degrees. Thrust vectoring was accomplished by the downward rotation of an upper flap with adaptive capability for internal contouring and a corresponding rotation of a center-pivoted lower flap. The static internal performance of configurations with pitch thrust-vector angles of 0 degrees, 60 degrees, and 105 degrees each with two throat areas, was investigated. The nozzle pressure ratio was varied from 1.5 to approximately 8.0 (5.0 for the maximum throat area configurations). Results of this study indicated that the nozzle configuration of the present investigation, when vectored, provided excellent flow-turning capability with relatively high levels of internal performance. In all cases, the thrust vector angle was a function of the nozzle pressure ratio. This result is expected because the flow is bounded by a single expansion surface on both vectored- and unvectored-nozzle geometries.

Arcjet nozzle design impacts

NASA Technical Reports Server (NTRS)

Curran, Francis M.; Sovie, Amy J.; Haag, Thomas W.

1989-01-01

The effect of nozzle configuration on the operating characteristics of a low power dc arcjet thruster was determined. A conical nozzle with a 30 deg converging angle, a 20 deg diverging angle, and an area ratio of 225 served as the baseline case. Variations on the geometry included bell-shaped contours both up and downstream, and a downstream trumpet-shaped contour. The nozzles were operated over a range of specific power near that anticipated for on-orbit operation. Mass flow rate, thrust, current, and voltage were monitored to provide accurate comparisons between nozzles. The upstream contour was found to have minimal effect on arcjet operation. It was determined that the contour of the divergent section of the nozzle, that serves as the anode, was very important in determining the location of arc attachment, and thus had a significant impact on arcjet performance. The conical nozzle was judged to have the optimal current/voltage characteristics and produced the best performance of the nozzles tested.

Arcjet Nozzle Design Impacts

NASA Technical Reports Server (NTRS)

Curran, Francis M.; Sovie, Amy J.; Haag, Thomas W.

1989-01-01

The effect of nozzle configuration on the operating characteristics of a low power dc arcjet thruster was determined. A conical nozzle with a 30 deg converging angle, a 20 deg diverging angle, and an area ratio of 225 served as the baseline case. Variations on the geometry included bell-shaped contours both up and downstream, and a downstream trumpet-shaped contour. The nozzles were operated over a range of specific power near that anticipated for on-orbit operation. Mass flow rate, thrust, current, and voltage were monitored to provide accurate comparisons between nozzles. The upstream contour was found to have minimal effect on arcjet operation. It was determined that the contour of the divergent section of the nozzle, that serves as the anode, was very important in determining the location of arc attachment, and thus had a significant impact on arcjet performance. The conical nozzle was judged to have the optimal current/voltage characteristics and produced the best performance of the nozzles tested.

Static internal performance of a two-dimensional convergent-divergent nozzle with thrust vectoring

NASA Technical Reports Server (NTRS)

Bare, E. Ann; Reubush, David E.

1987-01-01

A parametric investigation of the static internal performance of multifunction two-dimensional convergent-divergent nozzles has been made in the static test facility of the Langley 16-Foot Transonic Tunnel. All nozzles had a constant throat area and aspect ratio. The effects of upper and lower flap angles, divergent flap length, throat approach angle, sidewall containment, and throat geometry were determined. All nozzles were tested at a thrust vector angle that varied from 5.60 tp 23.00 deg. The nozzle pressure ratio was varied up to 10 for all configurations.

Aeroacoustics of contoured and solid/porous conical plug-nozzle supersonic jet flows

NASA Technical Reports Server (NTRS)

Dosanjh, D. S.; Das, I. S.

1985-01-01

The acoustic far field, the shock-associated noise and characteristics of the repetitive shock structure of supersonic jet flows issuing from a contoured plug-nozzle and uncontoured plug-nozzle having a short conical plug of either a solid or a combination of solid/porous surface with pointed termination operated at a range of supercritical pressure are reported. The contoured and the uncontoured plug-nozzles had the same throat area and the same annular-radius ratio.

Computational Fluid Dynamic (CFD) analysis of axisymmetric plume and base flow of film/dump cooled rocket nozzle

NASA Technical Reports Server (NTRS)

Tucker, P. K.; Warsi, S. A.

1993-01-01

Film/dump cooling a rocket nozzle with fuel rich gas, as in the National Launch System (NLS) Space Transportation Main Engine (STME), adds potential complexities for integrating the engine with the vehicle. The chief concern is that once the film coolant is exhausted from the nozzle, conditions may exist during flight for the fuel-rich film gases to be recirculated to the vehicle base region. The result could be significantly higher base temperatures than would be expected from a regeneratively cooled nozzle. CFD analyses were conduced to augment classical scaling techniques for vehicle base environments. The FDNS code with finite rate chemistry was used to simulate a single, axisymmetric STME plume and the NLS base area. Parallel calculations were made of the Saturn V S-1 C/F1 plume base area flows. The objective was to characterize the plume/freestream shear layer for both vehicles as inputs for scaling the S-C/F1 flight data to NLS/STME conditions. The code was validated on high speed flows with relevant physics. This paper contains the calculations for the NLS/STME plume for the baseline nozzle and a modified nozzle. The modified nozzle was intended to reduce the fuel available for recirculation to the vehicle base region. Plumes for both nozzles were calculated at 10kFT and 50kFT.

Flight effects on the aerodynamic and acoustic characteristics of inverted profile coannular nozzles, volume 3. [supersonic cruise aircraft research wind tunnel tests

NASA Technical Reports Server (NTRS)

Kozlowski, H.; Packman, A. B.

1978-01-01

Acoustic data from tests of the 0.75 area ratio coannular nozzle with ejector and the 1.2 area ratio coannular are presented in tables. Aerodynamic data acquired for the four test configurations are included.

Fuselage and nozzle pressure distributions of a 1/12-scale F-15 propulsion model at transonic speeds. Effect of fuselage modifications and nozzle variables

NASA Technical Reports Server (NTRS)

Pendergraft, O. C., Jr.; Carson, G. T., Jr.

1984-01-01

Static pressure coefficient distributions on the forebody, afterbody, and nozzles of a 1/12 scale F-15 propulsion model was determined in the 16 foot transonic tunnel for Mach numbers from 0.60 to 1.20, angles of attack from -2 deg to 7 deg and ratio of jet total pressure to free stream static pressure from 1 up to about 7, depending on Mach number. The effects of nozzle geometry and horizontal tail deflection on the pressure distributions were investigated. Boundary layer total pressure profiles were determined at two locations ahead of the nozzles on the top nacelle surface. Reynolds number varied from about 1.0 x 10 to the 7th power per meter, depending on Mach number.

Exhaust Nozzle Plume Effects on Sonic Boom Test Results for Isolated Nozzles

NASA Technical Reports Server (NTRS)

Castner, Raymond S.

2011-01-01

Reducing or eliminating the operational restrictions of supersonic aircraft over populated areas has led to extensive research at NASA. Restrictions were due to the disturbance of the sonic boom, caused by the coalescence of shock waves formed off the aircraft. Recent work has been performed to reduce the magnitude of the sonic boom N-wave generated by airplane components with focus on shock waves caused by the exhaust nozzle plume. Previous Computational Fluid Dynamics (CFD) analysis showed how the shock wave formed at the nozzle lip interacts with the nozzle boat-tail expansion wave. An experiment was conducted in the 1- by 1-ft Supersonic Wind Tunnel at the NASA Glenn Research Center to validate the computational study. Results demonstrated how the nozzle lip shock moved with increasing nozzle pressure ratio (NPR) and reduced the nozzle boat-tail expansion, causing a favorable change in the observed pressure signature. Experimental results were presented for comparison to the CFD results. The strong nozzle lip shock at high values of NPR intersected the nozzle boat-tail expansion and suppressed the expansion wave. Based on these results, it may be feasible to reduce the boat-tail expansion for a future supersonic aircraft with under-expanded nozzle exhaust flow by modifying nozzle pressure or nozzle divergent section geometry.

Quiet Clean Short-haul Experimental Engine (QCSEE) Under-The-Wing (UTW) composite nacelle subsystem test report. [to verify strength of selected composite materials

NASA Technical Reports Server (NTRS)

Stotler, C. L., Jr.; Johnston, E. A.; Freeman, D. S.

1977-01-01

The element and subcomponent testing conducted to verify the under the wing composite nacelle design is reported. This composite nacelle consists of an inlet, outer cowl doors, inner cowl doors, and a variable fan nozzle. The element tests provided the mechanical properties used in the nacelle design. The subcomponent tests verified that the critical panel and joint areas of the nacelle had adequate structural integrity.

Aerodynamic and acoustic tests of duct-burning turbofan exhaust nozzles

NASA Technical Reports Server (NTRS)

Kozlowski, H.; Packman, A. B.

1976-01-01

The static aerodynamic and acoustic characteristics of duct-burning turbofan (DBTF) exhaust nozzles are established. Scale models, having a total area equivalent to a 0.127 m diameter convergent nozzle, simulating unsuppressed coannular nozzles and mechanically suppressed nozzles with and without ejectors (hardwall and acoustically treated) were tested in a quiescent environment. The ratio of fan to primary area was varied from 0.75 to 1.2. Far field acoustic data, perceived noise levels, and thrust measurements were obtained for 417 test conditions. Pressure ratios were varied from 1.3 to 4.1 in the fan stream and from 1.53 to 2.5 in the primary stream. Total temperature varied from 395 to 1090 K in both streams. Jet noise reductions relative to synthesized prediction from 8 PNdB (with the unsuppressed coannular nozzle) to 15 PNdB (with a mechanically suppressed configuration) were observed at conditions typical of engines being considered under the Advanced Supersonic Technology program. The inherent suppression characteristic of the unsuppressed coannular nozzle is related to the rapid mixing in the jet wake caused by the velocity profiles associated with the DBTF. Since this can be achieved without a mechanical suppressor, significant reductions in aircraft weight or noise footprint can be realized.

Guidelines for Mass Casualty Decontamination During a HAZMAT/Weapon of Mass Destruction Incident. Volumes 1 and 2

DTIC Science & Technology

2009-04-01

Prioritization of victims for decontamination based on injury and evidence of contamination and/or exposure to the hazard. Fog nozzle - Firefighting hose nozzle...A thick, high-pressure hose used to carry water to a fire to extinguish it. Hot Zone - Contaminated area of HAZMAT incident that must be isolated and...evidence of contamination and/or exposure to the hazard. Fog nozzle - Firefighting hose nozzle that separates water into droplets. Hazardous Material

Droplet size distributions of adjuvant-amended sprays from an air-assisted five-port PWM nozzle

USDA-ARS?s Scientific Manuscript database

Verification of droplet size distributions is essential for the development of real-time variable-rate sprayers that synchronize spray outputs with canopy structures. Droplet sizes from a custom-designed, air-assisted, five-port nozzle coupled with a pulse-width-modulated (PWM) solenoid valve were m...

Practical Comparison of Cylindrical Nozzle and De Laval Nozzle for Wire Arc Spraying

NASA Astrophysics Data System (ADS)

Matz, Marc-Manuel; Aumiller, Markus

2014-12-01

In this article, two different nozzle designs (cylindrical nozzle and de Laval nozzle) are compared for use in wire arc spraying. The choice of nozzle is of particular importance because its geometry has a significant influence on the spraying result. The materials used for spraying are steel and copper. By using the de Laval atomizing gas nozzle, the aim is to improve adhesion on the one hand while reducing cost on the other. These objectives have been achieved for the most part, indicating that continued research and development in this area would be useful. Significant potential exists to optimize the efficiency of both the free gas jet and nozzle which have considerable impact on the gas velocity and thus, ultimately, on the spraying result. The measurements carried out have shown that there is a close correlation between the velocity of the gas flow and atomization of the droplets. An explanatory model for varying spraying results with different wire materials using open nozzle systems with de Laval orifice is given and confirmed. For new burner head constructions, an interaction of the atomizing gas nozzle, the contact tips, and wire materials must be considered to achieve all benefits of a de Laval nozzle.

CFD Simulations for Arc-Jet Panel Testing Capability Development Using Semi-Elliptical Nozzles

NASA Technical Reports Server (NTRS)

Gokcen, Tahir; Balboni, John A.; Hartman, G. Joseph

2016-01-01

This paper reports computational simulations in support of arc-jet panel testing capability development using semi-elliptical nozzles in a high enthalpy arc-jet facility at NASA Ames Research Center. Two different semi-elliptical nozzle configurations are proposed for testing panel test articles. Computational fluid dynamics simulations are performed to provide estimates of achievable panel surface conditions and useful test area for each configuration. The present analysis comprises three-dimensional simulations of the nonequilibrium flowfields in the semi-elliptical nozzles, test box and flowfield over the panel test articles. Computations show that useful test areas for the proposed two nozzle options are 20.32 centimeters by 20.32 centimeters (8 inches by 8 inches) and 43.18 centimeters by 43.18 centimeters (17 inches by 17 inches). Estimated values of the maximum cold-wall heat flux and surface pressure are 155 watts per centimeters squared and 39 kilopascals for the smaller panel test option, and 44 watts per centimeters squared and 7 kilopascals for the larger panel test option. Other important properties of the predicted flowfields are presented, and factors that limit the useful test area in the semi-free jet test configuration are discussed.

A Method for Estimating Noise from Full-Scale Distributed Exhaust Nozzles

NASA Technical Reports Server (NTRS)

Kinzie, Kevin W.; Schein, David B.

2004-01-01

A method to estimate the full-scale noise suppression from a scale model distributed exhaust nozzle (DEN) is presented. For a conventional scale model exhaust nozzle, Strouhal number scaling using a scale factor related to the nozzle exit area is typically applied that shifts model scale frequency in proportion to the geometric scale factor. However, model scale DEN designs have two inherent length scales. One is associated with the mini-nozzles, whose size do not change in going from model scale to full scale. The other is associated with the overall nozzle exit area which is much smaller than full size. Consequently, lower frequency energy that is generated by the coalesced jet plume should scale to lower frequency, but higher frequency energy generated by individual mini-jets does not shift frequency. In addition, jet-jet acoustic shielding by the array of mini-nozzles is a significant noise reduction effect that may change with DEN model size. A technique has been developed to scale laboratory model spectral data based on the premise that high and low frequency content must be treated differently during the scaling process. The model-scale distributed exhaust spectra are divided into low and high frequency regions that are then adjusted to full scale separately based on different physics-based scaling laws. The regions are then recombined to create an estimate of the full-scale acoustic spectra. These spectra can then be converted to perceived noise levels (PNL). The paper presents the details of this methodology and provides an example of the estimated noise suppression by a distributed exhaust nozzle compared to a round conic nozzle.

Effects of nozzle interfairing modifications on longitudinal aerodynamic characteristics of a twin jet, variable wing sweep fighter model

NASA Technical Reports Server (NTRS)

Reubush, D. E.; Mercer, C. E.

1975-01-01

A wind-tunnel investigation has been made to determine the effects of nozzle interfairing modifications on the longitudinal aerodynamic characteristics of a twin-jet, variable-wing-sweep fighter model. The model was tested in the Langley 16-foot transonic tunnel at Mach numbers of 0.6 to 1.3 and angles of attack from about minus 2 deg to 6 deg and in the Langley 4-foot supersonic presure tunnel at a Mach number of 2.2 and an angle of attack of 0 deg. Compressed air was used to simulate nozzle exhaust flow at jet total-pressure ratios from 1 (jet off) to about 21. The results of this investigation show that the aircraft drag can be significantly reduced by replacing the basic interfairing with a modified interfairing.

Calibration of the NASA GRC 16 In. Mass-Flow Plug

NASA Technical Reports Server (NTRS)

Davis, David O.; Friedlander, David J.; Saunders, J. David; Frate, Franco C.; Foster, Lancert E.

2012-01-01

The results of an experimental calibration of the NASA Glenn Research Center 16 in. Mass-Flow Plug (MFP) are presented and compared to a previously obtained calibration of a 15 in. Mass-Flow Plug. An ASME low-beta, long-radius nozzle was used as the calibration reference. The discharge coefficient for the ASME nozzle was obtained by numerically simulating the flow through the nozzle from the WIND-US code. The results showed agreement between the 15 in. and 16 in. MFPs for area ratios (MFP to pipe area ratio) greater than 0.6 but deviate at area ratios below this value for reasons that are not fully understood. A general uncertainty analysis was also performed and indicates that large uncertainties in the calibration are present for low MFP area ratios.

Calibration of the NASA Glenn Research Center 16 in. Mass-Flow Plug

NASA Technical Reports Server (NTRS)

Davis, David O.; Friedlander, David J.; Saunders, J. David; Frate, Franco C.; Foster, Lancert E.

2014-01-01

The results of an experimental calibration of the NASA Glenn Research Center 16 in. Mass-Flow Plug (MFP) are presented and compared to a previously obtained calibration of a 15 in. Mass-Flow Plug. An ASME low-beta, long-radius nozzle was used as the calibration reference. The discharge coefficient for the ASME nozzle was obtained by numerically simulating the flow through the nozzle from the WIND-US code. The results showed agreement between the 15 and 16 in. MFPs for area ratios (MFP to pipe area ratio) greater than 0.6 but deviate at area ratios below this value for reasons that are not fully understood. A general uncertainty analysis was also performed and indicates that large uncertainties in the calibration are present for low MFP area ratios.

Investigation of turbines for driving supersonic compressors II : performance of first configuration with 2.2 percent reduction in nozzle flow area / Warner L. Stewart, Harold J. Schum, Robert Y. Wong

NASA Technical Reports Server (NTRS)

Stewart, Warner L; Schum, Harold J; Wong, Robert Y

1952-01-01

The experimental performance of a modified turbine for driving a supersonic compressor is presented and compared with the performance of the original configuration to illustrate the effect of small changes in the ratio of nozzle-throat area to rotor-throat area. Performance is based on the performance of turbines designed to operate with both blade rows close to choking. On the basis of the results of this investigation, the ratio of areas is concluded to become especially critical in the design of turbines such as those designed to drive high-speed, high-specific weight-flow compressors where the turbine nozzles and rotor are both very close to choking.

On the selection of the nozzle geometry and other parameters for cutting corneal flaps with waterjets.

PubMed

Cadavid, Ricardo; Jean, Benedikt; Wüstenberg, Dieter

2009-06-01

A cutting waterjet to produce corneal flaps during refractive surgery or to slice donor corneas for corneal grafting was developed. Jets generated with several different nozzles were compared to determine the most appropriate nozzle geometry for this application. In this paper, it is also discussed how other variables, such as stand-off distance and transverse velocity, can affect the characteristics of the cut. The cutting mechanisms, giving bases for an application of waterjets for cutting other types of tissues, are also discussed.

Advanced supersonic propulsion study, phases 3 and 4. [variable cycle engines

NASA Technical Reports Server (NTRS)

Allan, R. D.; Joy, W.

1977-01-01

An evaluation of various advanced propulsion concepts for supersonic cruise aircraft resulted in the identification of the double-bypass variable cycle engine as the most promising concept. This engine design utilizes special variable geometry components and an annular exhaust nozzle to provide high take-off thrust and low jet noise. The engine also provides good performance at both supersonic cruise and subsonic cruise. Emission characteristics are excellent. The advanced technology double-bypass variable cycle engine offers an improvement in aircraft range performance relative to earlier supersonic jet engine designs and yet at a lower level of engine noise. Research and technology programs required in certain design areas for this engine concept to realize its potential benefits include refined parametric analysis of selected variable cycle engines, screening of additional unconventional concepts, and engine preliminary design studies. Required critical technology programs are summarized.

Analysis and documentation of QCSEE (Quiet Clean Short-haul Experimental Engine) over-the-wing exhaust system development

NASA Technical Reports Server (NTRS)

Ammer, R. C.; Kutney, J. T.

1977-01-01

A static scale model test program was conducted in the static test area of the NASA-Langley 9.14- by 18.29 m(30- by 60-ft) Full-Scale Wind Tunnel Facility to develop an over-the-wing (OTW) nozzle and reverser configuration for the Quiet Clean Short-Haul Experimental Engine (QCSEE). Three nozzles and one basic reverser configuration were tested over the QCSEE takeoff and approach power nozzle pressure ratio range between 1.1 and 1.3. The models were scaled to 8.53% of QCSEE engine size and tested behind two 13.97-cm (5.5-in.) diameter tip-turbine-driven fan simulators coupled in tandem. An OTW nozzle and reverser configuration was identified which satisfies the QCSEE experimental engine requirements in terms of nozzle cycle area variation capability and reverse thrust level, and provides good jet flow spreading over a wing upper surface for achievement of high propulsive lift performance.

Characterization of Three-Stream Jet Flow Fields

NASA Technical Reports Server (NTRS)

Henderson, Brenda S.; Wernet, Mark P.

2016-01-01

Flow-field measurements were conducted on single-, dual- and three-stream jets using two-component and stereo Particle Image Velocimetry (PIV). The flow-field measurements complimented previous acoustic measurements. The exhaust system consisted of externally-plugged, externally-mixed, convergent nozzles. The study used bypass-to-core area ratios equal to 1.0 and 2.5 and tertiary-to-core area ratios equal to 0.6 and 1.0. Axisymmetric and offset tertiary nozzles were investigated for heated and unheated high-subsonic conditions. Centerline velocity decay rates for the single-, dual- and three-stream axisymmetric jets compared well when axial distance was normalized by an equivalent diameter based on the nozzle system total exit area. The tertiary stream had a greater impact on the mean axial velocity for the small bypass-to-core area ratio nozzles than for large bypass-to-core area ratio nozzles. Normalized turbulence intensities were similar for the single-, dual-, and three-stream unheated jets due to the small difference (10 percent) in the core and bypass velocities for the dual-stream jets and the low tertiary velocity (50 percent of the core stream) for the three-stream jets. For heated jet conditions where the bypass velocity was 65 percent of the core velocity, additional regions of high turbulence intensity occurred near the plug tip which were not present for the unheated jets. Offsetting the tertiary stream moved the peak turbulence intensity levels upstream relative to those for all axisymmetric jets investigated.

Characterization of Three-Stream Jet Flow Fields

NASA Technical Reports Server (NTRS)

Henderson, Brenda S.; Wernet, Mark P.

2016-01-01

Flow-field measurements were conducted on single-, dual- and three-stream jets using two-component and stereo Particle Image Velocimetry (PIV). The flow-field measurements complimented previous acoustic measurements. The exhaust system consisted of externally-plugged, externally-mixed, convergent nozzles. The study used bypass-to-core area ratios equal to 1.0 and 2.5 and tertiary-to-core area ratios equal to 0.6 and 1.0. Axisymmetric and offset tertiary nozzles were investigated for heated and unheated high-subsonic conditions. Centerline velocity decay rates for the single-, dual- and three-stream axisymmetric jets compared well when axial distance was normalized by an equivalent diameter based on the nozzle system total exit area. The tertiary stream had a greater impact on the mean axial velocity for the small bypass-to-core area ratio nozzles than for large bypass-to-core area ratio nozzles. Normalized turbulence intensities were similar for the single-, dual-, and three-stream unheated jets due to the small difference (10%) in the core and bypass velocities for the dual-stream jets and the low tertiary velocity (50% of the core stream) for the three-stream jets. For heated jet conditions where the bypass velocity was 65% of the core velocity, additional regions of high turbulence intensity occurred near the plug tip which were not present for the unheated jets. Offsetting the tertiary stream moved the peak turbulence intensity levels upstream relative to those for all axisymmetric jets investigated.

Computational Study of Fluidic Thrust Vectoring using Separation Control in a Nozzle

NASA Technical Reports Server (NTRS)

Deere, Karen; Berrier, Bobby L.; Flamm, Jeffrey D.; Johnson, Stuart K.

2003-01-01

A computational investigation of a two- dimensional nozzle was completed to assess the use of fluidic injection to manipulate flow separation and cause thrust vectoring of the primary jet thrust. The nozzle was designed with a recessed cavity to enhance the throat shifting method of fluidic thrust vectoring. The structured-grid, computational fluid dynamics code PAB3D was used to guide the design and analyze over 60 configurations. Nozzle design variables included cavity convergence angle, cavity length, fluidic injection angle, upstream minimum height, aft deck angle, and aft deck shape. All simulations were computed with a static freestream Mach number of 0.05. a nozzle pressure ratio of 3.858, and a fluidic injection flow rate equal to 6 percent of the primary flow rate. Results indicate that the recessed cavity enhances the throat shifting method of fluidic thrust vectoring and allows for greater thrust-vector angles without compromising thrust efficiency.

A Computational Study of a New Dual Throat Fluidic Thrust Vectoring Nozzle Concept

NASA Technical Reports Server (NTRS)

Deere, Karen A.; Berrier, Bobby L.; Flamm, Jeffrey D.; Johnson, Stuart K.

2005-01-01

A computational investigation of a two-dimensional nozzle was completed to assess the use of fluidic injection to manipulate flow separation and cause thrust vectoring of the primary jet thrust. The nozzle was designed with a recessed cavity to enhance the throat shifting method of fluidic thrust vectoring. Several design cycles with the structured-grid, computational fluid dynamics code PAB3D and with experiments in the NASA Langley Research Center Jet Exit Test Facility have been completed to guide the nozzle design and analyze performance. This paper presents computational results on potential design improvements for best experimental configuration tested to date. Nozzle design variables included cavity divergence angle, cavity convergence angle and upstream throat height. Pulsed fluidic injection was also investigated for its ability to decrease mass flow requirements. Internal nozzle performance (wind-off conditions) and thrust vector angles were computed for several configurations over a range of nozzle pressure ratios from 2 to 7, with the fluidic injection flow rate equal to 3 percent of the primary flow rate. Computational results indicate that increasing cavity divergence angle beyond 10 is detrimental to thrust vectoring efficiency, while increasing cavity convergence angle from 20 to 30 improves thrust vectoring efficiency at nozzle pressure ratios greater than 2, albeit at the expense of discharge coefficient. Pulsed injection was no more efficient than steady injection for the Dual Throat Nozzle concept.

Static Thrust and Vectoring Performance of a Spherical Convergent Flap Nozzle with a Nonrectangular Divergent Duct

NASA Technical Reports Server (NTRS)

Wing, David J.

1998-01-01

The static internal performance of a multiaxis-thrust-vectoring, spherical convergent flap (SCF) nozzle with a non-rectangular divergent duct was obtained in the model preparation area of the Langley 16-Foot Transonic Tunnel. Duct cross sections of hexagonal and bowtie shapes were tested. Additional geometric parameters included throat area (power setting), pitch flap deflection angle, and yaw gimbal angle. Nozzle pressure ratio was varied from 2 to 12 for dry power configurations and from 2 to 6 for afterburning power configurations. Approximately a 1-percent loss in thrust efficiency from SCF nozzles with a rectangular divergent duct was incurred as a result of internal oblique shocks in the flow field. The internal oblique shocks were the result of cross flow generated by the vee-shaped geometric throat. The hexagonal and bowtie nozzles had mirror-imaged flow fields and therefore similar thrust performance. Thrust vectoring was not hampered by the three-dimensional internal geometry of the nozzles. Flow visualization indicates pitch thrust-vector angles larger than 10' may be achievable with minimal adverse effect on or a possible gain in resultant thrust efficiency as compared with the performance at a pitch thrust-vector angle of 10 deg.

Axisymmetric and non-axisymmetric exhaust jet induced effects on a V/STOL vehicle design. Part 2: Analysis of results

NASA Technical Reports Server (NTRS)

Schnell, W. C.

1982-01-01

A wind tunnel investigation, employing a 1/8 scale model in an 11 foot transonic wind tunnel (Mach 0.4-1.4), was conducted to determine the jet effects of several exhaust nozzles on the aeropropulsive performance of a V/STOL fighter design. The force and pressure data show that significant differences in aeropropulsion performance can be expected by varying the exhaust nozzle type, jet area and deflection angle on an underwing nacelle installation. At unvectored conditions, the single expansion ramp nozzles show large performance gains relative to a circular nozzle installation. Additionally, a further drag reduction is realized when the nonaxisymmetric nozzle is vectored through a 10 degree deflection angle. The combined payoff of the vectored nonaxisymmetric nozzle over the baseline circular nozzle installation is equivalent to 25 percent of zero lift drag.

Flight investigation of an underwing nacelle installation of an auxiliary-inlet ejector nozzle with a clamshell flow diverter from Mach 0.6 to 1.3

NASA Technical Reports Server (NTRS)

Head, V. L.

1972-01-01

A nozzle installation of general interest is a podded engine mounted near the aft lower surface of the wing. The effect of this installation on the performance of an auxiliary-inlet ejector nozzle with a clamshell flow diverter was investigated over a Mach number range of 0.6 to 1.3 by using a modified F-106B aircraft. The clamshell flow diverter was tested in a 17 deg position with double-hinged synchronized floating doors. The ejector nozzle trailing-edge flaps were simulated in the closed position with a rigid structure which provided a boattail angle of 10 deg. Primary nozzle area was varied as exhaust gas temperature was varied between 975 and 1561 K. With the nozzle in a subsonic cruise position, the nozzle gross thrust coefficient was 0.918 at a flight Mach number of 0.9.

Investigation of two-dimensional wedge exhaust nozzles for advanced aircraft

NASA Technical Reports Server (NTRS)

Maiden, D. L.; Petit, J. E.

1975-01-01

Two-dimensional wedge nozzle performance characteristics were investigated in a series of wind-tunnel tests. An isolated single-engine/nozzle model was used to study the effects of internal expansion area ratio, aftbody cowl boattail angle, and wedge length. An integrated twin-engine/nozzle model, tested with and without empenage surfaces, included cruise, acceleration, thrust vectoring and thrust reversing nozzle operating modes. Results indicate that the thrust-minus-aftbody drag performance of the twin two-dimensional nozzle integration is significantly higher, for speeds greater than Mach 0.8, than the performance achieved with twin axisymmetric nozzle installations. Significant jet-induced lift was obtained on an aft-mounted lifting surface using a cambered wedge center body to vector thrust. The thrust reversing capabilities of reverser panels installed on the two-dimensional wedge center body were very effective for static or in-flight operation.

Thrust modulation methods for a subsonic V/STOL aircraft

NASA Technical Reports Server (NTRS)

Woollett, R. R.

1981-01-01

Low speed wind tunnel tests were conducted to assess four methods for attaining thrust modulation for V/STOL aircraft. The four methods were: (1) fan speed change, (2) fan nozzle exit area change, (3) variable pitch rotor (VPR) fan, and (4) variable inlet guide vanes (VIGV). The interrelationships between inlet and thrust modulation system were also investigated using a double slotted inlet and thick lip inlet. Results can be summarized as: (1) the VPR and VIGV systems were the most promising, (2) changes in blade angle to obtain changes in fan thrust have significant implications for the inlet, and (3) both systems attained required level of thrust with acceptable levels of fan blade stress.

SST Technology Follow-On Program - Phase 2. Noise Suppressor/Nozzle Development. Volume 9. Performance Technology - Analysis of the Low Speed Performance of Multitube Suppressor/Ejector Nozzles (0-167 kn)

DTIC Science & Technology

1975-03-01

Layer Suction 18 Temperature and Pressure Profile at Charging Station |9 Roiind-Corivergent Reference Nozzle 20 Elliptical Ramps 21 37-Tube...between plumes of the jets in the outer row of a suppressor Homulary layer Discharge coelticient, accounting for temperature induced no/./Ie area...tunnel floor. The suppressor air tlow rate was measured with an A.S.M.H. long-radius flow nozzle. The boundary layer ihickness at the ejector inlet

Exhaust Nozzle Plume Effects on Sonic Boom Test Results for Vectored Nozzles

NASA Technical Reports Server (NTRS)

Castner, Raymond

2012-01-01

Reducing or eliminating the operational restrictions of supersonic aircraft over populated areas has led to extensive research at NASA. Restrictions were due to the disturbance of the sonic boom, caused by the coalescence of shock waves formed off the aircraft. Recent work has been performed to reduce the magnitude of the sonic boom N-wave generated by airplane components with a focus on shock waves caused by the exhaust nozzle plume. Previous Computational Fluid Dynamics (CFD) analysis showed how the shock wave formed at the nozzle lip interacts with the nozzle boat-tail expansion wave. An experiment was conducted in the 1- by 1-foot Supersonic Wind Tunnel (SWT) at the NASA Glenn Research Center. Results show how the shock generated at the nozzle lip affects the near field pressure signature, and thereby the potential sonic boom contribution for a nozzle at vector angles from 3 to 8 . The experiment was based on the NASA F-15 nozzle used in the Lift and Nozzle Change Effects on Tail Shock experiment, which possessed a large external boat-tail angle. In this case, the large boat-tail angle caused a dramatic expansion, which dominated the near field pressure signature. The impact of nozzle vector angle and nozzle pressure ratio are summarized.

Boundary layer development as a function of chamber pressure in the NASA Lewis 1030:1 area ratio rocket nozzle

NASA Technical Reports Server (NTRS)

Smith, Tamara A.

1988-01-01

Through the use of theoretical predictions of fluid properties and experimental heat transfer and thrust measurements, the zones of laminar, transitional, and turbulent boundary layer flow were defined for the NASA Lewis 1039:1 area ratio rocket nozzle. Tests were performed on the nozzle at chamber pressures from 350 to 100 psia. For these conditions, the throat diameter Reynolds numbers varied from 300,000 to 1 million. The propellants used were gaseous hydrogen and gaseous oxygen. Thrust measurements and nozzle outer wall temperature measurements were taken during the 3-sec test runs. Comparison of experimental heat transfer and thrust data with the corresponding predictions from the Two-Dimensional Kinetics (TDK) nozzle analysis program indicated laminar flow in the nozzle at a throat diameter Reynolds number of 320,000 or chamber pressure of 360 psia. Comparison of experimental and predicted heat transfer data indicated transitional flow up to and including a chamber pressure of 1000 psia. Predicted values of the axisymmetric acceleration parameter within the convergent and divergent nozzle were consistent with the above results. Based upon an extrapolation of the heat transfer data and predicted distributions of the axisymmetric acceleration parameter, transitional flow was predicted up to a throat diameter Reynolds number of 220,000 or 2600-psia chamber pressure. Above 2600-psia chamber pressure, fully developed turbulent flow was predicted.

Boundary layer development as a function of chamber pressure in the NASA Lewis 1030:1 area ratio rocket nozzle

NASA Technical Reports Server (NTRS)

Smith, Tamara A.

1988-01-01

Through the use of theoretical predictions of fluid properties and experimental heat transfer and thrust measurements, the zones of laminar, transitional, and turbulent boundary layer flow were defined for the NASA Lewis 1030:1 area ratio rocket nozzle. Tests were performed on the nozzle at chamber pressures from 350 to 100 psia. For these conditions, the throat diameter Reynolds numbers varied from 300,000 to 1 million. The propellants used were gaseous hydrogen and gaseous oxygen. Thrust measurements and nozzle outer wall temperature measurements were taken during the 3-sec test runs. Comparison of experimental heat transfer and thrust data with the corresponding predictions from the Two-Dimensional Kinetics (TDK) nozzle analysis program indicated laminar flow in the nozzle at a throat diameter Reynolds number of 320,000 or chamber pressure of 360 psia. Comparison of experimental and predicted heat transfer data indicated transitional flow up to and including a chamber pressure of 1000 psia. Predicted values of the axisymmetric acceleration parameter within the convergent and divergent nozzle were consistent with the above results. Based upon an extrapolation of the heat transfer data and predicted distributions of the axisymmetric acceleration parameter, transitional flow was predicted up to a throat diameter Reynolds number of 220,000 or 2600-psia chamber pressure. Above 2600-psia chamber pressure, fully developed turbulent flow was predicted.

X-ray radiography of cavitation in a beryllium alloy nozzle

DOE Office of Scientific and Technical Information (OSTI.GOV)

Duke, Daniel J.; Matusik, Katarzyna E.; Kastengren, Alan L.

In this study, making quantitative measurements of the vapor distribution in a cavitating nozzle is difficult, owing to the strong scattering of visible light at gas–liquid boundaries and wall boundaries, and the small lengths and time scales involved. The transparent models required for optical experiments are also limited in terms of maximum pressure and operating life. Over the past few years, x-ray radiography experiments at Argonne’s Advanced Photon Source have demonstrated the ability to perform quantitative measurements of the line of sight projected vapor fraction in submerged, cavitating plastic nozzles. In this paper, we present the results of new radiographymore » experiments performed on a submerged beryllium nozzle which is 520 μm in diameter, with a length/diameter ratio of 6. Beryllium is a light, hard metal that is very transparent to x-rays due to its low atomic number. We present quantitative measurements of cavitation vapor distribution conducted over a range of non-dimensional cavitation and Reynolds numbers, up to values typical of gasoline and diesel fuel injectors. A novel aspect of this work is the ability to quantitatively measure the area contraction along the nozzle with high spatial resolution. Analysis of the vapor distribution, area contraction and discharge coefficients are made between the beryllium nozzle and plastic nozzles of the same nominal geometry. When gas is dissolved in the fuel, the vapor distribution can be quite different from that found in plastic nozzles of the same dimensions, although the discharge coefficients are unaffected. In the beryllium nozzle, there were substantially fewer machining defects to act as nucleation sites for the precipitation of bubbles from dissolved gases in the fuel, and as such the effect on the vapor distribution was greatly reduced.« less

X-ray radiography of cavitation in a beryllium alloy nozzle

DOE PAGES

Duke, Daniel J.; Matusik, Katarzyna E.; Kastengren, Alan L.; ...

2017-01-17

In this study, making quantitative measurements of the vapor distribution in a cavitating nozzle is difficult, owing to the strong scattering of visible light at gas–liquid boundaries and wall boundaries, and the small lengths and time scales involved. The transparent models required for optical experiments are also limited in terms of maximum pressure and operating life. Over the past few years, x-ray radiography experiments at Argonne’s Advanced Photon Source have demonstrated the ability to perform quantitative measurements of the line of sight projected vapor fraction in submerged, cavitating plastic nozzles. In this paper, we present the results of new radiographymore » experiments performed on a submerged beryllium nozzle which is 520 μm in diameter, with a length/diameter ratio of 6. Beryllium is a light, hard metal that is very transparent to x-rays due to its low atomic number. We present quantitative measurements of cavitation vapor distribution conducted over a range of non-dimensional cavitation and Reynolds numbers, up to values typical of gasoline and diesel fuel injectors. A novel aspect of this work is the ability to quantitatively measure the area contraction along the nozzle with high spatial resolution. Analysis of the vapor distribution, area contraction and discharge coefficients are made between the beryllium nozzle and plastic nozzles of the same nominal geometry. When gas is dissolved in the fuel, the vapor distribution can be quite different from that found in plastic nozzles of the same dimensions, although the discharge coefficients are unaffected. In the beryllium nozzle, there were substantially fewer machining defects to act as nucleation sites for the precipitation of bubbles from dissolved gases in the fuel, and as such the effect on the vapor distribution was greatly reduced.« less

Turbojet-exhaust-nozzle secondary-airflow pumping as an exit control of an inlet-stability bypass system for a Mach 2.5 axisymmetric mixed-compression inlet. [Lewis 10- by 10-ft. supersonic wind tunnel test

NASA Technical Reports Server (NTRS)

Sanders, B. W.

1980-01-01

The throat of a Mach 2.5 inlet that was attached to a turbojet engine was fitted with large, porous bleed areas to provide a stability bypass system that would allow a large, stable airflow range. Exhaust-nozzle, secondary-airflow pumping was used as the exit control for the stability bypass airflow. Propulsion system response and stability bypass performance were obtained for several transient airflow disturbances, both internal and external. Internal airflow disturbances included reductions in overboard bypass airflow, power lever angle, and primary-nozzle area, as well as compressor stall. Nozzle secondary pumping as a stability bypass exit control can provide the inlet with a large stability margin with no adverse effects on propulsion system performance.

Qualitative spectroscopic study of magnetic nozzle flow

NASA Technical Reports Server (NTRS)

Umeki, T.; Turchi, P. J.

1992-01-01

The physics of the magnetic nozzle flow for a 100-kW-level quasi-steady MPD thruster was studied by photographic spectroscopy focusing on the plasma model in the flow and the acceleration mechanism. Spectroscopic visualization for the flow-species analysis indicates that the plasma-exhaust flow dominated by NII species were confined by the magnetic nozzle effect to collimate the flow for the better thruster performance. Inside the nozzle, the plasma flow was found to be in nonhomogeneous collisional-radiative condition. There appears to be a substantial flow acceleration from the magnetic nozzle inlet to the outlet with slight expansion. This suggests that the flow resembles that of constant area supersonic duct flow with cooling.

Performance of Single-Stage Turbine of Mark 25 Torpedo Power Plant with Two Special Nozzles. III; Efficiency with Standard Rotor Blades

NASA Technical Reports Server (NTRS)

Schum, Harold J.; Whitney, Warren J.

1949-01-01

A Mark 25 torpedo power plant modified to operate as a single-stage turbine was investigated to determine the performance with two nozzle designs and a standard first-stage rotor having 0.40-inch blades with a 17O met-air angle. Both nozzles had smaller port cross-sectional areas than those nozzles of similar design, which were previously investigated. The performance of the two nozzles was compared on the basis of blade, rotor, and brake efficiencies as a function of blade-jet speed ratio for pressure ratios of 8, 15 (design), and 20. At pressure ratios of 15 and 20, the blade efficiency obtained with the nozzle having circular passages (K) was higher than that obtained with the nozzle having rectangular passages (J). At a pressure ratio of 8, the efficiencies obtained with the two nozzles were comparable for blade-jet speed ratios of less than 0.260. For blade-jet speed ratios exceeding this value, nozzle K yielded slightly higher efficiencies. The maximum blade efficiency of 0.569 was obtained with nozzle K at a pressure ratio of 8 and a blade-jet speed ratio of 0.295. At design speed and pressure ratio, nozzle K yielded a maximum blade efficiency of 0.534, an increase of 0.031 over that obtained with nozzle J. When the blade efficiencies of the two nozzles were compared with those of four other nozzles previously investigated, the maximum difference for the six nozzles with this rotor was 0.050. From, this comparison, no specific effect of nozzles size or shape on over-all performance was discernible.

Aerodynamic Shape Optimization of a Dual-Stream Supersonic Plug Nozzle

NASA Technical Reports Server (NTRS)

Heath, Christopher M.; Gray, Justin S.; Park, Michael A.; Nielsen, Eric J.; Carlson, Jan-Renee

2015-01-01

Aerodynamic shape optimization was performed on an isolated axisymmetric plug nozzle sized for a supersonic business jet. The dual-stream concept was tailored to attenuate nearfield pressure disturbances without compromising nozzle performance. Adjoint-based anisotropic mesh refinement was applied to resolve nearfield compression and expansion features in the baseline viscous grid. Deformed versions of the adapted grid were used for subsequent adjoint-driven shape optimization. For design, a nonlinear gradient-based optimizer was coupled to the discrete adjoint formulation of the Reynolds-averaged Navier- Stokes equations. All nozzle surfaces were parameterized using 3rd order B-spline interpolants and perturbed axisymmetrically via free-form deformation. Geometry deformations were performed using 20 design variables shared between the outer cowl, shroud and centerbody nozzle surfaces. Interior volume grid deformation during design was accomplished using linear elastic mesh morphing. The nozzle optimization was performed at a design cruise speed of Mach 1.6, assuming core and bypass pressure ratios of 6.19 and 3.24, respectively. Ambient flight conditions at design were commensurate with 45,000-ft standard day atmosphere.

Free jet feasibility study of a thermal acoustic shield concept for AST/VCE application: Dual stream nozzles

NASA Technical Reports Server (NTRS)

Janardan, B. A.; Brausch, J. F.; Majjigi, R. K.

1985-01-01

The influence of selected geometric and aerodynamic flow variables of an unsuppressed coannular plug nozzle and a coannular plug nozzle with a 20-chute outer stream suppressor were experimentally determined. A total of 136 static and simulated flight acoustic test points were conducted with 9 scale model nozzles. Also, aerodynamic measurements of four selected plumes were made with a laser velocimeter. The presence of the 180 deg shield produced different mixing characteristics on the shield side compared to the unshield side because of the reduced mixing with ambient air on the shielded side. This resulted in a stretching of the jet, yielding a higher peak mean velocity up to a length of 10 equivalent diameters from the nozzle exit. The 180 deg shield in community orientation around the suppressed coannular plug nozzle yielded acoustic benefit at all observer angles for a simulated takeoff. While the effect of shield-to-outer stream velocity ratio was small at angles up to 120 deg, beyond this angle significant acoustic benefit was realized with a shield-to-outer stream velocity ratio of 0.64.

Design Enhancements of the Two-Dimensional, Dual Throat Fluidic Thrust Vectoring Nozzle Concept

NASA Technical Reports Server (NTRS)

Flamm, Jeffrey D.; Deere, Karen A.; Mason, Mary L.; Berrier, Bobby L.; Johnson, Stuart K.

2006-01-01

A Dual Throat Nozzle fluidic thrust vectoring technique that achieves higher thrust-vectoring efficiencies than other fluidic techniques, without sacrificing thrust efficiency has been developed at NASA Langley Research Center. The nozzle concept was designed with the aid of the structured-grid, Reynolds-averaged Navier-Stokes computational fluidic dynamics code PAB3D. This new concept combines the thrust efficiency of sonic-plane skewing with increased thrust-vectoring efficiencies obtained by maximizing pressure differentials in a separated cavity located downstream of the nozzle throat. By injecting secondary flow asymmetrically at the upstream minimum area, a new aerodynamic minimum area is formed downstream of the geometric minimum and the sonic line is skewed, thus vectoring the exhaust flow. The nozzle was tested in the NASA Langley Research Center Jet Exit Test Facility. Internal nozzle performance characteristics were defined for nozzle pressure ratios up to 10, with a range of secondary injection flow rates up to 10 percent of the primary flow rate. Most of the data included in this paper shows the effect of secondary injection rate at a nozzle pressure ratio of 4. The effects of modifying cavity divergence angle, convergence angle and cavity shape on internal nozzle performance were investigated, as were effects of injection geometry, hole or slot. In agreement with computationally predicted data, experimental data verified that decreasing cavity divergence angle had a negative impact and increasing cavity convergence angle had a positive impact on thrust vector angle and thrust efficiency. A curved cavity apex provided improved thrust ratios at some injection rates. However, overall nozzle performance suffered with no secondary injection. Injection holes were more efficient than the injection slot over the range of injection rates, but the slot generated larger thrust vector angles for injection rates less than 4 percent of the primary flow rate.

Use of ILTV Control Laws for LaNCETS Flight Research

NASA Technical Reports Server (NTRS)

Moua, Cheng

2010-01-01

A report discusses the Lift and Nozzle Change Effects on Tail Shock (LaNCETS) test to investigate the effects of lift distribution and nozzle-area ratio changes on tail shock strength of an F-15 aircraft. Specific research objectives are to obtain inflight shock strength for multiple combinations of nozzle-area ratio and lift distribution; compare results with preflight prediction tools; and update predictive tools with flight results. The objectives from a stability and control perspective are to ensure adequate aircraft stability for the changes in lift distribution and plume shape, and ensure manageable transient from engaging and disengaging the ILTV research control laws. In order to change the lift distribution and plume shape of the F-15 aircraft, a decade-old Inner Loop Thrust Vectoring (ILTV) research control law was used. Flight envelope expansion was performed for the test configuration and flight conditions prior to the probing test points. The approach for achieving the research objectives was to utilize the unique capabilities of NASA's NF-15B-837 aircraft to allow the adjustment of the nozzle-area ratio and/or canard positions by engaging the ILTV research control laws. The ILTV control laws provide the ability to add trim command biases to canard positions, nozzle area ratios, and thrust vectoring through the use of datasets. Datasets consist of programmed test inputs (PTIs) that define trims to change the nozzle-area ratio and/or canard positions. The trims are applied as increments to the normally commanded positions. A LaNCETS non-linear, six-degrees-of-freedom simulation capable of realtime pilot-in-the-loop, hardware-in-the-loop, and non-real-time batch support was developed and validated. Prior to first flight, extensive simulation analyses were performed to show adequate stability margins with the changes in lift distribution and plume shape. Additionally, engagement/disengagement transient analysis was also performed to show manageable transients.

Static Performance of a Fixed-Geometry Exhaust Nozzle Incorporating Porous Cavities for Shock-Boundary Layer Interaction Control

NASA Technical Reports Server (NTRS)

Asbury, Scott C.; Hunter, Craig A.

1999-01-01

An investigation was conducted in the model preparation area of the Langley 16-Foot Transonic Tunnel to determine the internal performance of a fixed-geometry exhaust nozzle incorporating porous cavities for shock-boundary layer interaction control. Testing was conducted at static conditions using a sub-scale nozzle model with one baseline and 27 porous configurations. For the porous configurations, the effects of percent open porosity, hole diameter, and cavity depth were determined. All tests were conducted with no external flow at nozzle pressure ratios from 1.25 to approximately 9.50. Results indicate that baseline nozzle performance was dominated by unstable, shock-induced, boundary-layer separation at over-expanded conditions. Porous configurations were capable of controlling off-design separation in the nozzle by either alleviating separation or encouraging stable separation of the exhaust flow. The ability of the porous nozzle concept to alternately alleviate separation or encourage stable separation of exhaust flow through shock-boundary layer interaction control offers tremendous off-design performance benefits for fixed-geometry nozzle installations. In addition, the ability to encourage separation on one divergent flap while alleviating it on the other makes it possible to generate thrust vectoring using a fixed-geometry nozzle.

Variable mixture ratio performance through nitrogen augmentation

NASA Technical Reports Server (NTRS)

Beichel, R.; Obrien, C. J.; Bair, E. K.

1988-01-01

High/variable mixture ratio O2/H2 candidate engine cycles are examined for earth-to-orbit vehicle application. Engine performance and power balance information are presented for the candidate cycles relative to chamber pressure, bulk density, and mixture ratio. Included in the cycle screening are concepts where a third fluid (liquid nitrogen) is used to achieve a variable mixture ratio over the trajectory from liftoff to earth orbit. The third fluid cycles offer a very low risk, fully reusable, low operation cost alternative to high/variable mixture ratio bipropellant cycles. Variable mixture ratio engines with extendible nozzle are slightly lower performing than a single mixture ratio engine (MR = 7:1) with extendible nozzle. Dual expander engines (MR = 7:1) have slightly better performance than the single mixture ratio engine. Dual fuel dual expander engines offer a 16 percent improvement over the single mixture ratio engine.

Investigation of Altitude Starting and Acceleration Characteristics of J47 Turbojet Engine

NASA Technical Reports Server (NTRS)

Golladay, Richard L; Bloomer, Harry E

1951-01-01

An investigation was conducted on an axial-flow-compressor type turbojet engine in the NACA Lewis altitude wind tunnel to determine the operational characteristics of several ignition systems, cross-fire tube configurations and fuel systems over a range of simulated flight conditions. The opposite-polarity-type spark plug provided the most satisfactory ignition. Increasing the cross-fire-tube diameter improved intercombustor flame propagation. At high windmilling speeds, accelerations to approximately 6200 rpm could be made at a preset constant throttle position. The use of a variable-area nozzle reduced acceleration time.

Low-noise nozzle valve

NASA Technical Reports Server (NTRS)

Gwin, Hal S. (Inventor); Aaron, James (Inventor)

1990-01-01

A low noise, variable discharage area, valve is constructed having opposed recesses within which a pair of gates are slidably disposed. Each of the gates is provided with upstream edges having a radius thereon, the radius enabling smooth, accelerated, low noise flow therebetween. The gates are further provided with tracks along each side, which in turn slide along splines set in the side walls of the valve. A threaded rod which rotates in a threaded insert in a rear wall of each of the gates, serves to move the gates within their respective recesses.

Evaluation of fuel preparation systems for lean premixing-prevaporizing combustors

NASA Technical Reports Server (NTRS)

Dodds, W. J.; Ekstedt, E. E.

1985-01-01

A series of experiments was carried out in order to produce design data for a premixing prevaporizing fuel-air mixture preparation system for aircraft gas turbine engine combustors. The fuel-air mixture uniformity of four different system design concepts was evaluated over a range of conditions representing the cruise operation of a modern commercial turbofan engine. Operating conditions including pressure, temperature, fuel-to-air ratio, and velocity, exhibited no clear effect on mixture uniformity of systems using pressure-atomizing fuel nozzles and large-scale mixing devices. However, the performance of systems using atomizing fuel nozzles and large-scale mixing devices was found to be sensitive to operating conditions. Variations in system design variables were also evaluated and correlated. Mixing uniformity was found to improve with system length, pressure drop, and the number of fuel injection points per unit area. A premixing system capable of providing mixing uniformity to within 15 percent over a typical range of cruise operating conditions is demonstrated.

Effects of Convoluted Divergent Flap Contouring on the Performance of a Fixed-Geometry Nonaxisymmetric Exhaust Nozzle

NASA Technical Reports Server (NTRS)

Asbury, Scott C.; Hunter, Craig A.

1999-01-01

An investigation was conducted in the model preparation area of the Langley 16-Foot Transonic Tunnel to determine the effects of convoluted divergent-flap contouring on the internal performance of a fixed-geometry, nonaxisymmetric, convergent-divergent exhaust nozzle. Testing was conducted at static conditions using a sub-scale nozzle model with one baseline and four convoluted configurations. All tests were conducted with no external flow at nozzle pressure ratios from 1.25 to approximately 9.50. Results indicate that baseline nozzle performance was dominated by unstable, shock-induced, boundary-layer separation at overexpanded conditions. Convoluted configurations were found to significantly reduce, and in some cases totally alleviate separation at overexpanded conditions. This result was attributed to the ability of convoluted contouring to energize and improve the condition of the nozzle boundary layer. Separation alleviation offers potential for installed nozzle aeropropulsive (thrust-minus-drag) performance benefits by reducing drag at forward flight speeds, even though this may reduce nozzle thrust ratio as much as 6.4% at off-design conditions. At on-design conditions, nozzle thrust ratio for the convoluted configurations ranged from 1% to 2.9% below the baseline configuration; this was a result of increased skin friction and oblique shock losses inside the nozzle.

Free-jet investigation of mechanically suppressed, high radius ratio coannular plug model nozzles

NASA Technical Reports Server (NTRS)

Janardan, B. A.; Majjigi, R. K.; Brausch, J. F.; Knott, P. R.

1985-01-01

The experimental and analytical acoustic results of a scale-model investigation or unsuppressed and mechanically suppressed high-radius ratio coannular plug nozzles with inverted velocity and temperature profiles are summarized. Nine coannular nozzle configurations along with a reference conical nozzle were evaluated in the Anechoic Free-Jet Facility for a total of 212 acoustic test points. Most of the tests were conducted at variable cycle engine conditions applicable to advanced high speed aircraft. The tested nozzles included coannular plug nozzles with both convergent and convergent-divergent (C-D) terminations in order to evaluate C-D effectiveness in the reduction of shock-cell noise and 20 and 40 shallow-chute mechanical suppressors in the outer stream in order to evaluate their effectiveness in the reduction of jet noise. In addition to the acoustic tests, mean and turbulent velocity measurements were made on selected plumes of the 20 shallow-chute configuration using a laser velocimeter. At a mixed jet velocity of 700 m/sec, the 20 shallow-chute suppressor configuration yielded peak aft quadrant suppression of 11.5 and 9 PNdB and forward quadrant suppression of 7 and 6 PNdB relative to a baseline conical nozzles during static and simulated flight, respectively. The C-D terminations were observed to reduce shock-cell noise. An engineering spectral prediction method was formulated for mechanically suppressed coannular plug nozzles.

Gas turbine exhaust nozzle. [for noise reduction

NASA Technical Reports Server (NTRS)

Straight, D. M. (Inventor)

1973-01-01

An elongated hollow string is disposed in an exhaust nozzle combustion chamber and communicates with an air source through hollow struts at one end. The other end of the string is bell-mouth shaped and extends over the front portion of a nozzle plug. The bell-mouth may be formed by pivotally mounted flaps or leaves which are used to vary the exhaust throat area and the area between the plug and the leaves. Air from the engine inlet flows into the string and also between the combustion chamber and a housing disposed around the chamber. The air cools the plug and serves as a low velocity inner core of secondary gas to provide noise reduction for the primary exhaust gas while the other air, when it exits from the nozzle, forms an outer low velocity layer to further reduce noise. The structure produces increased thrust in a turbojet or turbofan engine.

Effect of varying internal geometry on the static performance of rectangular thrust-reverser ports

NASA Technical Reports Server (NTRS)

Re, Richard J.; Mason, Mary L.

1987-01-01

An investigation has been conducted to evaluate the effects of several geometric parameters on the internal performance of rectangular thrust-reverser ports for nonaxisymmetric nozzles. Internal geometry was varied with a test apparatus which simulated a forward-flight nozzle with a single, fully deployed reverser port. The test apparatus was designed to simulate thrust reversal (conceptually) either in the convergent section of the nozzle or in the constant-area duct just upstream of the nozzle. The main geometric parameters investigated were port angle, port corner radius, port location, and internal flow blocker angle. For all reverser port geometries, the port opening had an aspect ratio (throat width to throat height) of 6.1 and had a constant passage area from the geometric port throat to the exit. Reverser-port internal performance and thrust-vector angles computed from force-balance measurements are presented.

1998 Calibration of the Mach 4.7 and Mach 6 Arc-Heated Scramjet Test Facility Nozzles

NASA Technical Reports Server (NTRS)

Witte, David W.; Irby, Richard G.; Auslender, Aaron H.; Rock, Kenneth E.

2004-01-01

A calibration of the Arc-Heated Scramjet Test Facility (AHSTF) Mach 4.7 and Mach 6 nozzles was performed in 1998. For each nozzle, three different typical facility operating test points were selected for calibration. Each survey consisted of measurements, at 340 separate locations across the 11 inch square nozzle exit plane, of pitot pressure, static pressure, and total temperature. Measurement density was higher (4/inch) in the boundary layer near the nozzle wall than in the core nozzle flow (1/inch). The results generated for each of these calibration surveys were contour plots at the nozzle exit plane of the measured and calculated flow properties which completely defined the thermodynamic state of the nozzle exit flow. An area integration of the mass flux at the nozzle exit for each survey was compared to the AHSTF mass flow meter results to provide an indication of the overall quality of the calibration performed. The percent difference between the integrated nozzle exit mass flow and the flow meter ranged from 0.0 to 1.3 percent for the six surveys. Finally, a comparison of this 1998 calibration was made with the 1986 calibration. Differences of less than 10 percent were found within the nozzle core flow while in the boundary layer differences on the order of 20 percent were quite common.

Design and Checkout of a High Speed Research Nozzle Evaluation Rig

NASA Technical Reports Server (NTRS)

Castner, Raymond S.; Wolter, John D.

1997-01-01

The High Flow Jet Exit Rig (HFJER) was designed to provide simulated mixed flow turbojet engine exhaust for one- seventh scale models of advanced High Speed Research test nozzles. The new rig was designed to be used at NASA Lewis Research Center in the Nozzle Acoustic Test Rig and the 8x6 Supersonic Wind Tunnel. Capabilities were also designed to collect nozzle thrust measurement, aerodynamic measurements, and acoustic measurements when installed at the Nozzle Acoustic Test Rig. Simulated engine exhaust can be supplied from a high pressure air source at 33 pounds of air per second at 530 degrees Rankine and nozzle pressure ratios of 4.0. In addition, a combustion unit was designed from a J-58 aircraft engine burner to provide 20 pounds of air per second at 2000 degrees Rankine, also at nozzle pressure ratios of 4.0. These airflow capacities were designed to test High Speed Research nozzles with exhaust areas from eighteen square inches to twenty-two square inches. Nozzle inlet flow measurement is available through pressure and temperature sensors installed in the rig. Research instrumentation on High Speed Research nozzles is available with a maximum of 200 individual pressure and 100 individual temperature measurements. Checkout testing was performed in May 1997 with a 22 square inch ASME long radius flow nozzle. Checkout test results will be summarized and compared to the stated design goals.

Static internal performance of convergent single-expansion-ramp nozzles with various combinations of internal geometric parameters

NASA Technical Reports Server (NTRS)

Bare, E. Ann; Capone, Francis J.

1989-01-01

An investigation was conducted in the Static Test Facility of the Langley 16-Foot Transonic Tunnel to determine the effects of five geometric design parameters on the internal performance of convergent single expansion ramp nozzles. The effects of ramp chordal angle, initial ramp angle, flap angle, flap length, and ramp length were determined. All nozzles tested has a nominally constant throat area and aspect ratio. Static pressure distributions along the centerlines of the ramp and flap were also obtained for each configuration. Nozzle pressure ratio was varied up to 10.0 for all configurations.

Plug cluster engine concept for in-space missions

NASA Technical Reports Server (NTRS)

Obrien, C. J.; Aukerman, C. A.

1979-01-01

The development of a suitable orbital transfer vehicle (OTV) engine is discussed. The OTV's dimensions are limited by those of the Space Shuttle payload bay on which it will be carried. An approach to utilize the available diameter to achieve high area ratio and thus high engine performance, is presented. Unconventional nozzles, such as clusters of small thrusters around a large diameter contoured plug, are investigated to arrive at engine designs which feature lower chamber pressures, with attendant lower heat flux, lower wall temperature, longer fatigue life, and less critical turbomachinery. Attention is also given to plug nozzle technology, high area ratio module- and scarfed bell- Plug Cluster Engine (PCE) concepts, as well as PCE performance, weight, and assessment. A conceptual design of a PCE formed from a cluster of high area ratio, scarfed, bell nozzles proved to be competitive with bell and spike nozzle engines. PCE advantages cited include increased payload length due to shorter engine length, ability to increase or decrease the number of modules and thereby the thrust, and low cost due to utilization of off-the-shelf technology.

High Pressure Water Stripping Using Multi-Orifice Nozzles

NASA Technical Reports Server (NTRS)

Hoppe, David T.

1998-01-01

The use of multi-orifice rotary nozzles not only increases the speed and stripping effectiveness of high pressure water blasting systems, but also greatly increases the complexity of selecting and optimizing the operating parameters. The rotational speed of the nozzle must be coupled with the transverse velocity of the nozzle as it passes across the surface of the substrate being stripped. The radial and angular positions of each orifice must be included in the analysis of the nozzle configuration. Since orifices at the outer edge of the nozzle head move at a faster rate than the orifice located near the center, the energy impact force of the water stream from the outer orifice is spread over a larger area than the water streams from the inner orifice. Utilizing a larger diameter orifice in the outer radial positions increases the energy impact to compensate for its wider force distribution. The total flow rate from the combination of orifices must be monitored and kept below the pump capacity while choosing an orifice to insert in each position. The energy distribution from the orifice pattern is further complicated since the rotary path of all orifices in the nozzle head pass through the center section, contributing to the stripping in this area while only the outer most orifice contributes to the stripping in the shell area at the extreme outside edge of the nozzle. From t he outer most shell to the center section, more orifices contribute to the stripping in each progressively reduced diameter shell. With all these parameters to configure and each parameter change affecting the others, a computer model was developed to track and coordinate these parameters. The computer simulation responds by graphically indicating the cumulative affect from each parameter selected. The results from the proper choices in parameters is a well designed, highly efficient stripping system. A poorly chosen set of parameters will cause the nozzle to strip aggressively in some areas while leaving the coating untouched in adjacent sections. The high pressure water stripping system can be set to extremely aggressive conditions allowing stripping of hard to remove adhesives, paint systems, cladding and chromate conversion coatings. The energy force can be reduced to strip coatings from thin aluminum substrates without causing damage or deterioration to the substrate's surface. High pressure water stripping of aerospace components have thus proven to be an efficient and cost effective method for cleaning and removing coatings.

Variable volume combustor with a conical liner support

DOE Office of Scientific and Technical Information (OSTI.GOV)

Johnson, Thomas Edward; McConnaughhay, Johnie Franklin; Keener, Chrisophter Paul

The present application provides a variable volume combustor for use with a gas turbine engine. The variable volume combustor may include a liner, a number of micro-mixer fuel nozzles positioned within the liner, and a conical liner support supporting the liner.

Hydrogen no-vent fill testing in a 5 cubic foot (142 liter) tank using spray nozzle and spray bar liquid injection

NASA Technical Reports Server (NTRS)

Moran, Matthew E.; Nyland, Ted W.

1992-01-01

A total of 38 hydrogen no-vent fill tests were performed in this test series using various size spray nozzles and a spray bar with different hole sizes in a 5 cubic foot receiver tank. Fill levels of 90 percent by volume or greater were achieved in 26 of the tests while maintaining a receiver tank pressure below 30 psia. Spray nozzles were mounted at the top of the tank, whereas, the spray bar was centered in the tank axially. The spray nozzle no-vent fills demonstrated tank pressure and temperature responses comparable to previous test series. Receiver tank pressure responses for the spray bar configuration were similar to the spray nozzle tests with the pressure initially rising rapidly, then leveling off as vapor condenses onto the discharging liquid streams, and finally ramping up near the end of the test due to ullage compression. Both liquid injection techniques tested were capable of filling the receiver tank to 90 percent under variable test conditions. Comparisons between the spray nozzle and spray bar configurations for well matched test conditions indicate the spray nozzle injection technique is more effective in minimizing the receiving tank pressure throughout a no-vent fill compared to the spray bar under normal gravity conditions.

V/STOL Tandem Fan transition section model test. [in the Lewis Research Center 10-by-10 foot wind tunnel

NASA Technical Reports Server (NTRS)

Simpkin, W. E.

1982-01-01

An approximately 0.25 scale model of the transition section of a tandem fan variable cycle engine nacelle was tested in the NASA Lewis Research Center 10-by-10 foot wind tunnel. Two 12-inch, tip-turbine driven fans were used to simulate a tandem fan engine. Three testing modes simulated a V/STOL tandem fan airplane. Parallel mode has two separate propulsion streams for maximum low speed performance. A front inlet, fan, and downward vectorable nozzle forms one stream. An auxilliary top inlet provides air to the aft fan - supplying the core engine and aft vectorable nozzle. Front nozzle and top inlet closure, and removal of a blocker door separating the two streams configures the tandem fan for series mode operations as a typical aircraft propulsion system. Transition mode operation is formed by intermediate settings of the front nozzle, blocker door, and top inlet. Emphasis was on the total pressure recovery and flow distortion at the aft fan face. A range of fan flow rates were tested at tunnel airspeeds from 0 to 240 knots, and angles-of-attack from -10 to 40 deg for all three modes. In addition to the model variables for the three modes, model variants of the top inlet were tested in the parallel mode only. These lip variables were: aft lip boundary layer bleed holes, and Three position turning vane. Also a bellmouth extension of the top inlet side lips was tested in parallel mode.

Directed spray mast

DOEpatents

Nance, Thomas A.; Siddall, Alvin A.; Cheng, William Y.; Counts, Kevin T.

2005-05-10

Disclosed is an elongated, tubular, compact high pressure sprayer apparatus for insertion into an access port of vessels having contaminated interior areas that require cleaning by high pressure water spray. The invention includes a spray nozzle and a camera adjacent thereto with means for rotating and raising and lowering the nozzle so that areas identified through the camera may be cleaned with a minimum production of waste water to be removed.

A Comparison of Shadowgraphy and X-ray Computed Tomography in Liquid Spray Analysis

DTIC Science & Technology

2014-11-14

atomizers and downstream of the nozzle exit gives insight into optimizing atomizers, particularly for combustion applications. The performance of gas ...regions near the spray nozzle [9, 10]. Because light refraction by liquid sheets is significant, these areas all cast a full shadow on the camera...hollow-cone pressure swirl design. Within this nozzle design, liquid swirls around an air-cored vortex. Upon exiting, the fluid expands due to its

Parametric Study of Sealant Nozzle

NASA Astrophysics Data System (ADS)

Yamamoto, Yoshimi

It has become apparent in recent years the advancement of manufacturing processes in the aerospace industry. Sealant nozzles are a critical device in the use of fuel tank applications for optimal bonds and for ground service support and repair. Sealants has always been a challenging area for optimizing and understanding the flow patterns. A parametric study was conducted to better understand geometric effects of sealant flow and to determine whether the sealant rheology can be numerically modeled. The Star-CCM+ software was used to successfully develop the parametric model, material model, physics continua, and simulate the fluid flow for the sealant nozzle. The simulation results of Semco sealant nozzles showed the geometric effects of fluid flow patterns and the influences from conical area reduction, tip length, inlet diameter, and tip angle parameters. A smaller outlet diameter induced maximum outlet velocity at the exit, and contributed to a high pressure drop. The conical area reduction, tip angle and inlet diameter contributed most to viscosity variation phenomenon. Developing and simulating 2 different flow models (Segregated Flow and Viscous Flow) proved that both can be used to obtain comparable velocity and pressure drop results, however; differences are seen visually in the non-uniformity of the velocity and viscosity fields for the Viscous Flow Model (VFM). A comprehensive simulation setup for sealant nozzles was developed so other analysts can utilize the data.

Comparison of options for reduction of noise in the test section of the NASA Langley 4x7m wind tunnel, including reduction of nozzle area

NASA Technical Reports Server (NTRS)

Hayden, R. E.

1984-01-01

The acoustically significant features of the NASA 4X7m wind tunnel and the Dutch-German DNW low speed tunnel are compared to illustrate the reasons for large differences in background noise in the open jet test sections of the two tunnels. Also introduced is the concept of reducing test section noise levels through fan and turning vane source reductions which can be brought about by reducing the nozzle cross sectional area, and thus the circuit mass flow for a particular exit velocity. The costs and benefits of treating sources, paths, and changing nozzle geometry are reviewed.

Free-jet feasibility study of a thermal acoustic shield concept for AST/VCE application-dual stream nozzles. Comprehensive data report. Volume 2: Laser velocimeter and suppressor. Base pressure data

NASA Technical Reports Server (NTRS)

Janardan, B. A.; Brausch, J. F.; Price, A. O.

1984-01-01

Acoustic and diagnostic data that were obtained to determine the influence of selected geometric and aerodynamic flow variables of coannular nozzles with thermal acoustic shields are summarized in this comprehensive data report. A total of 136 static and simulated flight acoustic test points were conducted with 9 scale-model nozzles. Aerodynamic laser velocimeter measurements were made for four selected plumes. In addition, static pressure data in the chute base region of the suppressor configurations were obtained to assess the influence of the shield stream on the suppressor base drag.

Investigation of impingement region and wall jets formed by the interaction of high aspect ratio lift jets and a ground plane

NASA Technical Reports Server (NTRS)

Kotansky, D. R.; Glaze, L. W.

1978-01-01

Flow characteristics of impinging jets emanating from rectangular exit area converging nozzles of exit area aspect ratio four, six, and eight were investigated. Azimuthal distributions of wall jet radial momentum flux in the ground plane were strongly directional and sensitive to rectangular nozzle exit area aspect ratio, jet impingement angle, and height above ground, H/D. Effects of jet exit velocity profile nonuniformities were also investigated. Data from the single nozzle rectangular jet impringement investigations were incorporated into an existing VTOL aircraft ground flow field computer program. It is suggested that this program together with the Douglas Neumann program modified for V/STOL applications may be used for the analysis and prediction of flow fields and resulting forces and moments on multijet V/STOL aircraft hovering in ground effect.

Phased-Array Study of Dual-Flow Jet Noise: Effect of Nozzles and Mixers

NASA Technical Reports Server (NTRS)

Soo Lee, Sang; Bridges, James

2006-01-01

A 16-microphone linear phased-array installed parallel to the jet axis and a 32-microphone azimuthal phased-array installed in the nozzle exit plane have been applied to identify the noise source distributions of nozzle exhaust systems with various internal mixers (lobed and axisymmetric) and nozzles (three different lengths). Measurements of velocity were also obtained using cross-stream stereo particle image velocimetry (PIV). Among the three nozzle lengths tested, the medium length nozzle was the quietest for all mixers at high frequency on the highest speed flow condition. Large differences in source strength distributions between nozzles and mixers occurred at or near the nozzle exit for this flow condition. The beamforming analyses from the azimuthal array for the 12-lobed mixer on the highest flow condition showed that the core flow and the lobe area were strong noise sources for the long and short nozzles. The 12 noisy spots associated with the lobe locations of the 12-lobed mixer with the long nozzle were very well detected for the frequencies 5 KHz and higher. Meanwhile, maps of the source strength of the axisymmetric splitter show that the outer shear layer was the most important noise source at most flow conditions. In general, there was a good correlation between the high turbulence regions from the PIV tests and the high noise source regions from the phased-array measurements.

The influences of tip clearance on the performance of nozzle blades of radial turbines - Experiment and performance prediction at three nozzle angles

NASA Astrophysics Data System (ADS)

Hyun, Yong-Ik; Yamaguchi, Michiteru; Hayami, Hiroshi; Senoo, Yasutoshi

1988-05-01

In order to study the influence of tip clearance on the turning angle and pressure loss of turbine nozzles, experimental results were obtained for nozzle angles at which the throat area was 0.8 and 1.4 times the rated condition. Contour maps of the total pressure loss and of the spanwise distributions of the mean exit-flow angle have been obtained. Although the two-layer flow model of Senoo et al., (1987) is shown to accurately predict the effects of tip clearance, it underestimates the clearance effect for a lightly loaded condition.

Control of supersonic wind-tunnel noise by laminarization of nozzle-wall boundary layer

NASA Technical Reports Server (NTRS)

Beckwith, I. E.; Harvey, W. D.; Harris, J. E.; Holley, B. B.

1973-01-01

One of the principal design requirements for a quiet supersonic or hypersonic wind tunnel is to maintain laminar boundary layers on the nozzle walls and thereby reduce disturbance levels in the test flow. The conditions and apparent reasons for laminar boundary layers which have been observed during previous investigations on the walls of several nozzles for exit Mach numbers from 2 to 20 are reviewed. Based on these results, an analysis and an assessment of nozzle design requirements for laminar boundary layers including low Reynolds numbers, high acceleration, suction slots, wall temperature control, wall roughness, and area suction are presented.

A preliminary assessment of the impact of 2-D exhaust-nozzle geometry on the cruise range of a hypersonic aircraft with top-mounted ramjet propulsion

NASA Technical Reports Server (NTRS)

Vahl, W. A.; Weidner, J. P.

1980-01-01

A theoretical study of full length and shortened, two dimensional, isentropic, exhaust nozzles integrated with top mounted ramjet propulsion nacelles were conducted. Both symmetric and asymmetric contoured nozzles with a range of angular orientations were considered. Performance comparisons to determine optimum installations for a representative hypersonic vehicle at Mach 5 cruise conditions are presented on the basis of cruise range, propulsive specific impulse, inlet area requirements, and overall lift drag ratio. The effect of approximating the nozzle internal contours with planar surfaces and the determination of viscous and frozen flow effects are also presented.

Comparison of Experimental Data and Computations Fluid Dynamics Analysis for a Three Dimensional Linear Plug Nozzle

NASA Technical Reports Server (NTRS)

Ruf, J. H.; Hagemann, G.; Immich, H.

2003-01-01

A three dimensional linear plug nozzle of area ratio 12.79 was designed by EADS Space Transportation (former Astrium Space Infrastructure). The nozzle was tested within the German National Technology Program 'LION' in a cold air wind tunnel by TU Dresden. The experimental hardware and test conditions are described. Experimental data was obtained for the nozzle without plug side wall fences at a nozzle pressure ratio of 116 and then with plug side wall fences at NPR 110. Schlieren images were recorded and axial profiles of plug wall static pressures were measured at several spanwise locations and on the plug base. Detailed CFD analysis was performed for these nozzle configurations at NPR 116 by NASA MSFC. The CFD exhibits good agreement with the experimental data. A detailed comparison of the CFD results and the experimental plug wall pressure data are given. Comparisons are made for both the without and with plug side wall fence configurations. Numerical results for density gradient are compared to experimental Schlieren images. Experimental nozzle thrust efficiencies are calculated based on the CFD results. The CFD results are used to illustrate the plug nozzle fluid dynamics. The effect of the plug side wall is emphasized.

Analysis of Exhaust Plume Effects on Sonic Boom for a 59-Degree Wing Body Model

NASA Technical Reports Server (NTRS)

Castner, Raymond S.

2011-01-01

Reducing or eliminating the operational restrictions of supersonic aircraft over populated areas has led to extensive research at NASA. Restrictions are due to the disturbance of the sonic boom, caused by the coalescence of shock waves formed off the aircraft. Recent work has been performed to reduce the magnitude of the sonic boom N-wave generated by airplane components with focus on shock waves caused by the exhaust nozzle plume. Previous Computational Fluid Dynamics (CFD) analyses showed how the shock wave formed at the nozzle lip interacted with the nozzle boat-tail expansion wave. The nozzle lip shock moved with increasing nozzle pressure ratio (NPR) and reduced the nozzle boat-tail expansion. Lip shock movement caused a favorable change in the observed pressure signature. These results were applied to a simplified supersonic vehicle geometry with no inlets and no tail, in which the goal was to demonstrate how under-expanded nozzle operation reduced the sonic boom signature by twelve percent. A secondary goal was to demonstrate the use of the Cart3D inviscid code for off-body pressure signatures including the nozzle plume effect.

SST Technology Follow-On Program-Phase 2, Noise Suppressor/Nozzle Development. Volume 2. Noise Technology

DTIC Science & Technology

1975-03-01

Loss Relationships 199 109 37-Tube, 4.5 Area Ratio Nozzle, Premergcd Jet Turbulence Noise 200 110 37-Tube Nozzle Premerged Jet Noise Peak...were obtained with the tunnel oil and at 165 knots. The tunnel air flows through a large , rectangular bell-mouth inlet, a (low straightening grid... ratio conditions on a fourteen-track annlog tape recorder for subsecjuent analysis after test com- pletion. Basic analysis of the recorded acoustic

Thrust Augmented Nozzle for a Hybrid Rocket with a Helical Fuel Port

NASA Astrophysics Data System (ADS)

Marshall, Joel H.

A thrust augmented nozzle for hybrid rocket systems is investigated. The design lever-ages 3-D additive manufacturing to embed a helical fuel port into the thrust chamber of a hybrid rocket burning gaseous oxygen and ABS plastic as propellants. The helical port significantly increases how quickly the fuel burns, resulting in a fuel-rich exhaust exiting the nozzle. When a secondary gaseous oxygen flow is injected into the nozzle downstream of the throat, all of the remaining unburned fuel in the plume spontaneously ignites. This secondary reaction produces additional high pressure gases that are captured by the nozzle and significantly increases the motor's performance. Secondary injection and combustion allows a high expansion ratio (area of the nozzle exit divided by area of the throat) to be effective at low altitudes where there would normally be significantly flow separation and possibly an embedded shock wave due. The result is a 15 percent increase in produced thrust level with no loss in engine efficiency due to secondary injection. Core flow efficiency was increased significantly. Control tests performed using cylindrical fuel ports with secondary injection, and helical fuel ports without secondary injection did not exhibit this performance increase. Clearly, both the fuel-rich plume and secondary injection are essential features allowing the hybrid thrust augmentation to occur. Techniques for better design optimization are discussed.

Numerical Simulations of Noise Generated by High Aspect Ratio Supersonic Rectangular Jets - Validation

NASA Astrophysics Data System (ADS)

Viswanath, Kamal; Johnson, Ryan; Kailasanath, Kailas; Malla, Bhupatindra; Gutmark, Ephraim

2017-11-01

The noise from high performance jet engines of both civilian and military aircraft is an area of active concern. Asymmetric exhaust nozzle configurations, in particular rectangular, potentially offer a passive way of modulating the farfield noise and are likely to become more important in the future. High aspect ratio nozzles offer the further benefit of easier airframe integration. In this study we validate the far field noise for ideally and over expanded supersonic jets issuing from a high aspect ratio rectangular nozzle geometry. Validation of the acoustic data is performed against experimentally recorded sound pressure level (SPL) spectra for a host of observer locations around the asymmetric nozzle. Data is presented for a slightly heated jet case for both nozzle pressure ratios. The contrast in the noise profile from low aspect ratio rectangular and circular nozzle jets are highlighted, especially the variation in the azimuthal direction that shows ``quiet'' and ``loud'' planes in the farfield in the peak noise direction. This variation is analyzed in the context of the effect of mixing at the sharp corners, the sense of the vortex pairs setup in the exit plane, and the evolution of the high aspect ratio exit cross-section as it propagates downstream including possible axis-switching. Supported by Office of Naval Research (ONR) through the Computational Physics Task Area under the NRL 6.1 Base Program.

The aerodynamic characteristics of large angled cones with retrorockets

NASA Technical Reports Server (NTRS)

Jarvinen, P. O.; Adams, R. H.

1970-01-01

Analytical and experimental phases of the subject investigation are described. The analytical program for the single jet determines the terminal shock location, the jet boundary, the interface profile, the bow shock profile, the shear layer growth and the dead air region pressure. The experimental program described was conducted over the range from free stream Mach 0.4 to 2.0 at angles-of-attack up to 18 deg and at thrusting coefficients up to C sub T = T/q sub infinity A sub m = 30. Variables investigated included aeroshell angle, number of nozzles, engine thrust, size of nozzles, nozzle throttling and gas composition. The influence of these variables on the aeroshell stability, drag, and loads was determined by integrating pressure measurements on the aeroshell. The total system forces consist of components due to pure thrust and components due to pressure on the aeroshell arising from the jet-free stream interaction. Shadowgraphs provided flow field geometries which proved to be within 10% of those predicted analytically.

Internal Performance of a Fixed-Shroud Nonaxisymmetric Nozzle Equipped with an Aft-Hood Exhaust Deflector

NASA Technical Reports Server (NTRS)

Asbury, Scott C.

1997-01-01

An investigation was conducted in the model preparation area of the Langley 16-Foot Transonic Tunnel to determine the internal performance of a fixed-shroud nonaxisymmetric nozzle equipped with an aft-hood exhaust deflector. Model geometric parameters investigated included nozzle power setting, aft-hood deflector angle, throat area control with the aft-hood deflector deployed, and yaw vector angle. Results indicate that cruise configurations produced peak performance in the range consistent with previous investigations of nonaxisymmetric convergent-divergent nozzles. The aft-hood deflector produced resultant pitch vector angles that were always less than the geometric aft-hood deflector angle when the nozzle throat was positioned upstream of the deflector exit. Significant losses in resultant thrust ratio occurred when the aft-hood deflector was deployed with an upstream throat location. At each aft-hood deflector angle, repositioning the throat to the deflector exit improved pitch vectoring performance and, in some cases, substantially improved resultant thrust ratio performance. Transferring the throat to the deflector exit allowed the flow to be turned upstream of the throat at subsonic Mach numbers, thereby eliminating losses associated with turning supersonic flow. Internal throat panel deflections were largely unsuccessful in generating yaw vectoring.

The Response of Cryogenic H2/O2 Coaxial Jet Flames to Acoustic Disturbances (POST PRINT)

DTIC Science & Technology

2015-01-05

in the presence of tuned or induced acoustic resonance . Anderson and coworkers have been investigating longitudinal combustion ... resonance that involved periodic blocking of a secondary exhaust nozzle. The periodic flow through the second- ary nozzle results in chamber pressure...Angeles, 2009. 10. Yu, Y., Koeglmeier, S., Sisco, J., and Anderson, W., " Combustion instability of gaseous fuels in a continuously variable resonance

PIV Measurements of Chevrons on F400-Series Tactical Aircraft Nozzle Model

NASA Technical Reports Server (NTRS)

Bridges, James; Wernet, Mark P.; Frate, Franco C.

2011-01-01

Reducing noise of tactical jet aircraft has taken on fresh urgency as core engine technologies allow higher specific-thrust engines and as society become more concerned for the health of its military workforce. Noise reduction on this application has lagged the commercial field as incentives for quieting military aircraft have not been as strong as in their civilian counterparts. And noise reduction strategies employed on civilian engines may not be directly applicable due to the differences in exhaust system architecture and mission. For instance, the noise reduction technology of chevrons, examined in this study, will need to be modified to take into account the special features of tactical aircraft nozzles. In practice, these nozzles have divergent slats that are tied to throttle position, and at take off the jet flow is highly overexpanded as the nozzle is optimized for cruise altitude rather than sea level. In simple oil flow visualization experiments conducted at the onset of the current test program flow barely stays attached at end of nozzle at takeoff conditions. This adds a new twist to the design of chevrons. Upon reaching the nozzle exit the flow shrinks inward radially, meaning that for a chevron to penetrate the flow it must extend much farther away from the baseline nozzle streamline. Another wrinkle is that with a variable divergence angle on the nozzle, the effective penetration will differ with throttle position and altitude. The final note of realism introduced in these experiments was to simulate the manner in which bypass flow is bled into the nozzle wall in real engines to cool the nozzle, which might cause very fat boundary layer at exit. These factors, along with several other issues specific to the application of chevrons to convergent-divergent nozzles have been explored with particle image velocimetry measurements and are presented in this paper.

Numerical analysis and design optimization of supersonic after-burning with strut fuel injectors for scramjet engines

NASA Astrophysics Data System (ADS)

Candon, M. J.; Ogawa, H.

2018-06-01

Scramjets are a class of hypersonic airbreathing engine that offer promise for economical, reliable and high-speed access-to-space and atmospheric transport. The expanding flow in the scramjet nozzle comprises of unburned hydrogen. An after-burning scheme can be used to effectively utilize the remaining hydrogen by supplying additional oxygen into the nozzle, aiming to augment the thrust. This paper presents the results of a single-objective design optimization for a strut fuel injection scheme considering four design variables with the objective of maximizing thrust augmentation. Thrust is found to be augmented significantly owing to a combination of contributions from aerodynamic and combustion effects. Further understanding and physical insights have been gained by performing variance-based global sensitivity analysis, scrutinizing the nozzle flowfields, analyzing the distributions and contributions of the forces acting on the nozzle wall, and examining the combustion efficiency.

Fluidic Thrust Vectoring of an Axisymmetric Exhaust Nozzle at Static Conditions

NASA Technical Reports Server (NTRS)

Wing, David J.; Giuliano, Victor J.

1997-01-01

A sub-scale experimental static investigation of an axisymmetric nozzle with fluidic injection for thrust vectoring was conducted at the NASA Langley Jet Exit Test Facility. Fluidic injection was introduced through flush-mounted injection ports in the divergent section. Geometric variables included injection-port geometry and location. Test conditions included a range of nozzle pressure ratios from 2 to 10 and a range of injection total pressure ratio from no-flow to 1.5. The results indicate that fluidic injection in an axisymmetric nozzle operating at design conditions produced significant thrust-vector angles with less reduction in thrust efficiency than that of a fluidically-vectored rectangular jet. The axisymmetric geometry promoted a pressure relief mechanism around the injection slot, thereby reducing the strength of the oblique shock and the losses associated with it. Injection port geometry had minimal effect on thrust vectoring.

Diagnostics of cathode material loss in cutting plasma torch

NASA Astrophysics Data System (ADS)

Gruber, J.; Šonský, J.; Hlína, J.

2014-07-01

A cutting plasma torch was observed in several ways by a high-speed camera with a focus on the cathode area. In the first experiment, the plasma arc between the nozzle tip and anode was recorded in a series of duty cycles ranging from new unworn cathodes to cathode failure due to wear and material loss. In the second experiment, we used a specially modified nozzle to observe the inside area between the cathode and the nozzle exit through a fused silica window. Finally, using tilted view, we observed a pool of molten hafnium at the cathode tip during the plasma torch operation. The process of cathode material melting, droplet formation, their expulsion and rate of cathode material loss was examined.

Performance and Thrust-to-Weight Optimization of the Dual-Expander Aerospike Nozzle Upper Stage Rocket Engine

DTIC Science & Technology

2012-06-01

calculates a constant convection heat transfer coefficient on the hot and cold side of the cooling jacket wall. The calculated maximum wall temperature for...regeneratively cools the combustion chamber and nozzle. The heat transferred to the fuel from cooling provides enough power to the turbine to power both... heat transfer at the throat compared to a bell nozzle. This increase in heat transfer surface area means more power to the turbine, increased chamber

A 1/10 Scale Model Test of a Fixed Chute Mixer-Ejector Nozzle in Unsuppressed Model. Part 1; Test Overview

NASA Technical Reports Server (NTRS)

Wolter, John D.

2007-01-01

This paper discusses a test of a nozzle concept for a high-speed commercial aircraft. While a great deal of effort has been expended to und erstand the noise-suppressed, take-off performance of mixer-ejector n ozzles, little has been done to assess their performance in unsuppressed mode at other flight conditions. To address this, a 1/10th scale m odel mixer-ejector nozzle in unsuppressed mode was tested at conditio ns representing transonic acceleration, supersonic cruise, subsonic cruise, and approach. Various configurations were tested to understand the effects of acoustic liners and several geometric parameters, such as throat area, expansion ratio, and nozzle length on nozzle performance. Thrust, flow, and internal pressures were measured. A statistica l model of the peak thrust coefficient results is presented and discussed.

Effect of facility variation on the acoustic characteristics of three single stream nozzles

NASA Technical Reports Server (NTRS)

Gutierrez, O. A.

1980-01-01

The characteristics of the jet noise produced by three single stream nozzles were investigated statistically at the NASA-Lewis Research Center outdoor jet acoustic facility. The nozzles consisted of a 7.6 cm diameter convergent conical, a 10.2 cm diameter convergent conical and an 8-lobe daisy nozzle with 7.6 cm equivalent diameter flow area. The same nozzles were tested previously at cold flow conditions in other facilities such as the Royal Aircraft Establishment (RAE) 7.3 m acoustic wind tunnel. The acoustic experiments at NASA covered pressure ratios from 1.4 to 2.5 at total temperatures of 811 K and ambient. The data obtained with four different microphone arrays are compared. The results are also compared with data taken at the RAE facility and with a NASA prediction procedure.

Simulation and characterization of silicon-based 0.5-MHz ultrasonic nozzles

NASA Astrophysics Data System (ADS)

Song, Y. L.; Tsai, S. C.; Chen, W. J.; Chou, Y. F.; Tseng, T. K.; Tsai, C. S.

2004-01-01

This paper compares the simulation results with the experimental results of impedance analysis and longitudinal vibration measurement of micro-fabricated 0.5 MHz silicon-based ultrasonic nozzles. Impedance analysis serves as a good diagnostic tool for evaluation of longitudinal vibration of the nozzles. Each nozzle is made of a piezoelectric drive section and a silicon-resonator consisting of multiple Fourier horns each with half wavelength design and twice amplitude magnification. The experimental results verified the simulation prediction of one pure longitudinal vibration mode at the resonant frequency in excellent agreement with the design value. Furthermore, at the resonant frequency, the measured longitudinal vibration amplitude gain at the nozzle tip increases as the number of Fourier horns (n) increases in good agreement with the theoretical value of 2n. Using this design, very high vibration amplitude at the nozzle tip can be achieved with no reduction in the tip cross sectional area. Therefore, the required electric drive power should be drastically reduced, decreasing the likelihood of transducer failure in ultrasonic atomization.

MULTI-LABORATORY STUDY OF FLOW-INDUCED HEMOLYSIS USING THE FDA BENCHMARK NOZZLE MODEL

PubMed Central

Herbertson, Luke H.; Olia, Salim E.; Daly, Amanda; Noatch, Christopher P.; Smith, William A.; Kameneva, Marina V.; Malinauskas, Richard A.

2015-01-01

Multilaboratory in vitro blood damage testing was performed on a simple nozzle model to determine how different flow parameters and blood properties affect device-induced hemolysis and to generate data for comparison with computational fluid dynamics-based predictions of blood damage as part of an FDA initiative for assessing medical device safety. Three independent laboratories evaluated hemolysis as a function of nozzle entrance geometry, flow rate, and blood properties. Bovine blood anticoagulated with acid citrate dextrose solution (2–80 h post-draw) was recirculated through nozzle-containing and paired nozzle-free control loops for 2 h. Controlled parameters included hematocrit (36 ± 1.5%), temperature (25°C), blood volume, flow rate, and pressure. Three nozzle test conditions were evaluated (n = 26–36 trials each): (i) sudden contraction at the entrance with a blood flow rate of 5 L/min, (ii) gradual cone at the entrance with a 6-L/min blood flow rate, and (iii) sudden-contraction inlet at 6 L/min. The blood damage caused only by the nozzle model was calculated by subtracting the hemolysis generated by the paired control loop test. Despite high intralaboratory variability, significant differences among the three test conditions were observed, with the sharp nozzle entrance causing the most hemolysis. Modified index of hemolysis (MIHnozzle) values were 0.292 ± 0.249, 0.021 ± 0.128, and 1.239 ± 0.667 for conditions i–iii, respectively. Porcine blood generated hemolysis results similar to those obtained with bovine blood. Although the interlaboratory hemolysis results are only applicable for the specific blood parameters and nozzle model used here, these empirical data may help to advance computational fluid dynamics models for predicting blood damage. PMID:25180887

Multiple-orifice liquid injection into hypersonic airstreams and applications to ram C-3 flight

NASA Technical Reports Server (NTRS)

Weaver, W. L.

1972-01-01

Experimental data are presented for the oblique injection of water and three electrophilic liquids (fluorocarbon compounds) through multiple-orifice nozzles from a flat plate and the sides of a hemisphere-cone (0.375 scale of RAM C spacecraft) into hypersonic airstreams. The nozzle patterns included single and multiple orifices, single rows of nozzles, and duplicates of the RAM C-III nozzles. The flat-plate tests were made at Mach 8. Total pressure was varied from 3.45 MN/m2 to 10.34 MN/m2, Reynolds number was varied form 9,840,000 per meter to 19,700,000 per meter, and liquid injection pressure was varied from 0.69 MN/m2 to 3.5 MN/m2. The hemisphere-cone tests were made at Mach 7.3. Total pressure was varied from 1.38 MN/m2, to 6.89 MN/m2, Reynolds number was varied from 3,540,000 per meter to 17,700,000 per meter, and liquid-injection pressure was varied from 0.34 MN/m2 to 4.14 MN/m2. Photographs of the tests and plots of liquid-penetration and spray cross-section area are presented. Maximum penetration was found to vary as the square root of the dynamic-pressure ratio and the square root of the total injection nozzle area. Spray cross-section area was linear with maximum penetration. The test results are used to compute injection parameters for the RAM C-3 flight injection experiment.

Design and experimental evaluation of a high temperature radial turbine with a moveable sidewall nozzle

NASA Technical Reports Server (NTRS)

Rogo, Casimir; Roelke, Richard J.

1987-01-01

The uncooled, 2.27 kg/sec mass flow radial turbine designed to operate at 1477 K in the gas generator of an advanced, variable-capacity 683 kW turboshaft engine was configured with a cooled, movable sidewall nozzle capable of changing the stage flow capacity from 50 to 100 percent of maximum. Overall performance test data were obtained in a turbine test rig that duplicated engine Reynolds numbers; attention is given to the changing of flow capacity by moving the hub or shroud sidewall, vane sidewall leakage, vaneless space sidewall geometry, and nozzle-cooling injection. Data are presented in the form of turbine flow, efficiency, work parameter, and performance mappings.

Comparison of theoretical and experimental thrust performance of a 1030:1 area ratio rocket nozzle at a chamber pressure of 2413 kN/m2 (350 psia)

NASA Technical Reports Server (NTRS)

Smith, Tamara A.; Pavli, Albert J.; Kacynski, Kenneth J.

1987-01-01

The joint Army. Navy, NASA. Air Force (JANNAF) rocket engine peformnace prediction procedure is based on the use of various reference computer programs. One of the reference programs for nozzle analysis is the Two-Dimensional Kinetics (TDK) Program. The purpose of this report is to calibrate the JANNAF procedure incorporated into the December l984 version of the TDK program for the high-area-ratio rocket engine regime. The calibration was accomplished by modeling the performance of a 1030:1 rocket nozzle tested at NASA Lewis Research Center. A detailed description of the experimental test conditions and TDK input parameters is given. The results show that the computer code predicts delivered vacuum specific impulse to within 0.12 to 1.9 percent of the experimental data. Vacuum thrust coefficient predictions were within + or - 1.3 percent of experimental results. Predictions of wall static pressure were within approximately + or - 5 percent of the measured values. An experimental value for inviscid thrust was obtained for the nozzle extension between area ratios of 427.5 and 1030 by using an integration of the measured wall static pressures. Subtracting the measured thrust gain produced by the nozzle between area ratios of 427.5 and 1030 from the inviscid thrust gain yielded experimental drag decrements of 10.85 and 27.00 N (2.44 and 6.07 lb) for mixture ratios of 3.04 and 4.29, respectively. These values correspond to 0.45 and 1.11 percent of the total vacuum thrust. At a mixture ratio of 4.29, the TDK predicted drag decrement was 16.59 N (3.73 lb), or 0.71 percent of the predicted total vacuum thrust.

Influence of Turbulent Flows in the Nozzle on Melt Flow Within a Slab Mold and Stability of the Metal-Flux Interface

NASA Astrophysics Data System (ADS)

Calderon-Ramos, Ismael; Morales, R. D.

2016-06-01

The design of the ports of a casting nozzle has profound effects on the fluid flow patterns in slab molds. The influence of these outlets have also considerable effects on the turbulent flow and turbulence variables inside the nozzle itself. To understand the effects of nozzle design, three approaches were employed: a theoretical analysis based on the turbulent viscosity hypothesis, dimensional analysis (both analyses aided by computer fluid dynamics), and experiments using particle image velocimetry. The first approach yields a linear relation between calculated magnitudes of scalar fields of ɛ (dissipation rate of kinetic energy) and k 2 (square of the turbulent kinetic energy), which is derived from the wall and the logarithmic-wall laws in the boundary layers. The smaller the slope of this linear relation is, the better the performance of a given nozzle is for maintaining the stability of the melt-flux interface. The second approach yields also a linear relation between flow rate of liquid metal and the cubic root of the dissipation rate of kinetic energy. In this case, the larger the slope of the linear relation is, the better the performance of a given nozzle is for maintaining the stability of the melt-flux interface. Finally, PIV measurements in a mold water model, together with equations for estimation of critical melt velocities for slag entrainment, were used to quantify the effects of nozzle design on the dynamics of the metal-slag interface. The three approaches agree in the characterization of turbulent flows in continuous casting molds using different nozzles.

The effect of forward speed on J85 engine noise from suppressor nozzles as measured in the NASA-Ames 40- by 80-foot wind tunnel

NASA Technical Reports Server (NTRS)

Atencio, A., Jr.

1977-01-01

An investigation to determine the effect of forward speed on the exhaust noise from a conical ejector nozzle and three suppressor nozzles mounted behind a J85 engine was performed in a 40- by 80-foot wind tunnel. The nozzles were tested at three engine power settings and at wind tunnel forward speeds up to 91 m/sec (300 ft/sec). In addition, outdoor static tests were conducted to determine (1) the differences between near field and far field measurements, (2) the effect of an airframe on the far field directivity of each nozzle, and (3) the relative suppression of each nozzle with respect to the baseline conical ejector nozzle. It was found that corrections to near field data are necessary to extrapolate to far field data and that the presence of the airframe changed the far field directivity as measured statically. The results show that the effect of forward speed was to reduce the noise from each nozzle more in the area of peak noise, but the change in forward quadrant noise was small or negligible. A comparison of wind tunnel data with available flight test data shows good agreement.

Experimental evaluation of heat transfer on a 1030:1 area ratio rocket nozzle

NASA Technical Reports Server (NTRS)

Kacynski, Kenneth J.; Pavli, Albert J.; Smith, Tamara A.

1987-01-01

A 1030:1 carbon steel, heat-sink nozzle was tested. The test conditions included a nominal chamber pressure of 2413 kN/sq m and a mixture ratio range of 2.78 to 5.49. The propellants were gaseous oxygen and gaseous hydrogen. Outer wall temperature measurements were used to calculate the inner wall temperature and the heat flux and heat rate to the nozzle at specified axial locations. The experimental heat fluxes were compared to those predicted by the Two-Dimensional Kinetics (TDK) computer model analysis program. When laminar boundary layer flow was assumed in the analysis, the predicted values were within 15 percent of the experimental values for the area ratios of 20 to 975. However, when turbulent boundary layer conditions were assumed, the predicted values were approximately 120 percent higher than the experimental values. A study was performed to determine if the conditions within the nozzle could sustain a laminar boundary layer. Using the flow properties predicted by TDK, the momentum-thickness Reynolds number was calculated, and the point of transition to turbulent flow was predicted. The predicted transition point was within 0.5 inches of the nozzle throat. Calculations of the acceleration parameter were then made to determine if the flow conditions could produce relaminarization of the boundary layer. It was determined that if the boundary layer flow was inclined to transition to turbulent, the acceleration conditions within the nozzle would tend to suppress turbulence and keep the flow laminar-like.

Experimental evaluation of heat transfer on a 1030:1 area ratio rocket nozzle

NASA Technical Reports Server (NTRS)

Kacynski, Kenneth J.; Pavli, Albert J.; Smith, Tamara A.

1987-01-01

A 1030:1 carbon steel, heat-sink nozzle was tested. The test conditions included a nominal chamber pressure of 2413 kN/sq m and a mixture ratio range of 2.78 to 5.49. The propellants were gaseous oxygen and gaseous hydrogen. Outer wall temperature measurements were used to calculate the inner wall temperature and the heat flux and heat rate to the nozzle at specified axial locations. The experimental heat fluxes were compared to those predicted by the Two-Dimensional Kinetics (TDK) computer model analysis program. When laminar boundary layer flow was assumed in the analysis, the predicted values were within 15% of the experimental values for the area ratios of 20 to 975. However, when turbulent boundary layer conditions were assumed, the predicted values were approximately 120% higher than the experimental values. A study was performed to determine if the conditions within the nozzle could sustain a laminar boundary layer. Using the flow properties predicted by TDK, the momentum-thickness Reynolds number was calculated, and the point of transition to turbulent flow was predicted. The predicted transition point was within 0.5 inches of the nozzle throat. Calculations of the acceleration parameter were then made to determine if the flow conditions could produce relaminarization of the boundary layer. It was determined that if the boundary layer flow was inclined to transition to turbulent, the acceleration conditions within the nozzle would tend to suppress turbulence and keep the flow laminar-like.

Interaction of Highly Underexpanded Jets with Simulated Lunar Surfaces

NASA Technical Reports Server (NTRS)

Stitt, Leonard E.

1961-01-01

Pressure distributions and erosion patterns on simulated lunar surfaces (hard and soft) and interference effects between the surface and two representative lunar vehicles (cylindrical and spherical) were obtained with cold-air jets at various descent heights and nozzle total-pressure ratios up to 288,000. Surface pressure distributions were dependent on both nozzle area ratio and, nozzle contour. Peak pressures obtained with a sonic nozzle agreed closely with those predicted theoretically for a near-sonic jet expanding into a vacuum. Short bell-shaped nozzles gave annular pressure distributions; the low center pressure resulted from the coalescence of shocks that originated within the nozzle. The high surface pressures were contained within a circle whose diameter was about 16 throat diameters, regardless of nozzle area ratio or contour. The peak pressure increased rapidly as the vehicle approached the surface; for example, at a descent height of 40 throat diameters the peak pressure was 0.4 percent of the chamber pressure, but increased to 6 percent at 13 throat diameters. The exhaust jet eroded a circular concave hole in white sand at descent heights from about 200 to 600 throat diameters. The hole diameter was about 225 throat diameters, while the depth was approximately 60 throat diameters. The sand particles, which formed a conical sheet at a semivertex angle of 50 deg, appeared to follow a ballistic trajectory and at no time struck the vehicle. An increase in pressure was measured on the base of the cylindrical lunar vehicle when it approached to within 14 throat diameters of the hard, flat surface. No interference effects were noted between the spherical model and the surface to descent heights as low as 8 throat diameters.

Supersonic Free-Jet Combustion in a Ramjet Burner

NASA Technical Reports Server (NTRS)

Trefny, Charles J.; Dippold, Vance F., III

2010-01-01

A new dual-mode ramjet combustor concept intended for operation over a wide flight Mach number range is described. Subsonic combustion mode is similar to that of a traditional ram combustor which allows operation at higher efficiency, and to lower flight Mach numbers than current dual-mode scramjets. High speed mode is characterized by supersonic combustion in a free-jet that traverses the subsonic combustion chamber to a variable nozzle. The maximum flight Mach number of this scheme is governed largely by the same physics as its classical counterpart. Although a variable combustor exit aperture is required, the need for fuel staging to accommodate the combustion process is eliminated. Local heating from shock-boundary-layer interactions on combustor walls is also eliminated. Given the parallel nature of the present scheme, overall flowpath length is less than that of present dual-mode configurations. Cycle analysis was done to define the flowpath geometry for computational fluid dynamics (CFD) analysis, and then to determine performance based on the CFD results. CFD results for Mach 5, 8, and 12 flight conditions indicate stable supersonic free-jet formation and nozzle reattachment, thereby establishing the basic feasibility of the concept. These results also reveal the structure of, and interactions between the free-jet and recirculating combustion chamber flows. Performance based on these CFD results is slightly less than that of the constant-pressure-combustion cycle analysis primarily due to these interactions. These differences are quantified and discussed. Additional CFD results at the Mach 8 flight condition show the effects of nozzle throat area variation on combustion chamber pressure, flow structure, and performance. Calculations with constant temperature walls were also done to evaluate heat flux and overall heat loads. Aspects of the concept that warrant further study are outlined. These include diffuser design, ramjet operation, mode transition, loss mechanisms, and the effects of secondary flow for wall cooling and combustion chamber pressurization. Also recommended is an examination of system-level aspects such as weight, thermal management and rocket integration as well as alternate geometries and variable geometry schemes.

Remtech SSME nozzle design TPS

NASA Technical Reports Server (NTRS)

Bancroft, Steven A.; Engel, Carl D.; Pond, John E.

1990-01-01

Thermal damage to the Space Shuttle Main Engine (SSME) aft manifold Thermal Protection System (TPS) has been observed for flights STS-8 through STS-13. This damaged area is located on the ME2 and ME3 and extends over a region of approximately one square foot. Total failure or burn-through of the TPS could lead to severe thermal damage of the SSME manifold and loss of an engine nozzle necessitating nozzle replacement causing significant schedule delays and cost increases. Thermal damage to the manifold can be defined as a situation where the manifold temperature becomes greater than 1300 F; thereby causing loss of heat treatment in the nozzle. Results of Orbiter/nozzle wind tunnel tests and Hot Gas Facility tests of the TPS are presented. Aerothermal and thermal analysis models for the SSME aft manifold are discussed along with the flight predictions, design trajectory and design environment. Finally, the TPS design concept and TPS thermal response are addressed.

Load calculation on the nozzle in a flue gas desulphurization system

NASA Astrophysics Data System (ADS)

Róbert, Olšiak; Zoltán, Fuszko; Zoltán, Csuka

2017-09-01

The desulphurization system is used to remove sulfur oxides from exhaust, so-called flue gases through absorbing them via the sprayed suspension. The suspension delivered from the pump system to the atmospheric bi-directional double hollow cone nozzle has the prescribed working pressure. The unknown mechanical load on the solid body of the nozzle is present through the change of moment due to the flow of the suspension through the bi-directional outflow areas [1], [4]. The calculation of the acting forces and torques in the 3 directions was carried out with the methods of computational fluid dynamics (CFD) in the software ANSYS Fluent. The geometric model of the flow areas of the nozzle were created with the methods of reverse engineering. The computational mesh required by the CFD solver was created, and its quality verified with the standard criteria. The used boundary conditions were defined by the hydraulic parameters of the pump system, the properties of the suspension present in the hydraulic system were specified by sample analysis. The post-processed and analyzed results of the CFD calculation, the pressure-field and the velocity magnitudes in particular directions were further used as input parameters at the mechanical analysis of the load on the bi-directional nozzle.

Development of acoustically lined ejector technology for multitube jet noise suppressor nozzles by model and engine tests over a wide range of jet pressure ratios and temperatures

NASA Technical Reports Server (NTRS)

Atvars, J.; Paynter, G. C.; Walker, D. Q.; Wintermeyer, C. F.

1974-01-01

An experimental program comprising model nozzle and full-scale engine tests was undertaken to acquire parametric data for acoustically lined ejectors applied to primary jet noise suppression. Ejector lining design technology and acoustical scaling of lined ejector configurations were the major objectives. Ground static tests were run with a J-75 turbojet engine fitted with a 37-tube, area ratio 3.3 suppressor nozzle and two lengths of ejector shroud (L/D = 1 and 2). Seven ejector lining configurations were tested over the engine pressure ratio range of 1.40 to 2.40 with corresponding jet velocities between 305 and 610 M/sec. One-fourth scale model nozzles were tested over a pressure ratio range of 1.40 to 4.0 with jet total temperatures between ambient and 1088 K. Scaling of multielement nozzle ejector configurations was also studied using a single element of the nozzle array with identical ejector lengths and lining materials. Acoustic far field and near field data together with nozzle thrust performance and jet aerodynamic flow profiles are presented.

Empirical source noise prediction method with application to subsonic coaxial jet mixing noise

NASA Technical Reports Server (NTRS)

Zorumski, W. E.; Weir, D. S.

1982-01-01

A general empirical method, developed for source noise predictions, uses tensor splines to represent the dependence of the acoustic field on frequency and direction and Taylor's series to represent the dependence on source state parameters. The method is applied to prediction of mixing noise from subsonic circular and coaxial jets. A noise data base of 1/3-octave-band sound pressure levels (SPL's) from 540 tests was gathered from three countries: United States, United Kingdom, and France. The SPL's depend on seven variables: frequency, polar direction angle, and five source state parameters: inner and outer nozzle pressure ratios, inner and outer stream total temperatures, and nozzle area ratio. A least-squares seven-dimensional curve fit defines a table of constants which is used for the prediction method. The resulting prediction has a mean error of 0 dB and a standard deviation of 1.2 dB. The prediction method is used to search for a coaxial jet which has the greatest coaxial noise benefit as compared with an equivalent single jet. It is found that benefits of about 6 dB are possible.

User's manual for the ALS base heating prediction code, volume 2

NASA Technical Reports Server (NTRS)

Reardon, John E.; Fulton, Michael S.

1992-01-01

The Advanced Launch System (ALS) Base Heating Prediction Code is based on a generalization of first principles in the prediction of plume induced base convective heating and plume radiation. It should be considered to be an approximate method for evaluating trends as a function of configuration variables because the processes being modeled are too complex to allow an accurate generalization. The convective methodology is based upon generalizing trends from four nozzle configurations, so an extension to use the code with strap-on boosters, multiple nozzle sizes, and variations in the propellants and chamber pressure histories cannot be precisely treated. The plume radiation is more amenable to precise computer prediction, but simplified assumptions are required to model the various aspects of the candidate configurations. Perhaps the most difficult area to characterize is the variation of radiation with altitude. The theory in the radiation predictions is described in more detail. This report is intended to familiarize a user with the interface operation and options, to summarize the limitations and restrictions of the code, and to provide information to assist in installing the code.

Boundary layer simulator improvement

NASA Technical Reports Server (NTRS)

Praharaj, S. C.; Schmitz, C.; Frost, C.; Engel, C. D.; Fuller, C. E.; Bender, R. L.; Pond, J.

1984-01-01

High chamber pressure expander cycles proposed for orbit transfer vehicles depend primarily on the heat energy transmitted from the combustion products through the thrust wall chamber wall. The heat transfer to the nozzle wall is affected by such variables as wall roughness, relamarization, and the presence of particles in the flow. Motor performance loss for these nozzles with thick boundary layers is inaccurate using the existing procedure coded BLIMPJ. Modifications and innovations to the code are examined. Updated routines are listed.

Performance characteristics of an isolated coannular plug nozzle at transonic speeds

NASA Technical Reports Server (NTRS)

Mercer, C. E.; Burley, J. R., II

1985-01-01

The Langley 16-Foot Transonic Tunnel was used to evaluate the performance characteristics of a coannular plug nozzle at static conditions (Mach number of 0) and at Mach numbers from 0.65 to 1.20. Jet total pressure ratio was varied from 1.0 (jet off) to 10.0. Thirty-seven configurations generated by the combination of three geometric variables - plug angle, shroud boattail length (fixed exit radius), and shroud extension length - were tested.

Slot Nozzle Effects for Reduced Sonic Boom on a Generic Supersonic Wing Section

NASA Technical Reports Server (NTRS)

Caster, Raymond S.

2010-01-01

NASA has conducted research programs to reduce or eliminate the operational restrictions of supersonic aircraft over populated areas. Restrictions are due to the disturbance from the sonic boom, caused by the coalescence of shock waves formed off the aircraft. Results from two-dimensional computational fluid dynamic (CFD) analyses (performed on a baseline Mach 2.0 nozzle in a simulated Mach 2.2 flow) indicate that over-expanded and under-expanded operation of the nozzle has an effect on the N-wave boom signature. Analyses demonstrate the feasibility of reducing the magnitude of the sonic boom N-wave by controlling the nozzle plume interaction with the nozzle boat tail shock structure. This work was extended to study the impact of integrating a high aspect ratio exhaust nozzle or long slot nozzle on the trailing edge of a supersonic wing. The nozzle is operated in a highly under-expanded condition, creating a large exhaust plume and a shock at the trailing edge of the wing. This shock interacts with and suppresses the expansion wave caused by the wing, a major contributor to the sonic boom signature. The goal was to reduce the near field pressures caused by the expansion using a slot nozzle located at the wing trailing edge. Results from CFD analysis on a simulated wing cross-section and a slot nozzle indicate potential reductions in sonic boom signature compared to a baseline wing with no propulsion or trailing edge exhaust. Future studies could investigate if this effect could be useful on a supersonic aircraft for main propulsion, auxiliary propulsion, or flow control.

Research Amplitudo Vibration On Holder Due To The Process Of Lathe Nozzle Rocket RX 450

NASA Astrophysics Data System (ADS)

Ediwan; Budi Djatmiko, Agus; Dody Arisandi, EfFendy; Purnomo, Heri; Ibadi, Mahfud

2018-04-01

The main function of the rocket nozzle is to convert the enthalpy efficiency from combustion gas to kinetic energy and also to make high velocity out of the gas. The rocket nozzle usually consists of a converging and diverging part. With a smaller area on the neck and enlarged at the exit area. The velocity flow through the nozzle enlarges into the speed of sound through the neck and then becomes super sonic in the divergent part. Nozzle making or machining using conventional lathes, first performed is drilling on a massive metal that is bonded to the veneer, then after a sufficient gap is done deep-boring. At the time of the process of lathe in the nozzle RX 450 there is an obstacle that is vibrating tool holder chisel or holder so it is worried about not precision of the process of lathe. This should not happen because it can cause failure in the latter for it needs to be studied and studied further so that the lathe process goes accordingly. The holder material of ST 60 with a modulus of elasticity 200 GPa and a nozzle material of AISI 4340 alloy steel with σyield = 470 MPa, Shear Modulus G = 80 GPa. The purpose of this research is to observe the amplitude of vibration on the holder due to RX- 450 nozzle lathe processing for the purpose of amplitude that occurs in accordance with the desired so that the nozzle structure is no damage process. The result of the research was obtained holder with length (L) 80cm, profile width (B) 5 cm, height of profile (H) 10 cm, turning machine ω = 8.98 rad / sec and natural holder frequency ωn = 89.8 rad / second, Amplitude of vibration of δ = 1.21 mm, while the amplitude of the design X = 1.22 mm From the results of this study it can be said that the holder of a chisel or holder can be used as a tool at the time of RX nozzle retrieval process and is quite safe because it works under the condition ω/ω n <0 3 (Plans and Specifications of A Suborbital Rocket Payload "AKPV Engineering University of Wyoming 2009 )

Refrigerant pressurization system with a two-phase condensing ejector

DOEpatents

Bergander, Mark [Madison, CT

2009-07-14

A refrigerant pressurization system including an ejector having a first conduit for flowing a liquid refrigerant therethrough and a nozzle for accelerating a vapor refrigerant therethrough. The first conduit is positioned such that the liquid refrigerant is discharged from the first conduit into the nozzle. The ejector includes a mixing chamber for condensing the vapor refrigerant. The mixing chamber comprises at least a portion of the nozzle and transitions into a second conduit having a substantially constant cross sectional area. The condensation of the vapor refrigerant in the mixing chamber causes the refrigerant mixture in at least a portion of the mixing chamber to be at a pressure greater than that of the refrigerant entering the nozzle and greater than that entering the first conduit.

Full Navier-Stokes analysis of a two-dimensional mixer/ejector nozzle for noise suppression

NASA Technical Reports Server (NTRS)

Debonis, James R.

1992-01-01

A three-dimensional full Navier-Stokes (FNS) analysis was performed on a mixer/ejector nozzle designed to reduce the jet noise created at takeoff by a future supersonic transport. The PARC3D computational fluid dynamics (CFD) code was used to study the flow field of the nozzle. The grid that was used in the analysis consisted of approximately 900,000 node points contained in eight grid blocks. Two nozzle configurations were studied: a constant area mixing section and a diverging mixing section. Data are presented for predictions of pressure, velocity, and total temperature distributions and for evaluations of internal performance and mixing effectiveness. The analysis provided good insight into the behavior of the flow.

An experimental investigation of internal area ruling for transonic and supersonic channel flow

NASA Technical Reports Server (NTRS)

Roberts, W. B.; Vanrintel, H. L.; Rizvi, G.

1982-01-01

A simulated transonic rotor channel model was examined experimentally to verify the flow physics of internal area ruling. Pressure measurements were performed in the high speed wind tunnel at transonic speeds with Mach 1.5 and Mach 2 nozzle blocks to get an indication of the approximate shock losses. The results showed a reduction in losses due to internal area ruling with the Mach 1.5 nozzle blocks. The reduction in total loss coefficient was of the order of 17 percent for a high blockage model and 7 percent for a cut-down model.

DOE Office of Scientific and Technical Information (OSTI.GOV)

Yu, Y.J.; Sohn, G.H.; Kim, Y.J.

Typical LBB (Leak-Before-Break) analysis is performed for the highest stress location for each different type of material in the high energy pipe line. In most cases, the highest stress occurs at the nozzle and pipe interface location at the terminal end. The standard finite element analysis approach to calculate J-Integral values at the crack tip utilizes symmetry conditions when modeling near the nozzle as well as away from the nozzle region to minimize the model size and simplify the calculation of J-integral values at the crack tip. A factor of two is typically applied to the J-integral value to accountmore » for symmetric conditions. This simplified analysis can lead to conservative results especially for small diameter pipes where the asymmetry of the nozzle-pipe interface is ignored. The stiffness of the residual piping system and non-symmetries of geometry along with different material for the nozzle, safe end and pipe are usually omitted in current LBB methodology. In this paper, the effects of non-symmetries due to geometry and material at the pipe-nozzle interface are presented. Various LBB analyses are performed for a small diameter piping system to evaluate the effect a nozzle has on the J-integral calculation, crack opening area and crack stability. In addition, material differences between the nozzle and pipe are evaluated. Comparison is made between a pipe model and a nozzle-pipe interface model, and a LBB PED (Piping Evaluation Diagram) curve is developed to summarize the results for use by piping designers.« less

Internal Designs Application for Inlet and Nozzle Aeroperformance Improvement

NASA Technical Reports Server (NTRS)

Gilinsky, M.; Blankson, I. M.

2000-01-01

The following research results are based on development of an approach previously proposed by the authors for optimum nozzle design to obtain maximum thrust. The design was denoted a Telescope nozzle. A Telescope nozzle contains one or several internal designs of certain location, which are inserted at certain locations into a divergent conical or planar main nozzle near its exit. Such a design provides additional thrust augmentation over 20% by comparison with the optimum single nozzle of equivalent lateral area. What is more, recent experimental acoustic tests have discovered an essential noise reduction due to Telescope nozzles application. In this paper, some additional theoretical results are presented for Telescope nozzles and a similar approach is applied for aeroperformance improvement of a supersonic inlet. In addition, a classic gas dynamics problem of a similar supersonic flow into a plate has been analyzed. In some particular cases, new exact analytical solutions are obtained for a flow into a wedge with an oblique shock wave. Numerical simulations were conducted for supersonic flow into a divergent portion of a 2D or axisymmetric nozzle with several plane or conical designs as well as into a 2D or axisymmetric supersonic inlet with a forebody. The 1st order Kryko-Godunov march- ing numerical scheme for inviscid supersonic flows was used. Several cases were tested using the NASA CFL3d code based on full Navier-Stokes equations. Numerical simulation results have confirmed essential benefits of Telescope design applications in propulsion systems.

Internal Designs Application for Inlet and Nozzle Aeroperformance Improvement

NASA Technical Reports Server (NTRS)

Gilinsky, M.; Blankson, I. M.

2000-01-01

The following research results are based on development of an approach previously proposed by the authors for optimum nozzle design to obtain maximum thrust. The design was denoted a Telescope nozzle. A Telescope nozzle contains one or several internal designs of certain location, which are inserted at certain locations into a divergent conical or planar main nozzle near its exit. Such a design provides additional thrust augmentation over 20% by comparison with the optimum single nozzle of equivalent lateral area. What is more, recent experimental acoustic tests have discovered an essential noise reduction due to Telescope nozzles application. In this paper, some additional theoretical results are presented for Telescope nozzles and a similar approach is applied for aeroperformance improvement of a supersonic inlet. In addition, a classic gas dynamics problem of a similar supersonic flow into a plate has been analyzed. In some particular cases, new exact analytical solutions are obtained for a flow into a wedge with an oblique shock wave. Numerical simulations were conducted for supersonic flow into a divergent portion of a 2D or axisymmetric nozzle with several plane or conuical designs as well as into a 2D or axisymmetric supersonic inlet with a forebody. The 1st order Kryko-Godunov marching numerical scheme for inviscid supersonic flows was used. Several cases were tested using the NASA CFL3d code based on full Navier-Stokes equations. Numerical simulation results have confirmed essential benefits of Telescope design applications in propulsion systems.

Internal Designs Application for Inlet and Nozzle Aeroperformance Improvement

NASA Technical Reports Server (NTRS)

Gilinsky, M.; Blankson, I. M.

2000-01-01

The following research results are based on development of an approach previously proposed by the authors for optimum nozzle design to obtain maximum thrust. The design was denoted a Telescope nozzle. A Telescope nozzle contains one or several internal designs of certain location, which are inserted at certain locations into a divergent conical or planar main nozzle near its exit. Such a design provides additional thrust augmentation over 20% by comparison with the optimum single nozzle of equivalent lateral area. What is more, recent experimental acoustic tests have discovered an essential noise reduction due to Telescope nozzles application. In this paper, some additional theoretical results are presented for Telescope nozzles and a similar approach is applied for aeroperformance improvement of a supersonic inlet. In addition, a classic gas dynamics problem of a similar supersonic flow into a plate has been analyzed. In some particular cases, new exact analytical solutions are obtained for a flow into a wedge with an oblique shock wave. Numerical simulations were conducted for supersonic flow into a divergent portion of a 2D or axisymmetric nozzle with several plane or conical designs as well as into a 2D or axisymmetric supersonic inlet with a forebody. The 1st order Kryko-Godunov marching numerical scheme for inviscid supersonic flows was used. Several cases were tested using the NASA CFL3d code based on full Navier-Stokes equations. Numerical simulation results have confirmed essential benefits of Telescope design applications in propulsion systems.

Effects of nozzle-strut integrated design concepton on the subsonic turbine stage flowfield

NASA Astrophysics Data System (ADS)

Liu, Jun; Du, Qiang; Liu, Guang; Wang, Pei; Zhu, Junqiang

2014-10-01

In order to shorten aero-engine axial length, substituting the traditional long chord thick strut design accompanied with the traditional low pressure(LP) stage nozzle, LP turbine is integrated with intermediate turbine duct (ITD). In the current paper, five vanes of the first stage LP turbine nozzle is replaced with loaded struts for supporting the engine shaft, and providing oil pipes circumferentially which fulfilled the areo-engine structure requirement. However, their bulky geometric size represents a more effective obstacle to flow from high pressure (HP) turbine rotor. These five struts give obvious influence for not only the LP turbine nozzle but also the flowfield within the ITD, and hence cause higher loss. Numerical investigation has been undertaken to observe the influence of the Nozzle-Strut integrated design concept on the flowfield within the ITD and the nearby nozzle blades. According to the computational results, three main conclusions are finally obtained. Firstly, a noticeable low speed area is formed near the strut's leading edge, which is no doubt caused by the potential flow effects. Secondly, more severe radial migration of boundary layer flow adjacent to the strut's pressure side have been found near the nozzle's trailing edge. Such boundary layer migration is obvious, especially close to the shroud domain. Meanwhile, radial pressure gradient aggravates this phenomenon. Thirdly, velocity distribution along the strut's pressure side on nozzle's suction surface differs, which means loading variation of the nozzle. And it will no doubt cause nonuniform flowfield faced by the downstream rotor blade.

Some tests on small-scale rectangular throat ejector. [thrust augmentation for V/STOL aircraft

NASA Technical Reports Server (NTRS)

Dean, W. N., Jr.; Franke, M. E.

1979-01-01

A small scale rectangular throat ejector with plane slot nozzles and a fixed throat area was tested to determine the effects of diffuser sidewall length, diffuser area ratio, and sidewall nozzle position on thrust and mass augmentation. The thrust augmentation ratio varied from approximately 0.9 to 1.1. Although the ejector did not have good thrust augmentation performance, the effects of the parameters studied are believed to indicate probable trends in thrust augmenting ejectors.

Model aerodynamic test results for two variable cycle engine coannular exhaust systems at simulated takeoff and cruise conditions. [Lewis 8 by 6-foot supersonic wind tunnel tests

NASA Technical Reports Server (NTRS)

Nelson, D. P.

1980-01-01

Wind tunnel tests were conducted to evaluate the aerodynamic performance of a coannular exhaust nozzle for a proposed variable stream control supersonic propulsion system. Tests were conducted with two simulated configurations differing primarily in the fan duct flowpaths: a short flap mechanism for fan stream control with an isentropic contoured flow splitter, and an iris fan nozzle with a conical flow splitter. Both designs feature a translating primary plug and an auxiliary inlet ejector. Tests were conducted at takeoff and simulated cruise conditions. Data were acquired at Mach numbers of 0, 0.36, 0.9, and 2.0 for a wide range of nozzle operating conditions. At simulated supersonic cruise, both configurations demonstrated good performance, comparable to levels assumed in earlier advanced supersonic propulsion studies. However, at subsonic cruise, both configurations exhibited performance that was 6 to 7.5 percent less than the study assumptions. At take off conditions, the iris configuration performance approached the assumed levels, while the short flap design was 4 to 6 percent less.

Plug nozzles: The ultimate customer driven propulsion system

NASA Technical Reports Server (NTRS)

Aukerman, Carl A.

1991-01-01

This paper presents the results of a study applying the plug cluster nozzle concept to the propulsion system for a typical lunar excursion vehicle. Primary attention for the design criteria is given to user defined factors such as reliability, low volume, and ease of propulsion system development. Total thrust and specific impulse are held constant in the study while other parameters are explored to minimize the design chamber pressure. A brief history of the plug nozzle concept is included to point out the advanced level of technology of the concept and the feasibility of exploiting the variables considered in this study. The plug cluster concept looks very promising as a candidate for consideration for the ultimate customer driven propulsion system.

Plug nozzles - The ultimate customer driven propulsion system. [applied to manned lunar and Martian landers

NASA Technical Reports Server (NTRS)

Aukerman, Carl A.

1991-01-01

This paper presents the results of a study applying the plug cluster nozzle concept to the propulsion system for a typical lunar excursion vehicle. Primary attention for the design criteria is given to user defined factors such as reliability, low volume, and ease of propulsion system development. Total thrust and specific impulse are held constant in the study while other parameters are explored to minimize the design chamber pressure. A brief history of the plug nozzle concept is included to point out the advanced level of technology of the concept and the feasibility of exploiting the variables considered in the study. The plug cluster concept looks very promising as a candidate for consideration for the ultimate customer driven propulsion system.

Aeroacoustics of supersonic jet flows from contoured and solid/porous conical plug-nozzles

NASA Technical Reports Server (NTRS)

Dosanjh, Darshan S.; Das, Indu S.

1987-01-01

The results of an experimental study of the acoustic far-field, the shock associated noise, and the nature of the repetitive shock structure of supersonic jet flows issuing from plug-nozzles having externally-expanded plugs with pointed termination operated at a range of supercritical pressure ratios Xi approaching 2 to 4.5 are reported. The plug of one of these plug-nozzles was contoured. The other plug-nozzles had short conical plugs with either a solid surface or a combination of solid/porous surface of different porosities. The contoured and the uncontoured plug-nozzles had the same throat area and the same annulus-radius ratio K = R sub p/R sub N = 0.43. As the result of modifications of the shock structure, the acoustic performance of improperly expanded jet flows of an externally-expanded short uncontoured plug of an appropriate geometry with suitably perforated plug and a pointed termination, is shown to approach the acoustic performance of a shock-free supersonic jet issuing from an equivalent externally-expanded contoured plug-nozzle.

Static performance of vectoring/reversing non-axisymmetric nozzles

NASA Technical Reports Server (NTRS)

Willard, C. M.; Capone, F. J.; Konarski, M.; Stevens, H. L.

1977-01-01

An experimental program sponsored by the Air Force Flight Dynamics Laboratory is currently in progress to determine the internal and installed performance characteristics of five different thrust vectoring/reversing non-axisymmetric nozzle concepts for tactical fighter aircraft applications. Internal performance characteristics for the five non-axisymmetric nozzles and an advanced technology axisymmetric baseline nozzle were determined in static tests conducted in January 1977 at the NASA-Langley Research Center. The non-axisymmetric nozzle models were tested at thrust deflection angles of up to 30 degrees from horizontal at throat areas associated with both dry and afterburning power. In addition, dry power reverse thrust geometries were tested for three of the concepts. The best designs demonstrated internal performance levels essentially equivalent to the baseline axisymmetric nozzle at unvectored conditions. The best designs also gave minimum performance losses due to vectoring, and reverse thrust levels up to 50% of maximum dry power forward thrust. The installed performance characteristics will be established based on wind tunnel testing to be conducted at Arnold Engineering Development Center in the fall of 1977.

An all-at-once reduced Hessian SQP scheme for aerodynamic design optimization

NASA Technical Reports Server (NTRS)

Feng, Dan; Pulliam, Thomas H.

1995-01-01

This paper introduces a computational scheme for solving a class of aerodynamic design problems that can be posed as nonlinear equality constrained optimizations. The scheme treats the flow and design variables as independent variables, and solves the constrained optimization problem via reduced Hessian successive quadratic programming. It updates the design and flow variables simultaneously at each iteration and allows flow variables to be infeasible before convergence. The solution of an adjoint flow equation is never needed. In addition, a range space basis is chosen so that in a certain sense the 'cross term' ignored in reduced Hessian SQP methods is minimized. Numerical results for a nozzle design using the quasi-one-dimensional Euler equations show that this scheme is computationally efficient and robust. The computational cost of a typical nozzle design is only a fraction more than that of the corresponding analysis flow calculation. Superlinear convergence is also observed, which agrees with the theoretical properties of this scheme. All optimal solutions are obtained by starting far away from the final solution.

Acoustic and aerodynamic performance of a 1.83 meter (6 foot) diameter 1.2 pressure ratio fan (QF-6). [for short takeoff aircraft

NASA Technical Reports Server (NTRS)

Woodward, R. P.; Lucas, J. G.; Stakolich, E. G.

1974-01-01

A 1.2-pressure-ratio, 1.83-meter-(6-ft-) diameter experimental fan stage with characteristics suitable for use in STOL aircraft engines was tested for acoustic and aerodynamic performance. The design incorporated features for low noise, including absence of inlet guide vanes, low rotor-blade-tip speed, low aerodynamic blade loading, and long axial spacing between the rotor and stator rows. The stage was run with four nozzles of different area. The perceived noise along a 152.4 meter (500-ft) sideline was rear-quadrant dominated with a maximum design-point level of 103.9 PNdb. The acoustic 1/3-octave results were analytically separated into broadband and pure-tone components. It was found that the stage noise levels generally increase with a decrease in nozzle area, with this increase observed primarily in the broadband noise component. A stall condition was documented acoustically with a 90-percent-of-design-area nozzle.

Measurement and analysis of a small nozzle plume in vacuum

NASA Technical Reports Server (NTRS)

Penko, P. F.; Boyd, I. D.; Meissner, D. L.; Dewitt, K. J.

1993-01-01

Pitot pressures and flow angles are measured in the plume of a nozzle flowing nitrogen and exhausting to a vacuum. Total pressures are measured with Pitot tubes sized for specific regions of the plume and flow angles measured with a conical probe. The measurement area for total pressure extends 480 mm (16 exit diameters) downstream of the nozzle exit plane and radially to 60 mm (1.9 exit diameters) off the plume axis. The measurement area for flow angle extends to 160 mm (5 exit diameters) downstream and radially to 60 mm. The measurements are compared to results from a numerical simulation of the flow that is based on kinetic theory and uses the direct-simulation Monte Carlo (DSMC) method. Comparisons of computed results from the DSMC method with measurements of flow angle display good agreement in the far-field of the plume and improve with increasing distance from the exit plane. Pitot pressures computed from the DSMC method are in reasonably good agreement with experimental results over the entire measurement area.

Exact Solution of the Two-Dimensional Problem on an Impact Ideal-Liquid Jet

NASA Astrophysics Data System (ADS)

Belik, V. D.

2018-05-01

The two-dimensional problem on the collision of a potential ideal-liquid jet, outflowing from a reservoir through a nozzle, with an infinite plane obstacle was considered for the case where the distance between the nozzle exit section and the obstacle is finite. An exact solution of this problem has been found using methods of the complex-variable function theory. Simple analytical expressions for the complex velocity of the liquid, its flow rate, and the force of action of the jet on the obstacle have been obtained. The velocity distributions of the liquid at the nozzle exit section, in the region of spreading of the jet, and at the obstacle have been constructed for different distances between the nozzle exit section and the obstacle. Analytical expressions for the thickness of the boundary layer and the Nusselt number at the point of stagnation of the jet have been obtained. A number of distributions of the local friction coefficient and the Nusselt number of the indicated jet are presented.

Jet Noise Reduction Potential from Emerging Variable Cycle Technologies

NASA Technical Reports Server (NTRS)

Henderson, Brenda; Bridges, James; Wernet, Mark

2012-01-01

Acoustic and flow-field experiments were conducted on exhaust concepts for the next generation supersonic, commercial aircraft. The concepts were developed by Lockheed Martin (LM), Rolls-Royce Liberty Works (RRLW), and General Electric Global Research (GEGR) as part of an N+2 (next generation forward) aircraft system study initiated by the Supersonics Project in NASA s Fundamental Aeronautics Program. The experiments were conducted in the Aero-Acoustic Propulsion Laboratory at the NASA Glenn Research Center. The exhaust concepts utilized ejectors, inverted velocity profiles, and fluidic shields. One of the ejector concepts was found to produce stagnant flow within the ejector and the other ejector concept produced discrete-frequency tones that degraded the acoustic performance of the model. The concept incorporating an inverted velocity profile and fluid shield produced overall-sound-pressure-level reductions of 6 dB relative to a single stream nozzle at the peak jet noise angle for some nozzle pressure ratios. Flow separations in the nozzle degraded the acoustic performance of the inverted velocity profile model at low nozzle pressure ratios.

Jet Noise Reduction Potential From Emerging Variable Cycle Technologies

NASA Technical Reports Server (NTRS)

2012-01-01

Acoustic and flow-field experiments were conducted on exhaust concepts for the next generation supersonic, commercial aircraft. The concepts were developed by Lockheed Martin (LM), Rolls-Royce Liberty Works (RRLW), and General Electric Global Research (GEGR) as part of an N+2 (next generation forward) aircraft system study initiated by the Supersonics Project in NASA s Fundamental Aeronautics Program. The experiments were conducted in the Aero-Acoustic Propulsion Laboratory at the NASA Glenn Research Center. The exhaust concepts utilized ejectors, inverted velocity profiles, and fluidic shields. One of the ejector concepts was found to produce stagnant flow within the ejector and the other ejector concept produced discrete-frequency tones that degraded the acoustic performance of the model. The concept incorporating an inverted velocity profile and fluid shield produced overall-sound-pressure-level reductions of 6 dB relative to a single stream nozzle at the peak jet noise angle for some nozzle pressure ratios. Flow separations in the nozzle degraded the acoustic performance of the inverted velocity profile model at low nozzle pressure ratios.

Aerodynamic characteristics of a large-scale semispan model with a swept wing and an augmented jet flap with hypermixing nozzles. [Ames 40- by 80-Foot Wind Tunnel and Static Test Facility

NASA Technical Reports Server (NTRS)

Aiken, T. N.; Falarski, M. D.; Koenin, D. G.

1979-01-01

The aerodynamic characteristics of the augmentor wing concept with hypermixing primary nozzles were investigated. A large-scale semispan model in the Ames 40- by 80-Foot Wind Tunnel and Static Test Facility was used. The trailing edge, augmentor flap system occupied 65% of the span and consisted of two fixed pivot flaps. The nozzle system consisted of hypermixing, lobe primary nozzles, and BLC slot nozzles at the forward inlet, both sides and ends of the throat, and at the aft flap. The entire wing leading edge was fitted with a 10% chord slat and a blowing slot. Outboard of the flap was a blown aileron. The model was tested statically and at forward speed. Primary parameters and their ranges included angle of attack from -12 to 32 degrees, flap angles of 20, 30, 45, 60 and 70 degrees, and deflection and diffuser area ratios from 1.16 to 2.22. Thrust coefficients ranged from 0 to 2.73, while nozzle pressure ratios varied from 1.0 to 2.34. Reynolds number per foot varied from 0 to 1.4 million. Analysis of the data indicated a maximum static, gross augmentation of 1.53 at a flap angle of 45 degrees. Analysis also indicated that the configuration was an efficient powered lift device and that the net thrust was comparable with augmentor wings of similar static performance. Performance at forward speed was best at a diffuser area ratio of 1.37.

Low Noise Exhaust Nozzle Technology Development

NASA Technical Reports Server (NTRS)

Majjigi, R. K.; Balan, C.; Mengle, V.; Brausch, J. F.; Shin, H.; Askew, J. W.

2005-01-01

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). Development of a propulsion system that satisfies strict airport noise regulations and provides high levels of cruise and transonic performance with adequate takeoff performance, at an acceptable weight, is critical to the success of any HSCT program. The principal objectives were to: 1. Develop a preliminary design of an innovative 2-D exhaust nozzle with the goal of meeting FAR36 Stage III noise levels and providing high levels of cruise performance with a high specific thrust for Mach 2.4 HSCT with a range of 5000 nmi and a payload of 51,900 lbm, 2. Employ advanced acoustic and aerodynamic codes during preliminary design, 3. Develop a comprehensive acoustic and aerodynamic database through scale-model testing of low-noise, high-performance, 2-D nozzle configurations, based on the preliminary design, and 4. Verify acoustic and aerodynamic predictions by means of scale-model testing. The results were: 1. The preliminary design of a 2-D, convergent/divergent suppressor ejector nozzle for a variable-cycle engine powered, Mach 2.4 HSCT was evolved, 2. Noise goals were predicted to be achievable for three takeoff scenarios, and 3. Impact of noise suppression, nozzle aerodynamic performance, and nozzle weight on HSCT takeoff gross weight were assessed.

Optimization of the inter-tablet coating uniformity for an active coating process at lab and pilot scale.

PubMed

Just, Sarah; Toschkoff, Gregor; Funke, Adrian; Djuric, Dejan; Scharrer, Georg; Khinast, Johannes; Knop, Klaus; Kleinebudde, Peter

2013-11-30

The objective of this study was to enhance the inter-tablet coating uniformity in an active coating process at lab and pilot scale by statistical design of experiments. The API candesartan cilexetil was applied onto gastrointestinal therapeutic systems containing the API nifedipine to obtain fixed dose combinations of these two drugs with different release profiles. At lab scale, the parameters pan load, pan speed, spray rate and number of spray nozzles were examined. At pilot scale, the parameters pan load, pan speed, spray rate, spray time, and spray pressure were investigated. A low spray rate and a high pan speed improved the coating uniformity at both scales. The number of spray nozzles was identified as the most influential variable at lab scale. With four spray nozzles, the highest CV value was equal to 6.4%, compared to 13.4% obtained with two spray nozzles. The lowest CV of 4.5% obtained with two spray nozzles was further reduced to 2.3% when using four spray nozzles. At pilot scale, CV values between 2.7% and 11.1% were achieved. Since the test of uniformity of dosage units accepts CV values of up to 6.25%, this active coating process is well suited to comply with the pharmacopoeial requirements. Copyright © 2013 Elsevier B.V. All rights reserved.

Effects of Fuel and Nozzle Characteristics on Micro Gas Turbine System: A Review

NASA Astrophysics Data System (ADS)

Akasha Hashim, Muhammad; Khalid, Amir; Salleh, Hamidon; Sunar, Norshuhaila Mohamed

2017-08-01

For many decades, gas turbines have been used widely in the internal combustion engine industry. Due to the deficiency of fossil fuel and the concern of global warming, the used of bio-gas have been recognized as one of most clean fuels in the application of engine to improve performance of lean combustion and minimize the production of NOX and PM. This review paper is to understand the combustion performance using dual-fuel nozzle for a micro gas turbine that was basically designed as a natural gas fuelled engine, the nozzle characteristics of the micro gas turbine has been modelled and the effect of multi-fuel used were investigated. The used of biogas (hydrogen) as substitute for liquid fuel (methane) at constant fuel injection velocity, the flame temperature is increased, but the fuel low rate reduced. Applying the blended fuel at constant fuel rate will increased the flame temperature as the hydrogen percentages increased. Micro gas turbines which shows the uniformity of the flow distribution that can be improved without the increase of the pressure drop by applying the variable nozzle diameters into the fuel supply nozzle design. It also identifies the combustion efficiency, better fuel mixing in combustion chamber using duel fuel nozzle with the largest potential for the future. This paper can also be used as a reference source that summarizes the research and development activities on micro gas turbines.

Flame structure of wall-impinging diesel fuel sprays injected by group-hole nozzles

DOE Office of Scientific and Technical Information (OSTI.GOV)

Gao, Jian; Moon, Seoksu; Nishida, Keiya

This paper describes an investigation of the flame structure of wall-impinging diesel sprays injected by group-hole nozzles in a constant-volume combustion vessel at experimental conditions typical of a diesel engine. The particular emphasis was on the effect of the included angle between two orifices (0-15 deg. in current study) on the flame structure and combustion characteristics under various simulated engine load conditions. The laser absorption scattering (LAS) technique was applied to analyze the spray and mixture properties. Direct flame imaging and OH chemiluminescence imaging were utilized to quantify the ignition delay, flame geometrical parameters, and OH chemiluminescence intensity. The imagesmore » show that the asymmetric flame structure emerges in wall-impinging group-hole nozzle sprays as larger included angle and higher engine load conditions are applied, which is consistent with the spray shape observed by LAS. Compared to the base nozzle, group-hole nozzles with large included angles yield higher overall OH chemiluminescence intensity, wider flame area, and greater proportion of high OH intensity, implying the better fuel/air mixing and improved combustion characteristics. The advantages of group-hole nozzle are more pronounced under high load conditions. Based on the results, the feasibility of group-hole nozzle for practical direct injection diesel engines is also discussed. It is concluded that the asymmetric flame structure of a group-hole nozzle spray is favorable to reduce soot formation over wide engine loads. However, the hole configuration of the group-hole nozzle should be carefully considered so as to achieve proper air utilization in the combustion chamber. Stoichiometric diesel combustion is another promising application of group-hole nozzle. (author)« less

Wind Tunnel Model Design for the Study of Plume Effects on Sonic Boom for Isolated Exhaust Nozzles

NASA Technical Reports Server (NTRS)

Castner, Raynold S.

2010-01-01

A low cost test capability was developed at the NASA Glenn Research Center 1- by 1-Foot Supersonic Wind Tunnel (SWT), with a goal to reduce the disturbance caused by supersonic aircraft flight over populated areas. This work focused on the shock wave structure caused by the exhaust nozzle plume. Analysis and design was performed on a new rig to test exhaust nozzle plume effects on sonic boom signature. Test capability included a baseline nozzle test article and a wind tunnel model consisting of a strut, a nosecone and an upper plenum. Analysis was performed on the external and internal aerodynamic configuration, including the shock reflections from the wind tunnel walls caused by the presence of the model nosecone. This wind tunnel model was designed to operate from Mach 1.4 to Mach 3.0 with nozzle pressure ratios from 6 to 12 and altitudes from 30,000 ft (4.36 psia) to 50,000 ft (1.68 psia). The model design was based on a 1 in. outer diameter, was 9 in. in overall length, and was mounted in the wind tunnel on a 3/8 in. wide support strut. For test conditions at 50,000 ft the strut was built to supply 90 psia of pressure, and to achieve 20 psia at the nozzle inlet with a maximum nozzle pressure of 52 psia. Instrumentation was developed to measure nozzle pressure ratio, and an external static pressure probe was designed to survey near field static pressure profiles at one nozzle diameter above the rig centerline. Model layout placed test nozzles between two transparent sidewalls in the 1 1 SWT for Schlieren photography and comparison to CFD analysis.

Wind Tunnel Model Design for the Study of Plume Effects on Sonic Boom for Isolated Exhaust Nozzles

NASA Technical Reports Server (NTRS)

Castner, Raymond S.

2009-01-01

A low cost test capability was developed at the NASA Glenn Research Center 1- by 1-Foot Supersonic Wind Tunnel (SWT), with a goal to reduce the disturbance caused by supersonic aircraft flight over populated areas. This work focused on the shock wave structure caused by the exhaust nozzle plume. Analysis and design was performed on a new rig to test exhaust nozzle plume effects on sonic boom signature. Test capability included a baseline nozzle test article and a wind tunnel model consisting of a strut, a nose cone and an upper plenum. Analysis was performed on the external and internal aerodynamic configuration, including the shock reflections from the wind tunnel walls caused by the presence of the model nosecone. This wind tunnel model was designed to operate from Mach 1.4 to Mach 3.0 with nozzle pressure ratios from 6 to 12 and altitudes from 30,000 ft (4.36 psia) to 50,000 ft (1.68 psia). The model design was based on a 1 in. outer diameter, was 9 in. in overall length, and was mounted in the wind tunnel on a 3/8 in. wide support strut. For test conditions at 50,000 ft the strut was built to supply 90 psia of pressure, and to achieve 20 psia at the nozzle inlet with a maximum nozzle pressure of 52 psia. Instrumentation was developed to measure nozzle pressure ratio, and an external static pressure probe was designed to survey near field static pressure profiles at one nozzle diameter above the rig centerline. Model layout placed test nozzles between two transparent sidewalls in the 1x1 SWT for Schlieren photography and comparison to CFD analysis.

Hot streak characterization in serpentine exhaust nozzles

NASA Astrophysics Data System (ADS)

Crowe, Darrell S.

Modern aircraft of the United States Air Force face increasingly demanding cost, weight, and survivability requirements. Serpentine exhaust nozzles within an embedded engine allow a weapon system to fulfill mission survivability requirements by providing denial of direct line-of-sight into the high-temperature components of the engine. Recently, aircraft have experienced material degradation and failure along the aft deck due to extreme thermal loading. Failure has occurred in specific regions along the aft deck where concentrations of hot gas have come in contact with the surface causing hot streaks. The prevention of these failures will be aided by the accurate prediction of hot streaks. Additionally, hot streak prediction will improve future designs by identifying areas of the nozzle and aft deck surfaces that require thermal management. To this end, the goal of this research is to observe and characterize the underlying flow physics of hot streak phenomena. The goal is accomplished by applying computational fluid dynamics to determine how hot streak phenomena is affected by changes in nozzle geometry. The present research first validates the computational methods using serpentine inlet experimental and computational studies. A design methodology is then established for creating six serpentine exhaust nozzles investigated in this research. A grid independent solution is obtained on a nozzle using several figures of merit and the grid-convergence index method. An investigation into the application of a second-order closure turbulence model is accomplished. Simulations are performed for all serpentine nozzles at two flow conditions. The research introduces a set of characterization and performance parameters based on the temperature distribution and flow conditions at the nozzle throat and exit. Examination of the temperature distribution on the upper and lower nozzle surfaces reveals critical information concerning changes in hot streak phenomena due to changes in nozzle geometry.

Development of digital flow control system for multi-channel variable-rate sprayers

USDA-ARS?s Scientific Manuscript database

Precision modulation of nozzle flow rates is a critical step for variable-rate spray applications in orchards and ornamental nurseries. An automatic flow rate control system activated with microprocessors and pulse width modulation (PWM) controlled solenoid valves was developed to control flow rates...

Comparison of experimental surface pressures with theoretical predictions on twin two-dimensional convergent-divergent nozzles

NASA Technical Reports Server (NTRS)

Carlson, J. R.; Pendergraft, O. C., Jr.; Burley, J. R., II

1986-01-01

A three-dimensional subsonic aerodynamic panel code (VSAERO) was used to predict the effects of upper and lower external nozzle flap geometry on the external afterbody/nozzle pressure coefficient distributions and external nozzle drag of nonaxisymmetric convergent-divergent exhaust nozzles having parallel external sidewalls installed on a generic twin-engine high performance aircraft model. Nozzle static pressure coefficient distributions along the upper and lower surfaces near the model centerline and near the outer edges (corner) of the two surfaces were calculated, and nozzle drag was predicted using these surface pressure distributions. A comparison between the theoretical predictions and experimental wind tunnel data is made to evaluate the utility of the code in calculating the flow about these types of non-axisymmetric afterbody configurations. For free-stream Mach numbers of 0.60 and 0.90, the conditions where the flows were attached on the boattails yielded the best comparison between the theoretical predictions and the experimental data. For the Boattail terminal angles of greater than 15 deg., the experimental data for M = 0.60 and 0.90 indicated areas of separated flow, so the theoretical predictions failed to match the experimental data. Even though calculations of regions of separated flows are within the capabilities of the theoretical method, acceptable solutions were not obtained.

Study of unconventional propulsion system concepts for use in a long range transport

NASA Technical Reports Server (NTRS)

Champagne, G. A.

1973-01-01

The noise level of an uncoventional propulsion system for the next generation of subsonic, long-range transport aircraft is discussed. The desired noise level may be achieved by: (1) a fixed geometry, high bypass ratio turbofan with a geared two-stage fan and advanced acoustic treatment or (2) a moderate bypass ratio turbofan with a variable pitch two-stage fan, variable primary and duct nozzles, and advanced acoustic treatment. The geared fan system meets the noise goal with minimum economic penalty. Comparison of the noise levels at takeoff and landing in combination with the economic penalties required to achieve the lower noise levels at specific noise measuring stations, indicate that both area reduction and current certification prodedures should be used to ascertain the point of diminishing returns in establishing future noise goals.

Comparison of Rocket Performance using Exhaust Diffuser and Conventional Techniques for Altitude Simulation

NASA Technical Reports Server (NTRS)

Sivo, Joseph N.; Peters, Daniel J.

1959-01-01

A rocket engine with an exhaust-nozzle area ratio of 25 was operated at a constant chamber pressure of 600 pounds per square inch absolute over a range of oxidant-fuel ratios at an altitude pressure corresponding to approximately 47,000 feet. At this condition, the nozzle flow is slightly underexpanded as it leaves the nozzle. The altitude simulation was obtained first through the use of an exhaust diffuser coupled with the rocket engine and secondly, in an altitude test chamber where separate exhauster equipment provided the altitude pressure. A comparison of performance data from these two tests has established that a diffuser used with a rocket engine operating at near-design nozzle pressure ratio can be a valid means of obtaining altitude performance data for rocket engines.

Design of a Mach-15 Total-Enthalpy Nozzle With Non-uniform Inflow Using Rotational MOC

NASA Technical Reports Server (NTRS)

Gaffney, Richard L., Jr.

2004-01-01

A new computer program to design nozzles with non-uniform inflow has been developed using the rotational method of characteristics (MOC). This program has been used to design a nozzle for the NASA's HYPULSE shock-expansion tunnel for use in scramjet engine tests at a Mach-15 flight-enthalpy condition. The nozzle has an area ratio of 9.5:1 that expands the inflow from Mach 6 along the centerline to Mach 8.7. Although the density and Mach number vary radially at the exit due to the non-uniformities of the inflow, the MOC procedure produces exit flow that is parallel and has uniform static pressure. The design has been verified with CFD which compares favorably with the MOC solution.

Static noise tests on augmentor wing jet STOL research aircraft (C8A Buffalo)

NASA Technical Reports Server (NTRS)

Marrs, C. C.; Harkonen, D. L.; Okeefe, J. V.

1974-01-01

Results are presented for full scale ground static acoustic tests of over-area conical nozzles and a lobe nozzle installed on the Augmentor Wing Jet STOL Research Aircraft, a modified C8A Buffalo. The noise levels and spectrums of the test nozzles are compared against those of the standard conical nozzle now in use on the aircraft. Acoustic evaluations at 152 m (500 ft), 304 m (1000 ft), and 1216 m (4000 ft) are made at various engine power settings with the emphasis on approach and takeoff power. Appendix A contains the test log and propulsion calculations. Appendix B gives the original test plan, which was closely adhered to during the test. Appendix C describes the acoustic data recording and reduction systems, with calibration details.

Ultra low injection angle fuel holes in a combustor fuel nozzle

DOEpatents

York, William David

2012-10-23

A fuel nozzle for a combustor includes a mixing passage through which fluid is directed toward a combustion area and a plurality of swirler vanes disposed in the mixing passage. Each swirler vane of the plurality of swirler vanes includes at least one fuel hole through which fuel enters the mixing passage in an injection direction substantially parallel to an outer surface of the plurality of swirler vanes thereby decreasing a flameholding tendency of the fuel nozzle. A method of operating a fuel nozzle for a combustor includes flowing a fluid through a mixing passage past a plurality of swirler vanes and injecting a fuel into the mixing passage in an injection direction substantially parallel to an outer surface of the plurality of swirler vanes.

Exhaust Plume Effects on Sonic Boom for a Delta Wing and a Swept Wing-Body Model

NASA Technical Reports Server (NTRS)

Castner, Raymond; Lake, Troy

2012-01-01

Supersonic travel is not allowed over populated areas due to the disturbance caused by the sonic boom. Research has been performed on sonic boom reduction and has included the contribution of the exhaust nozzle plume. Plume effect on sonic boom has progressed from the study of isolated nozzles to a study with four exhaust plumes integrated with a wing-body vehicle. This report provides a baseline analysis of the generic wing-body vehicle to demonstrate the effect of the nozzle exhaust on the near-field pressure profile. Reductions occurred in the peak-to-peak magnitude of the pressure profile for a swept wing-body vehicle. The exhaust plumes also had a favorable effect as the nozzles were moved outward along the wing-span.

Corrugated and Composite Nozzle-Inlets for Thrust and Noise Benefits

NASA Technical Reports Server (NTRS)

Gilinsky, M.; Blankson, I. M.; Gromov, V. G.; Sakharov, V. I.

2004-01-01

The following research results are based on development of an approach previously proposed and investigated in for optimum nozzle design to obtain maximum thrust. The design was denoted a Telescope nozzle. A Telescope nozzle contains one or several internal designs, which are inserted at certain locations into a divergent conical or planar main nozzle near its exit. Such a design provides additional thrust augmentation over 20% by comparison with the optimum single nozzle of equivalent lateral area, What is more, experimental acoustic tests have discovered an essential noise reduction due to application of Telescope nozzles. In this paper, some additional theoretical results are presented for Telescope nozzles and a similar approach is applied for aero-performance improvement of a supersonic inlet. Numerical simulations were conducted for supersonic flow into the divergent portion of a 2D or axisymmetric nozzle with several plane or conical designs as well as into a 2D or axisymmetric supersonic inlet with a forebody. The Kryko-Godunov marching numerical scheme for inviscid supersonic flows was used. Several cases were tested using the NASA CFL3d and IM/MSU Russian codes based on the full Navier-Stokes equations. Numerical simulations were conducted for non reacting flows (both codes) as well as for real high temperature gas flows with non-equilibrium chemical reactions (the latter code). In general, these simulations have confirmed essential benefits of Telescope design applications in propulsion system. Some preliminary numerical simulations of several typical inlet designs were conducted with the goal of inlet design optimization for maneuvering flight conditions.

Computer Tomography Analysis of Fastrac Composite Thrust Chamber Assemblies

NASA Technical Reports Server (NTRS)

Beshears, Ronald D.

2000-01-01

Computed tomography (CT) inspection has been integrated into the production process for NASA's Fastrac composite thrust chamber assemblies (TCAs). CT has been proven to be uniquely qualified to detect the known critical flaw for these nozzles, liner cracks that are adjacent to debonds between the liner and overwrap. CT is also being used as a process monitoring tool through analysis of low density indications in the nozzle overwraps. 3d reconstruction of CT images to produce models of flawed areas is being used to give program engineers better insight into the location and nature of nozzle flaws.

Noise, Turbulence, and Thrust of Subsonic Free Jets from Lobed Nozzles

NASA Technical Reports Server (NTRS)

Zaman, K. B. M. Q.; Wang, F. Y.

2002-01-01

A study of noise benefit, vis-a-vis thrust penalty, and its correlation to turbulence intensities was conducted for free jets issuing from lobed nozzles. Four convergent nozzles with constant exit area were used in the experiments. Three of these were of rectangular lobed configuration having six, ten and fourteen lobes; the fourth was a circular nozzle. Increasing the number of lobes resulted in a progressive reduction in the turbulence intensities as well as in the overall radiated noise. The noise reduction was pronounced at the low frequency end of the spectrum. However, there was an increase in the high frequency noise that rendered the overall benefit less attractive when compared on a scaled-up A-weighted basis. A reduction in noise was accompanied by a commensurate reduction in the turbulent kinetic energy in the flow field. As expected, increasing the number of lobes involved progressive reduction in the thrust coefficient. Among the cases studied, the six-lobed nozzle had the optimum reduction in turbulence and noise with the least thrust penalty.

Static Internal Performance of a Two-Dimensional Convergent-Divergent Nozzle with External Shelf

NASA Technical Reports Server (NTRS)

Lamb, Milton; Taylor, John G.; Frassinelli, Mark C.

1996-01-01

An investigation was conducted in the static test facility of the Langley 16-Foot Transonic Tunnel to determine the internal performance of a two-dimensional convergent-divergent nozzle. The nozzle design was tested with dry and afterburning throat areas, which represent different power settings and three expansion ratios. For each of these configurations, three trailing-edge geometries were tested. The baseline geometry had a straight trailing edge. Two different shaping techniques were applied to the baseline nozzle design to reduce radar observables: the scarfed design and the sawtooth design. A flat plate extended downstream of the lower divergent flap trailing edge parallel to the model centerline to form a shelf-like expansion surface. This shelf was designed to shield the plume from ground observation (infrared radiation (IR) signature suppression). The shelf represents the part of the aircraft structure that might be present in an installed configuration. These configurations were tested at nozzle pressure ratios from 2.0 to 12.0.

VCE testbed program planning and definition study

NASA Technical Reports Server (NTRS)

Westmoreland, J. S.; Godston, J.

1978-01-01

The flight definition of the Variable Stream Control Engine (VSCE) was updated to reflect design improvements in the two key components: (1) the low emissions duct burner, and (2) the coannular exhaust nozzle. The testbed design was defined and plans for the overall program were formulated. The effect of these improvements was evaluated for performance, emissions, noise, weight, and length. For experimental large scale testing of the duct burner and coannular nozzle, a design definition of the VCE testbed configuration was made. This included selecting the core engine, determining instrumentation requirements, and selecting the test facilities, in addition to defining control system and assembly requirements. Plans for a comprehensive test program to demonstrate the duct burner and nozzle technologies were formulated. The plans include both aeroacoustic and emissions testing.

Supersonic propulsion technology. [variable cycle engines

NASA Technical Reports Server (NTRS)

Powers, A. G.; Coltrin, R. E.; Stitt, L. E.; Weber, R. J.; Whitlow, J. B., Jr.

1979-01-01

Propulsion concepts for commercial supersonic transports are discussed. It is concluded that variable cycle engines, together with advanced supersonic inlets and low noise coannular nozzles, provide good operating performance for both supersonic and subsonic flight. In addition, they are reasonably quiet during takeoff and landing and have acceptable exhaust emissions.

Hydrazine Blending and Storage Facility Wastewater Treatment and Decommissioning Assessment. Version 3.1

DTIC Science & Technology

1988-06-01

Valve Pit Number 2 Location One near hydrazine/aerozine tank area .nd one near wastewater tank area *There is a variety of underground piping at the...loading station (wipe of drum filling nozzles/connectors) 3,475 19.u I W-1U Tank HAS-?, drain value (Tank pit valve ) <S ɘ.2 W-110 Tank HAS-I, control... valve (on top) sample bottle broken W-111 Tank Truck Station, Truck loading filler nozzle and boom sample bottle broken I/ UUMH - 1,1

Warm fog dissipation using large volume water sprays

NASA Technical Reports Server (NTRS)

Keller, Vernon W. (Inventor)

1988-01-01

To accomplish the removal of warm fog about an area such as an airport runway, a plurality of nozzles along a line adjacent the area propelled water jets through the fog to heights of approximately twenty-five meters. Each water jet breaks up forming a water drop size distribution that falls through the fog overtaking, colliding, and coalescing with individual fog droplets and thereby removes the fog. A water retrieval system is used to collect the water and return it to reservoirs for pumping it to the nozzles once again.

A Passive Cavity Concept for Improving the Off-Design Performance of Fixed-Geometry Exhaust Nozzles

NASA Technical Reports Server (NTRS)

Asbury, Scott C.; Gunther, Christopher L.; Hunter, Craig A.

1996-01-01

An investigation was conducted in the model preparation area of the Langley 16-Foot Transonic Tunnel to study a passive cavity concept for improving the off-design performance of fixed-geometry exhaust nozzles. Passive cavity ventilation (through a porous surface) was applied to divergent flap surfaces and tested at static conditions in a sub-scale, nonaxisymmetric, convergent-divergent nozzle. As part of a comprehensive investigation, force, moment and pressure measurements were taken and focusing schlieren flow visualization was obtained for a baseline configuration and D passive cavity configurations. All tests were conducted with no external flow and high-pressure air was used to simulate jet-exhaust flow at nozzle pressure ratios from 1.25 to approximately 9.50. Results indicate that baseline nozzle performance was dominated by unstable shock-induced boundary-layer separation at off-design conditions, which came about through the natural tendency of overexpanded exhaust flow to satisfy conservation requirements by detaching from the nozzle divergent flaps. Passive cavity ventilation added the ability to control off-design separation in the nozzle by either alleviating separation or encouraging stable separation of the exhaust flow. Separation alleviation offers potential for installed nozzle performance benefits by reducing drag at forward flight speeds, even though it may reduce off-design static thrust efficiency as much as 3.2 percent. Encouraging stable separation of the exhaust flow offers significant performance improvements at static, low NPR and low Mach number flight conditions by improving off-design static thrust efficiency as much as 2.8 percent. By designing a fixed-geometry nozzle with fully porous divergent flaps, where both cavity location and percent open porosity of the flaps could be varied, passive flow control would make it possible to improve off-design nozzle performance across a wide operating range. In addition, the ability to encourage separation on one flap while alleviating it on the other makes it possible to generate thrust vectoring in the nozzle through passive flow control.

Calculation of Propulsive Nozzle Flowfields in Multidiffusing Chemically Reacting Environments. Ph.D. Thesis - Purdue Univ.

NASA Technical Reports Server (NTRS)

Kacynski, Kenneth John

1994-01-01

An advanced engineering model has been developed to aid in the analysis and design of hydrogen/oxygen chemical rocket engines. The complete multispecies, chemically reacting and multidiffusing Navier-Stokes equations are modelled, including the Soret thermal diffusion and the Dufour energy transfer terms. In addition to the spectrum of multispecies aspects developed, the model developed in this study is also conservative in axisymmetric flow for both inviscid and viscous flow environments and the boundary conditions employ a viscous, chemically reacting, reference plane characteristics method. Demonstration cases are presented for a 1030:1 area ratio nozzle, a 25 lbf film cooled nozzle, and a transpiration cooled plug and spool rocket engine. The results indicate that the thrust coefficient predictions of the 1030:1 and the 25 lbf film cooled nozzle are within 0.2 to 0.5 percent, respectively, of experimental measurements when all of the chemical reaction and diffusion terms are considered. Further, the model's predictions agree very well with the heat transfer measurements made in all of the nozzle test cases. The Soret thermal diffusion term is demonstrated to have a significant effect on the predicted mass fraction of hydrogen along the wall of the nozzle in both the laminar flow 1030:1 nozzle and the turbulent flow plug and spool nozzle analysis cases performed. Further, the Soret term was shown to represent an important fraction of the diffusion fluxes occurring in a transpiration cooled rocket engine.

Status of flow separation prediction in liquid propellant rocket nozzles

NASA Technical Reports Server (NTRS)

Schmucker, R. H.

1974-01-01

Flow separation which plays an important role in the design of a rocket engine nozzle is discussed. For a given ambient pressure, the condition of no flow separation limits the area ratio and, therefore, the vacuum performance. Avoidance of performance loss due to area ratio limitation requires a correct prediction of the flow separation conditions. To provide a better understanding of the flow separation process, the principal behavior of flow separation in a supersonic overexpanded rocket nozzle is described. The hot firing separation tests from various sources are summarized, and the applicability and accuracy of the measurements are described. A comparison of the different data points allows an evaluation of the parameters that affect flow separation. The pertinent flow separation predicting methods, which are divided into theoretical and empirical correlations, are summarized and the numerical results are compared with the experimental points.

The utilization of an infrared imaging system as a cooling slot blockage detector in the inspection of a transpiration cooled nozzle

NASA Technical Reports Server (NTRS)

Borg, Stephen E.; Wright, Robert E., Jr.; Alderfer, David W.; Whipple, Janet C.

1990-01-01

A comprehensive examination of the 8 foot temperature tunnel's transpiration cooled nozzle was completed using an infrared imaging radiometer to locate regions of cooling flow irregularities caused by obstruction of three or more adjacent cooling slots. Restrictions in the cooling flow were found and cataloged. Blockages found were due primarily to the presence of residual phosphoric acid being discharged from some of the cooling slots. This acid was used during construction of the nozzle components and was to have been purged prior to its delivery to the NASA Langley Research Center (LaRC). In addition, a radial displacement of one selection of discs located in the spool piece was inspected and cataloged for future reference. There did not seem to be a serious restriction of flow in this defect, but evidence from the infrared images indicated reduced slot activity within the gouge. The radiometer survey uncovered regions where closer inspection is recommended but did not cover the entire surface area of the three nozzle subsections due to equipment limitations. A list of areas with suspected problems is included in Appendix A.

Noise of Embedded High Aspect Ratio Nozzles

NASA Technical Reports Server (NTRS)

Bridges, James E.

2011-01-01

A family of high aspect ratio nozzles were designed to provide a parametric database of canonical embedded propulsion concepts. Nozzle throat geometries with aspect ratios of 2:1, 4:1, and 8:1 were chosen, all with convergent nozzle areas. The transition from the typical round duct to the rectangular nozzle was designed very carefully to produce a flow at the nozzle exit that was uniform and free from swirl. Once the basic rectangular nozzles were designed, external features common to embedded propulsion systems were added: extended lower lip (a.k.a. bevel, aft deck), differing sidewalls, and chevrons. For the latter detailed Reynolds-averaged Navier-Stokes (RANS) computational fluid dynamics (CFD) simulations were made to predict the thrust performance and to optimize parameters such as bevel length, and chevron penetration and azimuthal curvature. Seventeen of these nozzles were fabricated at a scale providing a 2.13 inch diameter equivalent area throat." ! The seventeen nozzles were tested for far-field noise and a few data were presented here on the effect of aspect ratio, bevel length, and chevron count and penetration. The sound field of the 2:1 aspect ratio rectangular jet was very nearly axisymmetric, but the 4:1 and 8:1 were not, the noise on their minor axes being louder than the major axes. Adding bevel length increased the noise of these nozzles, especially on their minor axes, both toward the long and short sides of the beveled nozzle. Chevrons were only added to the 2:1 rectangular jet. Adding 4 chevrons per wide side produced some decrease at aft angles, but increased the high frequency noise at right angles to the jet flow. This trend increased with increasing chevron penetration. Doubling the number of chevrons while maintaining their penetration decreased these effects. Empirical models of the parametric effect of these nozzles were constructed and quantify the trends stated above." Because it is the objective of the Supersonics Project that future design work be done more by physics-based computations and less by experiments, several codes under development were evaluated against these test cases. Preliminary results show that the RANS-based code JeNo predicts the spectral directivity of the low aspect ratio jets well, but has no capability to predict the non-axisymmetry. An effort to address this limitations, used in the RANS-based code of Leib and Goldstein, overpredicted the impact of aspect ratio. The broadband shock noise code RISN, also limited to axisymmetric assumptions, did a good job of predicting the spectral directivity of underexpanded 2:1 cold jet case but was not as successful on high aspect ratio jets, particularly when they are hot. All results are preliminary because the underlying CFD has not been validated yet. An effort using a Large Eddy Simulation code by Stanford University predicted noise that agreed with experiments to within a few dB.

Transient Three-Dimensional Side Load Analysis of a Film Cooled Nozzle

NASA Technical Reports Server (NTRS)

Wang, Ten-See; Guidos, Mike

2008-01-01

Transient three-dimensional numerical investigations on the side load physics for an engine encompassing a film cooled nozzle extension and a regeneratively cooled thrust chamber, were performed. The objectives of this study are to identify the three-dimensional side load physics and to compute the associated aerodynamic side load using an anchored computational methodology. The computational methodology is based on an unstructured-grid, pressure-based computational fluid dynamics formulation, and a transient inlet history based on an engine system simulation. Ultimately, the computational results will be provided to the nozzle designers for estimating of effect of the peak side load on the nozzle structure. Computations simulating engine startup at ambient pressures corresponding to sea level and three high altitudes were performed. In addition, computations for both engine startup and shutdown transients were also performed for a stub nozzle, operating at sea level. For engine with the full nozzle extension, computational result shows starting up at sea level, the peak side load occurs when the lambda shock steps into the turbine exhaust flow, while the side load caused by the transition from free-shock separation to restricted-shock separation comes at second; and the side loads decreasing rapidly and progressively as the ambient pressure decreases. For the stub nozzle operating at sea level, the computed side loads during both startup and shutdown becomes very small due to the much reduced flow area.

A New 3D Printing Strategy by Harnessing Deformation, Instability, and Fracture of Viscoelastic Inks.

PubMed

Yuk, Hyunwoo; Zhao, Xuanhe

2018-02-01

Direct ink writing (DIW) has demonstrated great potential as a multimaterial multifunctional fabrication method in areas as diverse as electronics, structural materials, tissue engineering, and soft robotics. During DIW, viscoelastic inks are extruded out of a 3D printer's nozzle as printed fibers, which are deposited into patterns when the nozzle moves. Hence, the resolution of printed fibers is commonly limited by the nozzle's diameter, and the printed pattern is limited by the motion paths. These limits have severely hampered innovations and applications of DIW 3D printing. Here, a new strategy to exceed the limits of DIW 3D printing by harnessing deformation, instability, and fracture of viscoelastic inks is reported. It is shown that a single nozzle can print fibers with resolution much finer than the nozzle diameter by stretching the extruded ink, and print various thickened or curved patterns with straight nozzle motions by accumulating the ink. A quantitative phase diagram is constructed to rationally select parameters for the new strategy. Further, applications including structures with tunable stiffening, 3D structures with gradient and programmable swelling properties, all printed with a single nozzle are demonstrated. The current work demonstrates that the mechanics of inks plays a critical role in developing 3D printing technology. © 2017 WILEY-VCH Verlag GmbH & Co. KGaA, Weinheim.

Multi-Objective Optimization of a Turbofan for an Advanced, Single-Aisle Transport

NASA Technical Reports Server (NTRS)

Berton, Jeffrey J.; Guynn, Mark D.

2012-01-01

Considerable interest surrounds the design of the next generation of single-aisle commercial transports in the Boeing 737 and Airbus A320 class. Aircraft designers will depend on advanced, next-generation turbofan engines to power these airplanes. The focus of this study is to apply single- and multi-objective optimization algorithms to the conceptual design of ultrahigh bypass turbofan engines for this class of aircraft, using NASA s Subsonic Fixed Wing Project metrics as multidisciplinary objectives for optimization. The independent design variables investigated include three continuous variables: sea level static thrust, wing reference area, and aerodynamic design point fan pressure ratio, and four discrete variables: overall pressure ratio, fan drive system architecture (i.e., direct- or gear-driven), bypass nozzle architecture (i.e., fixed- or variable geometry), and the high- and low-pressure compressor work split. Ramp weight, fuel burn, noise, and emissions are the parameters treated as dependent objective functions. These optimized solutions provide insight to the ultrahigh bypass engine design process and provide information to NASA program management to help guide its technology development efforts.

Magnetic Field Effects on Plasma Plumes

NASA Technical Reports Server (NTRS)

Ebersohn, F.; Shebalin, J.; Girimaji, S.; Staack, D.

2012-01-01

Here, we will discuss our numerical studies of plasma jets and loops, of basic interest for plasma propulsion and plasma astrophysics. Space plasma propulsion systems require strong guiding magnetic fields known as magnetic nozzles to control plasma flow and produce thrust. Propulsion methods currently being developed that require magnetic nozzles include the VAriable Specific Impulse Magnetoplasma Rocket (VASIMR) [1] and magnetoplasmadynamic thrusters. Magnetic nozzles are functionally similar to de Laval nozzles, but are inherently more complex due to electromagnetic field interactions. The two crucial physical phenomenon are thrust production and plasma detachment. Thrust production encompasses the energy conversion within the nozzle and momentum transfer to a spacecraft. Plasma detachment through magnetic reconnection addresses the problem of the fluid separating efficiently from the magnetic field lines to produce maximum thrust. Plasma jets similar to those of VASIMR will be studied with particular interest in dual jet configurations, which begin as a plasma loops between two nozzles. This research strives to fulfill a need for computational study of these systems and should culminate with a greater understanding of the crucial physics of magnetic nozzles with dual jet plasma thrusters, as well as astrophysics problems such as magnetic reconnection and dynamics of coronal loops.[2] To study this problem a novel, hybrid kinetic theory and single fluid magnetohydrodynamic (MHD) solver known as the Magneto-Gas Kinetic Method is used.[3] The solver is comprised of a "hydrodynamic" portion based on the Gas Kinetic Method and a "magnetic" portion that accounts for the electromagnetic behaviour of the fluid through source terms based on the resistive MHD equations. This method is being further developed to include additional physics such as the Hall effect. Here, we will discuss the current level of code development, as well as numerical simulation results

Systems Modeling of a Hypothetical SSME Channel-Wall Nozzle

NASA Technical Reports Server (NTRS)

Greene, William D.; Thames, Mignon P.; Polsgrove, Robert H.

2003-01-01

A future upgrade to the Space Shuttle Main Engine (SSME) may be the replacement of the current regenerative cooled tube-wall nozzle with a nozzle using a regeneratively-cooled channel-wall design. The current tube-wall design represents the only major piece of SSME hardware that has not been dramatically updated throughout thc long history of the engine. There are a number of advantages to a channel-wall design including the promise of faster and lower cost fabrication and greater reliability in the field. The technical obstacles in the path of making this happen are many, particularly in the realms of metallurgy and manufacturing techniques. However, one technical area that can and should be addressed in the near term as part of the development of detailed component requirements is a systems type model of the fluid flow and heat transfer processes to which the new design will be exposed. This paper presents the results of an effort to develop a mathematical model of the internal flow for a generic channel-wall nozzle functioning as a direct replacement for the current tube-wall nozzle with a minimum of systems-level changes. Comparisons will be made to mathematical modeling results for the current tube-wall design and the results of various geometrical trade studies will be presented. It is the intent of this work to examine the feasibility of the concept of a direct replacement component with minimum systems-!eve impacts and to highlight potential areas of concern requiring further work in the future.

Comparison of jet Mach number decay data with a correlation and jet spreading contours for a large variety of nozzles

NASA Technical Reports Server (NTRS)

Groesbeck, D. E.; Huff, R. G.; Vonglahn, U. H.

1977-01-01

Small-scale circular, noncircular, single- and multi-element nozzles with flow areas as large as 122 sq cm were tested with cold airflow at exit Mach numbers from 0.28 to 1.15. The effects of multi-element nozzle shape and element spacing on jet Mach number decay were studied in an effort to reduce the noise caused by jet impingement on externally blown flap (EBF) STOL aircraft. The jet Mach number decay data are well represented by empirical relations. Jet spreading and Mach number decay contours are presented for all configurations tested.

Theoretical Performance of Hydrogen-Oxygen Rocket Thrust Chambers

NASA Technical Reports Server (NTRS)

Sievers, Gilbert K.; Tomazic, William A.; Kinney, George R.

1961-01-01

Data are presented for liquid-hydrogen-liquid-oxygen thrust chambers at chamber pressures from 15 to 1200 pounds per square inch absolute, area ratios to approximately 300, and percent fuel from about 8 to 34 for both equilibrium and frozen composition during expansion. Specific impulse in vacuum, specific impulse, combustion-chamber temperature, nozzle-exit temperature, characteristic velocity, and the ratio of chamber-to-nozzle-exit pressure are included. The data are presented in convenient graphical forms to allow quick calculation of theoretical nozzle performance with over- or underexpansion, flow separation, and introduction of the propellants at various initial conditions or heat loss from the combustion chamber.

Effect of Jet-nozzle-expansion Ratio on Drag of Parabolic Afterbodies

NASA Technical Reports Server (NTRS)

Englert, Gerald W; Vargo, Donald J; Cubbison, Robert W

1954-01-01

The interaction of the flow from one convergent and two convergent-divergent nozzles on parabolic afterbodies was studied at free-stream Mach numbers of 2.0, 1.6, and 0.6 over a range of jet pressure ratio. The influence of the jet on boattail and base drag was very pronounced. Study of the total external afterbody drag values at supersonic speeds indicated that, over most of the high-pressure-ratio range, increasing the nozzle design expansion ratio increased the drag even though the boattail area was reduced. Increasing the pressure ratio tended to increase slightly the total-drag increment caused by angle-of-attack operation.

Coannular plug nozzle noise reduction and impact of exhaust system designs

NASA Technical Reports Server (NTRS)

Lee, R.

1976-01-01

Reducing the noise generated by high velocity jets has confronted engine designers and acoustics workers alike for the past fifteen years. Some of the jet noise suppressor configurations that are investigated are shown. With the exception of the early CJ-805 daisy suppressor nozzle which found successful application on the Convair 990 airplane, the others were developmental hardware at different stages of the effort in the past eight years - all aiming at potential supersonic cruise aircraft applications. Some significant progress was made as the result of work supported by NASA and FAA in the past two to three years. This work pertains to the concept demonstration and scale model testing of coannular plug nozzles with inverted velocity profile, and to the preliminary study of its application to advanced variable cycle engines (VCE) appropriate for supersonic cruise aircraft.

Single stage, low noise, advanced technology fan. Volume 1: Aerodynamic design

NASA Technical Reports Server (NTRS)

Sullivan, T. J.; Younghans, J. L.; Little, D. R.

1976-01-01

The aerodynamic design for a half-scale fan vehicle, which would have application on an advanced transport aircraft, is described. The single stage advanced technology fan was designed to a pressure ratio of 1.8 at a tip speed of 503 m/sec 11,650 ft/sec). The fan and booster components are designed in a scale model flow size convenient for testing with existing facility and vehicle hardware. The design corrected flow per unit annulus area at the fan face is 215 kg/sec sq m (44.0 lb m/sec sq ft) with a hub-tip ratio of 0.38 at the leading edge of the fan rotor. This results in an inlet corrected airflow of 117.9 kg/sec (259.9 lb m/sec) for the selected rotor tip diameter if 90.37 cm (35.58 in.). The variable geometry inlet is designed utilizing a combination of high throat Mach number and acoustic treatment in the inlet diffuser for noise suppression (hybrid inlet). A variable fan exhaust nozzle was assumed in conjunction with the variable inlet throat area to limit the required area change of the inlet throat at approach and hence limit the overall diffusion and inlet length. The fan exit duct design was primarily influenced by acoustic requirements, including length of suppressor wall treatment; length, thickness and position on a duct splitter for additional suppressor treatment; and duct surface Mach numbers.

Impact of Turbine Modulation on Variable-Cycle Engine Performance. Phase 4. Additional Hardware Design and Fabrication, Engine Modification, and Altitude Test. Part 3 B

DTIC Science & Technology

1974-12-01

urbofan engine performance. An AiKesearch Model TFE731 -2 Turbofan Engine was modified to incorporate production-type variable-geometry hardware...reliability was shown for the variable- geometry components. The TFE731 , modified to include variable geometry, proved to be an inexpensive...Atm at a Met Thrust of 3300 LBF 929 85 Variable-Cycle Engine TFE731 Exhaust-Nozzle Performance 948 86 Analytical Model Comparisons, Aerodynamic

Method of making a rocket nozzle

NASA Technical Reports Server (NTRS)

Campbell, D. H. (Inventor)

1969-01-01

A method is described for forming the interior of a nozzle having uneven walls so that a throat of smooth converging and diverging sides is provided for passing flow. A metallic insert material is placed within the flow passageway adjacent to the area where the sharper throat constriction is to be formed, so that the material will flow through the inlet into the throat space when liquefied.

Rotary and High-Pressure Nozzle Spray Plume Droplet Analysis For Aerially Applied Mosquito Adulticides: Laser Diffraction Characterization.

PubMed

Hornby, Jonathan A; Robinson, Jim; Sterling, Milton

2017-03-01

The droplet spectrum of a mosquito adulticide spray plume determines its ability to drift through the target area, impinge on the mosquito, deliver a toxic dose, and the risk of environmental contamination. This paper provides data on droplet spectra produced from 6 nozzles in a high-pressure nozzle spray system and 5 rotary nozzle systems for common mosquito adulticides. Spray plume spectra were measured by laser diffraction. High-pressure nozzles were evaluated at pressures ranging from 500 psi to 6,000 psi. Rotary nozzles were evaluated at rotational speeds ranging from 500 rpm to 24,000 rpm. Measurements were made at wind speeds of 129 km/h (80 mph) to 225 km/h (140 mph). Adulticides included were Fyfanon ® , Aqua-Reslin ® , Dibrom ® , Duet ® , Permanone ® , and the inert mineral oil, Orchex ® 796. High-pressure nozzles produced spray plumes within the US Environmental Protection Agency (EPA) label requirements for all configurations tested except for one at a wind speed of 225 km/h, BETE ® MW125. Air speed had no significant effect on the spray plume volume median diameter (Dv (0.5) ) at the speeds tested with Fyfanon ® . The spray plume 90% drop volume diameter (Dv (0.9) ) significantly decreased, 13% at the higher wind speed of 225 km/h. Drop size was inversely related to pressure. Dilution of the product formulations increased the Dv (0.5) of the spray plume but it did not exceed the label requirements. For the PJ15 nozzle, orientation of the nozzle into the wind of up to 135° showed a significant increase in Dv (0.5) at 500 psi, 750 psi, and 1,500 psi. The Dv (0.5) varied <5 μm over the 3 angles examined for any specific pressure. Rotary nozzles produced spray plumes within the EPA label requirements for all test configurations examined. Air speed had no significant effect on Dv (0.5) or Dv (0.9) of the plume at speeds tested with Fyfanon for the ASC A20 nozzle. The rotary AU5000 nozzle using Orchex 796 produced plumes of larger drops in all configurations than any of the rotary nozzles of similar configurations using active ingredient formulations and within EPA label requirements.

An overview of current Navy programs to develop thrust augmenting ejectors

NASA Technical Reports Server (NTRS)

Green, K. A.

1979-01-01

The primary objective of Navy sponsored research in thrust augmentation is the development of an improved augmenter for V/STOL application. In support of this goal, a data base is being established to provide an accurate prediction capability for use in ejector design. A general technology development of ejectors and associated effects presently is split into the more specific areas of lift and control, since thrust augmenting ejectors may be suitable for both. Research areas examined include advanced diffuser and end wall design; advanced primary nozzles; analytic studies; augmenting reaction controls; and nozzle design.

Combustion devices technology team - An overview and status of STME-related activities

NASA Technical Reports Server (NTRS)

Tucker, P. K.; Croteau-Gillespie, Margie

1992-01-01

The Consortium for CFD applications in propulsion technology has been formed at NASA/Marshall Space Flight Center. The combustion devices technology team is one of the three teams that constitute the Consortium. While generally aiming to advance combustion devices technology for rocket propulsion, the team's efforts for the last 1 and 1/2 years have been focused on issues relating to the Space Transportation Main Engine (STME) nozzle. The nozzle design uses hydrogen-rich turbine exhaust to cool the wall in a film/dump scheme. This method of cooling presents challenges and associated risks for the nozzle designers and the engine/vehicle integrators. Within the nozzle itself, a key concern is the ability to effectively and efficiently film cool the wall. From the National Launch System vehicle base standpoint, there are concerns with dumping combustible gases at the nozzle exit and their potential adverse effects on the base thermal environment. The Combustion Team has developed and is implementing plans to use validated CFD tools to aid in risk mitigation for both areas.

High-frequency, silicon-based ultrasonic nozzles using multiple Fourier horns.

PubMed

Tsai, Shirley C; Song, Yu L; Tseng, Terry K; Chou, Yuan F; Chen, Wei J; Tsai, Chen S

2004-03-01

This paper presents the design, simulation, and characterization of microfabricated 0.5 MHz, silicon-based, ultrasonic nozzles. Each nozzle is made of a piezoelectric drive section and a silicon resonator consisting of multiple Fourier horns, each with half wavelength design and twice amplitude magnification. Results of finite element three-dimensional (3-D) simulation using a commercial program predicted existence of one resonant frequency of pure longitudinal vibration. Both impedance analysis and measurement of longitudinal vibration confirmed the simulation results with one pure longitudinal vibration mode at the resonant frequency in excellent agreement with the design value. Furthermore, at the resonant frequency, the measured longitudinal vibration amplitude at the nozzle tip increases as the number of Fourier horns (n) increases in good agreement with the theoretical values of 2(n). Using this design, very high vibration amplitude gain at the nozzle tip can be achieved with no reduction in the tip cross-sectional area for contact of liquid to be atomized. Therefore, the required electric drive power should be drastically reduced, decreasing the likelihood of transducer failure in ultrasonic atomization.

Radial cold trap

DOEpatents

Grundy, Brian R.

1981-01-01

The radial cold trap comprises a housing having a plurality of mesh bands disposed therein. The mesh bands comprise concentrically arranged bands of mesh with the mesh specific surface area of each band increasing from the outermost mesh band to the innermost mesh band. An inlet nozzle is attached to the outside section of the housing while an outlet nozzle is attached to the inner portion of the housing so as to be concentrically connected to the innermost mesh band. An inlet baffle having orifices therein may be disposed around the outermost mesh band and within the housing for directing the flow of the fluid from the inlet nozzle to the outermost mesh band in a uniform manner. The flow of fluid passes through each consecutive mesh band and into the outlet nozzle. The circular pattern of the symmetrically arranged mesh packing allows for better utilization of the entire cold trap volume.

Radial cold trap

DOEpatents

Grundy, B.R.

1981-09-29

The radial cold trap comprises a housing having a plurality of mesh bands disposed therein. The mesh bands comprise concentrically arranged bands of mesh with the mesh specific surface area of each band increasing from the outermost mesh band to the innermost mesh band. An inlet nozzle is attached to the outside section of the housing while an outlet nozzle is attached to the inner portion of the housing so as to be concentrically connected to the innermost mesh band. An inlet baffle having orifices therein may be disposed around the outermost mesh band and within the housing for directing the flow of the fluid from the inlet nozzle to the outermost mesh band in a uniform manner. The flow of fluid passes through each consecutive mesh band and into the outlet nozzle. The circular pattern of the symmetrically arranged mesh packing allows for better utilization of the entire cold trap volume. 2 figs.

Quadratic curvature terms and deformed Schwarzschild-de Sitter black hole analogues in the laboratory

NASA Astrophysics Data System (ADS)

da Rocha, R.; Sobreiro, R. F.; Tomaz, A. A.

2017-12-01

Sound waves on a fluid stream, in a de Laval nozzle, are shown to correspond to quasinormal modes emitted by black holes that are physical solutions in a quadratic curvature gravity with cosmological constant. Sound waves patterns in transsonic regimes at a laboratory are employed here to provide experimental data regarding generalized theories of gravity, comprised by the exact de Sitter-like solution and a perturbative solution around the Schwarzschild-de Sitter standard solution as well. Using the classical tests of General Relativity to bound free parameters in these solutions, acoustic perturbations on fluid flows in nozzles are then regarded, to study quasinormal modes of these black holes solutions, providing deviations of the de Laval nozzle cross-sectional area, when compared to the Schwarzschild solution. The fluid sonic point in the nozzle, for sound waves in the fluid, is shown to implement the acoustic event horizon corresponding to quasinormal modes.

Shock unsteadiness in a thrust optimized parabolic nozzle

NASA Astrophysics Data System (ADS)

Verma, S. B.

2009-07-01

This paper discusses the nature of shock unsteadiness, in an overexpanded thrust optimized parabolic nozzle, prevalent in various flow separation modes experienced during start up {(δ P0 /δ t > 0)} and shut down {(δ P0/δ t < 0)} sequences. The results are based on simultaneously acquired data from real-time wall pressure measurements using Kulite pressure transducers, high-speed schlieren (2 kHz) of the exhaust flow-field and from strain-gauges installed on the nozzle bending tube. Shock unsteadiness in the separation region is seen to increase significantly just before the onset of each flow transition, even during steady nozzle operation. The intensity of this measure ( rms level) is seen to be strongly influenced by relative locations of normal and overexpansion shock, the decrease in radial size of re-circulation zone in the back-flow region, and finally, the local nozzle wall contour. During restricted shock separation, the pressure fluctuations in separation region exhibit periodic characteristics rather than the usually observed characteristics of intermittent separation. The possible physical mechanisms responsible for the generation of flow unsteadiness in various separation modes are discussed. The results are from an experimental study conducted in P6.2 cold-gas subscale test facility using a thrust optimized parabolic nozzle of area-ratio 30.

Weed Control Sprayers: Calibration and Maintenance. Special Circular 81.

ERIC Educational Resources Information Center

Myers, Arthur L.

This manual covers aspects of calibration and maintenance of weed control sprayers including variables affecting application rate, the pre-calibration check, calculations, band spraying, nozzle tip selection, agitation, and cleaning. (BB)

The variable polarity plasma arc welding process: Characteristics and performance

NASA Technical Reports Server (NTRS)

Hung, R. J.; Zhu, G. J.

1991-01-01

Significant advantages of the Variable Polarity Plasma Arc (VPPA) Welding Process include faster welding, fewer repairs, less joint preparation, reduced weldment distortion, and absence of porosity. The power distribution was analyzed for an argon plasma gas flow constituting the fluid in the VPPA Welding Process. The major heat loss at the torch nozzle is convective heat transfer; in the space between the outlet of the nozzle and the workpiece; radiative heat transfer; and in the keyhole in the workpiece, convective heat transfer. The power absorbed at the workpiece produces the molten puddle that solidifies into the weld bead. Crown and root widths, and crown and root heights of the weld bead are predicted. The basis is provided for an algorithm for automatic control of VPPA welding machine parameters to obtain desired weld bead dimensions.

Low-Thrust Bipropellant Engine Technology.

DTIC Science & Technology

1980-08-01

Non-Destructive Testing OD Outside Diameter xv tr. GLOSSARY (cont.J ODE One Dimensional Equilibrium ODK One Dimensional Kinetics Pc Thrust Chamber...performance (280 sec steady- state, 220 sec pulsing) have not yet been collectively achieved, but should be obtainable with further development activities...even at nozzle area ratios up to 400:1. The influence of nozzle kinetics (i.e., equilibrium versus frozen flow and ODK ) are noted to be a much more

Liquid Engine Design: Effect of Chamber Dimensions on Specific Impulse

NASA Technical Reports Server (NTRS)

Hoggard, Lindsay; Leahy, Joe

2009-01-01

Which assumption of combustion chemistry - frozen or equilibrium - should be used in the prediction of liquid rocket engine performance calculations? Can a correlation be developed for this? A literature search using the LaSSe tool, an online repository of old rocket data and reports, was completed. Test results of NTO/Aerozine-50 and Lox/LH2 subscale and full-scale injector and combustion chamber test results were found and studied for this task. NASA code, Chemical Equilibrium with Applications (CEA) was used to predict engine performance using both chemistry assumptions, defined here. Frozen- composition remains frozen during expansion through the nozzle. Equilibrium- instantaneous chemical equilibrium during nozzle expansion. Chamber parameters were varied to understand what dimensions drive chamber C* and Isp. Contraction Ratio is the ratio of the nozzle throat area to the area of the chamber. L is the length of the chamber. Characteristic chamber length, L*, is the length that the chamber would be if it were a straight tube and had no converging nozzle. Goal: Develop a qualitative and quantitative correlation for performance parameters - Specific Impulse (Isp) and Characteristic Velocity (C*) - as a function of one or more chamber dimensions - Contraction Ratio (CR), Chamber Length (L ) and/or Characteristic Chamber Length (L*). Determine if chamber dimensions can be correlated to frozen or equilibrium chemistry.

Effect of flame-tube head structure on combustion chamber performance

NASA Technical Reports Server (NTRS)

Gu, Minqqi

1986-01-01

The experimental combustion performance of a premixed, pilot-type flame tube with various head structures is discussed. The test study covers an extensive area: efficiency of the combustion chamber, quality of the outlet temperature field, limit of the fuel-lean blowout, ignition performance at ground starting, and carbon deposition. As a result of these tests, a nozzle was found which fits the premixed pilot flame tube well. The use of this nozzle optimized the performance of the combustion chamber. The tested models had premixed pilot chambers with two types of air-film-cooling structures, six types of venturi-tube structures, and secondary fuel nozzles with two small spray-cone angles.

Navier-Stokes calculations of scramjet-nozzle-afterbody flowfields

NASA Technical Reports Server (NTRS)

Baysal, Oktay

1991-01-01

A comprehensive computational fluid dynamics effort was conducted from 1987 to 1990 to properly design a nozzle and lower aft end of a generic hypersonic vehicle powered by a scramjet engine. The interference of the exhaust on the control surfaces of the vehicle can have adverse effects on its stability. Two-dimensional Navier-Stokes computations were performed, where the exhaust gas was assumed to be air behaving as a perfect gas. Then the exhaust was simulated by a mixture of Freon-12 and argon, which required solving the Navier-Stokes equations for four species: (nitrogen, oxygen, Freon-12, and argon). This allowed gamma to be a field variable during the mixing of the multispecies gases. Two different mixing models were used and comparisons between them as well as the perfect gas air calculations were made to assess their relative merits. Finally, the three dimensional Navier-Stokes computations were made for the full-span scramjet nozzle afterbody module.

Navier-Stokes calculations of scramjet-nozzle-afterbody flowfields

NASA Astrophysics Data System (ADS)

Baysal, Oktay

1991-07-01

A comprehensive computational fluid dynamics effort was conducted from 1987 to 1990 to properly design a nozzle and lower aft end of a generic hypersonic vehicle powered by a scramjet engine. The interference of the exhaust on the control surfaces of the vehicle can have adverse effects on its stability. Two-dimensional Navier-Stokes computations were performed, where the exhaust gas was assumed to be air behaving as a perfect gas. Then the exhaust was simulated by a mixture of Freon-12 and argon, which required solving the Navier-Stokes equations for four species: (nitrogen, oxygen, Freon-12, and argon). This allowed gamma to be a field variable during the mixing of the multispecies gases. Two different mixing models were used and comparisons between them as well as the perfect gas air calculations were made to assess their relative merits. Finally, the three dimensional Navier-Stokes computations were made for the full-span scramjet nozzle afterbody module.

A computational model for the prediction of jet entrainment in the vicinity of nozzle boattails (the BOAT code)

NASA Technical Reports Server (NTRS)

Dash, S. M.; Pergament, H. S.

1978-01-01

The development of a computational model (BOAT) for calculating nearfield jet entrainment, and its incorporation in an existing methodology for the prediction of nozzle boattail pressures, is discussed. The model accounts for the detailed turbulence and thermochemical processes occurring in the mixing layer formed between a jet exhaust and surrounding external stream while interfacing with the inviscid exhaust and external flowfield regions in an overlaid, interactive manner. The ability of the BOAT model to analyze simple free shear flows is assessed by comparisons with fundamental laboratory data. The overlaid procedure for incorporating variable pressures into BOAT and the entrainment correction employed to yield an effective plume boundary for the inviscid external flow are demonstrated. This is accomplished via application of BOAT in conjunction with the codes comprising the NASA/LRC patched viscous/inviscid methodology for determining nozzle boattail drag for subsonic/transonic external flows.

Large-Eddy Simulations of Noise Generation in Supersonic Jets at Realistic Engine Temperatures

NASA Astrophysics Data System (ADS)

Liu, Junhui; Corrigan, Andrew; Kailasanath, K.; Taylor, Brian

2015-11-01

Large-eddy simulations (LES) have been carried out to investigate the noise generation in highly heated supersonic jets at temperatures similar to those observed in high-performance jet engine exhausts. It is found that the exhaust temperature of high-performance jet engines can range from 1000K at an intermediate power to above 2000K at a maximum afterburning power. In low-temperature jets, the effects of the variation of the specific heat ratio as well as the radial temperature profile near the nozzle exit are small and are ignored, but it is not clear whether those effects can be also ignored in highly heated jets. The impact of the variation of the specific heat ratio is assessed by comparing LES results using a variable specific heat ratio with those using a constant specific heat ratio. The impact on both the flow field and the noise distributions are investigated. Because the total temperature near the nozzle wall can be substantially lower than the nozzle total temperature either due to the heating loss through the nozzle wall or due to the cooling applied near the wall, this lower wall temperature may impact the temperature in the shear layer, and thus impact the noise generation. The impact of the radial temperature profile on the jet noise generation is investigated by comparing results of lower nozzle wall temperatures with those of the adiabatic wall condition.

Effect of formulation- and administration-related variables on deposition pattern of nasal spray pumps evaluated using a nasal cast.

PubMed

Kundoor, Vipra; Dalby, Richard N

2011-08-01

To systematically evaluate the effect of formulation- and administration-related variables on nasal spray deposition using a nasal cast. Deposition pattern was assessed by uniformly coating a transparent nose model with Sar-Gel®, which changes from white to purple on contact with water. Sprays were subsequently discharged into the cast, which was then digitally photographed. Images were quantified using Adobe® Photoshop. The effects of formulation viscosity (which influences droplet size), simulated administration techniques (head orientation, spray administration angle, spray nozzle insertion depth), spray pump design and metering volume on nasal deposition pattern were investigated. There was a significant decrease in the deposition area associated with sprays of increasing viscosity. This appeared to be mediated by an increase in droplet size and a narrowing of the spray plume. Administration techniques and nasal spray pump design also had a significant effect on the deposition pattern. This simple color-based method provides quantitative estimates of the effects that different formulation and administration variables may have on the nasal deposition area, and provides a rational basis on which manufacturers of nasal sprays can base their patient instructions or post approval changes when it is impractical to optimize these using a clinical study.

Subscale Carbon-Carbon Nozzle Extension Development and Hot Fire Testing in Support of Upper Stage Liquid Rocket Engines

NASA Technical Reports Server (NTRS)

Gradl, Paul; Valentine, Peter; Crisanti, Matthew; Greene, Sandy Elam

2016-01-01

Upper stage and in-space liquid rocket engines are optimized for performance through the use of high area ratio nozzles to fully expand combustion gases to low exit pressures increasing exhaust velocities. Due to the large size of such nozzles and the related engine performance requirements, carbon-carbon (C/C) composite nozzle extensions are being considered for use in order to reduce weight impacts. NASA and industry partner Carbon-Carbon Advanced Technologies (C-CAT) are working towards advancing the technology readiness level of large-scale, domestically-fabricated, C/C nozzle extensions. These C/C extensions have the ability to reduce the overall costs of extensions relative to heritage metallic and composite extensions and to decrease weight by 50%. Material process and coating developments have advanced over the last several years, but hot fire testing to fully evaluate C/C nozzle extensions in relevant environments has been very limited. NASA and C-CAT have designed, fabricated and hot fire tested multiple subscale nozzle extension test articles of various C/C material systems, with the goal of assessing and advancing the manufacturability of these domestically producible materials as well as characterizing their performance when subjected to the typical environments found in a variety of liquid rocket and scramjet engines. Testing at the MSFC Test Stand 115 evaluated heritage and state-of-the-art C/C materials and coatings, demonstrating the capabilities of the high temperature materials and their fabrication methods. This paper discusses the design and fabrication of the 1.2k-lbf sized carbon-carbon nozzle extensions, provides an overview of the test campaign, presents results of the hot fire testing, and discusses potential follow-on development work.

Comparison of theoretical and experimental thrust performance of a 1030:1 area ratio rocket nozzle at a chamber pressure of 2413 kN/sq m (350 psia)

NASA Technical Reports Server (NTRS)

Smith, Tamara A.; Pavli, Albert J.; Kacynski, Kenneth J.

1987-01-01

The Joint Army, Navy, NASA, Air Force (JANNAF) rocket-engine performance-prediction procedure is based on the use of various reference computer programs. One of the reference programs for nozzle analysis is the Two-Dimensional Kinetics (TDK) Program. The purpose of this report is to calibrate the JANNAF procedure that has been incorporated into the December 1984 version of the TDK program for the high-area-ratio rocket-engine regime. The calibration was accomplished by modeling the performance of a 1030:1 rocket nozzle tested at NASA Lewis. A detailed description of the test conditions and TDK input parameters is given. The reuslts indicate that the computer code predicts delivered vacuum specific impulse to within 0.12 to 1.9 percent of the experimental data. Vacuum thrust coefficient predictions were within + or - 1.3 percent of experimental results. Predictions of wall static pressure were within approximately + or - 5 percent of the measured values.

An application of digital image processing techniques to the characterization of liquid petroleum gas (LPG) spray

NASA Astrophysics Data System (ADS)

Qi, Y. L.; Xu, B. Y.; Cai, S. L.

2006-12-01

To control fuel injection, optimize combustion and reduce emissions for LPG (liquefied petroleum gas) engines, it is necessary and important to understand the characteristics of LPG sprays. The present work investigates the geometry of LPG sprays, including spray tip penetration, spray angle, projected spray area and spray volume, by using schlieren photography and digital image processing techniques. Two types of single nozzle injectors were studied, with the same nozzle diameter, but one with and one without a double-hole flow-split head. A code developed to analyse the results directly from the digitized images is shown to be more accurate and efficient than manual measurement and analysis. Test results show that a higher injection pressure produces a longer spray tip penetration, a larger projected spray area and spray volume, but a smaller spray cone angle. The injector with the double-hole split-head nozzle produces better atomization and shorter tip penetration at medium and late injection times, but longer tip penetration in the early stage.

Preliminary design of a prototype particulate stack sampler. [For stack gas temperature under 300/sup 0/C

DOE Office of Scientific and Technical Information (OSTI.GOV)

Elder, J.C.; Littlefield, L.G.; Tillery, M.I.

1978-06-01

A preliminary design of a prototype particulate stack sampler (PPSS) has been prepared, and development of several components is under way. The objective of this Environmental Protection Agency (EPA)-sponsored program is to develop and demonstrate a prototype sampler with capabilities similar to EPA Method 5 apparatus but without some of the more troublesome aspects. Features of the new design include higher sampling flow; display (on demand) of all variables and periodic calculation of percent isokinetic, sample volume, and stack velocity; automatic control of probe and filter heaters; stainless steel surfaces in contact with the sample stream; single-point particle size separationmore » in the probe nozzle; null-probe capability in the nozzle; and lower weight in the components of the sampling train. Design considerations will limit use of the PPSS to stack gas temperatures under approximately 300/sup 0/C, which will exclude sampling some high-temperature stacks such as incinerators. Although need for filter weighing has not been eliminated in the new design, introduction of a variable-slit virtual impactor nozzle may eliminate the need for mass analysis of particles washed from the probe. Component development has shown some promise for continuous humidity measurement by an in-line wet-bulb, dry-bulb psychrometer.« less

Historical problem areas lessons learned

NASA Technical Reports Server (NTRS)

Sackheim, Bob; Fester, Dale A.

1991-01-01

Historical problem areas in space transportation propulsion technology are identified in viewgraph form. Problem areas discussed include materials compatibility, contamination, pneumatic/feed system flow instabilities, instabilities in rocket engine combustion and fuel sloshing, exhaust plume interference, composite rocket nozzle failure, and freeze/thaw damage.

Effect of chevron nozzle penetration on aero-acoustic characteristics of jet at M = 0.8

NASA Astrophysics Data System (ADS)

Nikam, S. R.; Sharma, S. D.

2017-12-01

Aero-acoustic characteristics of a high-speed jet with chevron nozzles are experimentally investigated at a Mach number of 0.8. The main focus is to examine the effects of the extent of chevron penetration and its position in the mixing layer. Chevron nozzles with three different levels of penetration employed at three different longitudinal locations from the nozzle lip are tested, and the results are compared with those of a plain baseline nozzle. The chevrons are found to produce a lobed shear layer through the notched region, thereby increasing the surface area of the jet, particularly in the close vicinity of the nozzle, which increases the mixing and reduces the potential core length. This effect becomes more prominent with increasing penetration closer to the nozzle lip in the thinner mixing layer. Near field and far field noise measurements show distinctly different acoustic features due to chevrons. The chevrons are found to effectively shift the dominant noise source upstream closer to the nozzle. Present investigation proposes a simpler method for locating the dominant noise source from the peak of the centerline velocity decay rate. The overall noise levels registered along the jet edge immediately downstream of the chevrons are higher, but further downstream they are reduced in comparison with the plain baseline nozzle. Also, the chevrons beam the noise towards higher polar angles at higher frequencies. At shallow polar angles with respect to the jet axis in the far field, chevrons suppress the noise at low frequencies with increasing penetration, but for higher polar angles, while they continue to suppress the low frequency noise, at higher frequencies the trend is found to reverse. The noise measured in the near field close to the jet edge is composed of two components: acoustic and hydrodynamic. Of these two components, the chevrons are found to reduce the hydrodynamic component in comparison with the acoustic one.

Theoretical Rocket Performance of Liquid Methane with Several Fluorine-Oxygen Mixtures Assuming Frozen Composition

NASA Technical Reports Server (NTRS)

Gordon, Sanford; Kastner, Michael E

1958-01-01

Theoretical rocket performance for frozen composition during expansion was calculated for liquid methane with several fluorine-oxygen mixtures for a range of pressure ratios and oxidant-fuel ratios. The parameters included are specific impulse, combustion-chamber temperature, nozzle-exit temperature molecular weight, characteristic velocity, coefficient of thrust, ratio of nozzle-exit area to throat area, specific heat at constant pressure, isentropic exponent, viscosity, and thermal conductivity. The maximum calculated value of specific impulse for a chamber pressure of 600 pounds per square inch absolute (40.827atm) and an exit pressure of 1 atmosphere is 315.3 for 79.67 percent fluorine in the oxidant.

Pressure distribution in a converging-diverging nozzle during two-phase choked flow of subcooled nitrogen

NASA Technical Reports Server (NTRS)

Simoneau, R. J.

1975-01-01

Choked flow rates and axial pressure distributions were measured for subcooled nitrogen in a converging-diverging nozzle with a constant area section in the throat region. Stagnation pressures ranged from slightly above saturation to twice the thermodynamic critical pressure. Stagnation temperatures ranged from 0.75 to 1.03 times the thermodynamic critical temperature. The choking plane is at the divergence end of the constant area throat section. At high stagnation pressures the fluid stays liquid well into the constant area throat region; at near saturation stagnation pressures it appears that vaporization occurs at or before the entrance to the constant area throat region. The throat-to-stagnation pressure ratio data exhibits an anomalous flat region, and this anomaly is related to the two-phase process. The fluid is metastably all liquid below the saturation pressure.

On the Scaling of Small, Heat Simulated Jet Noise Measurements to Moderate Size Exhaust Jets

NASA Technical Reports Server (NTRS)

McLaughlin, Dennis K.; Bridges, James; Kuo, Ching-Wen

2010-01-01

Modern military aircraft jet engines are designed with variable geometry nozzles to provide optimum thrust in different operating conditions, depending on the flight envelope. However, the acoustic measurements for such nozzles are scarce, due to the cost involved in making full scale measurements and the lack of details about the exact geometry of these nozzles. Thus the present effort at The Pennsylvania State University and the NASA Glenn Research Center- in partnership with GE Aviation is aiming to study and characterize the acoustic field produced by supersonic jets issuing from converging-diverging military style nozzles. An equally important objective is to validate methodology for using data obtained from small and moderate scale experiments to reliably predict the most important components of full scale engine noise. The experimental results presented show reasonable agreement between small scale and moderate scale jet acoustic data, as well as between heated jets and heat-simulated ones. Unresolved issues however are identified that are currently receiving our attention, in particular the effect of the small bypass ratio airflow. Future activities will identify and test promising noise reduction techniques in an effort to predict how well such concepts will work with full scale engines in flight conditions.

Electrosprayed Cerium Oxide Nanoparticles

NASA Astrophysics Data System (ADS)

Azar, Pedram Bagherzadeh; Tavanai, Hossein; Allafchian, Ali Reza

2018-04-01

Cerium oxide nanoparticles were fabricated via the calcination of electrosprayed polyvinyl alcohol (PVA)/cerium nitrate nanoparticles. The effect of material variables of PVA/cerium nitrate electrospraying solution, i.e. viscosity, surface tension and electrical conductivity, as well as important process variables like voltage, nozzle-collector distance and feed rate on cerium oxide nanoparticle size, are investigated. Scanning electron microscopy and Fourier-transform infrared (FTIR) spectroscopy analysis have also been carried out. The results showed that electrospraying of PVA/cerium nitrate (25% w/v) was only possible with PVA concentrations in the range of 5-8% w/v. With other conditions constant, decreasing PVA concentration, decreasing feed rate, increasing nozzle-collector distance and increasing voltage decreased the size of the final cerium oxide nanoparticles. The gross average size of all cerium oxide nanoparticles obtained in this work was about 80 nm. FTIR analysis proved the formation of cerium oxide after the calcination process.

Improved ablative materials for the ASRM nozzle

NASA Technical Reports Server (NTRS)

Canfield, A.; Clinton, R. G.; Armour, W.; Koenig, J.

1992-01-01

Rayon precursor carbon-cloth phenolic was developed more than 30 years ago and is used in most nozzles today including the Poseidon, Trident, Peacekeeper, Small ICBM, Space Shuttle, and numerous tactical and space systems. Specifications and manufacturing controls were placed on these materials and, once qualified, a no-change policy was instituted. The current material is acceptable; however, prepreg variability does not always accommodate the requirements of automation. The advanced solid rocket motor requires material with less variability for automated manufacturing. An advanced solid rocket motor materials team, composed of NASA, Thiokol, Aerojet, SRI, and Lockheed specialists, along with materials suppliers ICI Fiberite/Polycarbon, BP Chemicals/Hitco, and Amoco, embarked on a program to improve the current materials. The program consisted of heat treatment studies and standard and low-density material improvements evaluation. Improvements evaluated included fiber/fabric heat treatments, weave variations, resin application methods, process controls, and monitors.

Advanced Gas Turbine (AGT): Power-train system development

NASA Technical Reports Server (NTRS)

Helms, H. E.; Johnson, R. A.; Gibson, R. K.; Smith, L. B.

1983-01-01

Technical work on the design and effort leading to the testing of a 74.5 kW (100 hp) automotive gas turbine is described. The general effort was concentrated on building an engine for test starting in July. The buildup progressed with only routine problems and the engine was delivered to the test stand 9 July. In addition to the engine build effort, work continued in selected component areas. Ceramic turbine parts were built and tested. Burst tests of ceramic rotors show strengths are approaching that achieved in test bars; proof testing is required for acceptable strength ceramic vanes. Over 25 hours was accumulated on the combustor rig in three test modes: pilot nozzle only, start nozzle, and main nozzle operation. Satisfactory ignition was achieved for a wide range of starting speeds and the lean blowout limit was as low as 0.06 kg/b (0.14 lb/hr). Lean blowout was more a function of nozzle atomization than fuel/air ratio. A variety of cycle points were tested. Transition from start nozzle flow to main nozzle flow was done manually without difficulty. Regenerator parts were qualification tested without incident and the parts were assembled on schedule. Rig based performance matched first build requirements. Repeated failures in the harmonic drive gearbox during rig testing resulted in that concept being abandoned for an alternate scheme.

Microfabricated airflow nozzle for microencapsulation of living cells into 150 micrometer microcapsules.

PubMed

Sugiura, Shinji; Oda, Tatsuya; Aoyagi, Yasuyuki; Matsuo, Ryota; Enomoto, Tsuyoshi; Matsumoto, Kunio; Nakamura, Toshikazu; Satake, Mitsuo; Ochiai, Atsushi; Ohkohchi, Nobuhiro; Nakajima, Mitsutoshi

2007-02-01

Microencapsulation of genetically engineered cells has attracted much attention as an alternative nonviral strategy to gene therapy. Though smaller microcapsules (i.e. less than 300 microm) theoretically have various advantages, technical limitations made it difficult to prove this notion. We have developed a novel microfabricated device, namely a micro-airflow-nozzle (MAN), to produce 100 to 300 microm alginate microcapsules with a narrow size distribution. The MAN is composed of a nozzle with a 60 microm internal diameter for an alginate solution channel and airflow channels next to the nozzle. An alginate solution extruded through the nozzle was sheared by the airflow. The resulting alginate droplets fell directly into a CaCl2 solution, and calcium alginate beads were formed. The device enabled us to successfully encapsulate living cells into 150 microm microcapsules, as well as control microcapsule size by simply changing the airflow rate. The encapsulated cells had a higher growth rate and greater secretion activity of marker protein in 150 microm microcapsules compared to larger microcapsules prepared by conventional methods because of their high diffusion efficiency and effective scaffold surface area. The advantages of smaller microcapsules offer new prospects for the advancement of microencapsulation technology.

MHD Simulation of Magnetic Nozzle Plasma with the NIMROD Code: Applications to the VASIMR Advanced Space Propulsion Concept

NASA Astrophysics Data System (ADS)

Tarditi, Alfonso G.; Shebalin, John V.

2002-11-01

A simulation study with the NIMROD code [1] is being carried on to investigate the efficiency of the thrust generation process and the properties of the plasma detachment in a magnetic nozzle. In the simulation, hot plasma is injected in the magnetic nozzle, modeled as a 2D, axi-symmetric domain. NIMROD has two-fluid, 3D capabilities but the present runs are being conducted within the MHD, 2D approximation. As the plasma travels through the magnetic field, part of its thermal energy is converted into longitudinal kinetic energy, along the axis of the nozzle. The plasma eventually detaches from the magnetic field at a certain distance from the nozzle throat where the kinetic energy becomes larger than the magnetic energy. Preliminary NIMROD 2D runs have been benchmarked with a particle trajectory code showing satisfactory results [2]. Further testing is here reported with the emphasis on the analysis of the diffusion rate across the field lines and of the overall nozzle efficiency. These simulation runs are specifically designed for obtaining comparisons with laboratory measurements of the VASIMR experiment, by looking at the evolution of the radial plasma density and temperature profiles in the nozzle. VASIMR (Variable Specific Impulse Magnetoplasma Rocket, [3]) is an advanced space propulsion concept currently under experimental development at the Advanced Space Propulsion Laboratory, NASA Johnson Space Center. A plasma (typically ionized Hydrogen or Helium) is generated by a RF (Helicon) discharge and heated by an Ion Cyclotron Resonance Heating antenna. The heated plasma is then guided into a magnetic nozzle to convert the thermal plasma energy into effective thrust. The VASIMR system has no electrodes and a solenoidal magnetic field produced by an asymmetric mirror configuration ensures magnetic insulation of the plasma from the material surfaces. By powering the plasma source and the heating antenna at different levels it is possible to vary smoothly of the thrust-to-specific impulse ratio while maintaining maximum power utilization. [1] http://www.nimrodteam.org [2] A. V. Ilin et al., Proc. 40th AIAA Aerospace Sciences Meeting, Reno, NV, Jan. 2002 [3] F. R. Chang-Diaz, Scientific American, p. 90, Nov. 2000

Shock capturing finite-difference and characteristic reference plane techniques for the prediction of three-dimensional nozzle-exhaust flowfields

NASA Technical Reports Server (NTRS)

Dash, S.; Delguidice, P.

1978-01-01

This report summarizes work accomplished under Contract No. NAS1-12726 towards the development of computational procedures and associated numerical. The flow fields considered were those associated with airbreathing hypersonic aircraft which require a high degree of engine/airframe integration in order to achieve optimized performance. The exhaust flow, due to physical area limitations, was generally underexpanded at the nozzle exit; the vehicle afterbody undersurface was used to provide additional expansion to obtain maximum propulsive efficiency. This resulted in a three dimensional nozzle flow, initialized at the combustor exit, whose boundaries are internally defined by the undersurface, cowling and walls separating individual modules, and externally, by the undersurface and slipstream separating the exhaust flow and external stream.

Results of a space shuttle pulme impingement investigation at stage separation in the NASA-MSFC impulse base flow facility

NASA Technical Reports Server (NTRS)

Mccanna, R. W.; Sims, W. H.

1972-01-01

Results are presented for an experimental space shuttle stage separation plume impingement program conducted in the NASA-Marshall Space Flight Center's impulse base flow facility (IBFF). Major objectives of the investigation were to: (1)determine the degree of dual engine exhaust plume simulation obtained using the equivalent engine; (2) determine the applicability of the analytical techniques; and (3) obtain data applicable for use in full-scale studies. The IBFF tests determined the orbiter rocket motor plume impingement loads, both pressure and heating, on a 3 percent General Dynamics B-15B booster configuration in a quiescent environment simulating a nominal staging altitude of 73.2 km (240,00 ft). The data included plume surveys of two 3 percent scale orbiter nozzles, and a 4.242 percent scaled equivalent nozzle - equivalent in the sense that it was designed to have the same nozzle-throat-to-area ratio as the two 3 percent nozzles and, within the tolerances assigned for machining the hardware, this was accomplished.

AST Critical Propulsion and Noise Reduction Technologies for Future Commercial Subsonic Engines Area of Interest 1.0: Reliable and Affordable Control Systems

NASA Technical Reports Server (NTRS)

Myers, William; Winter, Steve

2006-01-01

The General Electric Reliable and Affordable Controls effort under the NASA Advanced Subsonic Technology (AST) Program has designed, fabricated, and tested advanced controls hardware and software to reduce emissions and improve engine safety and reliability. The original effort consisted of four elements: 1) a Hydraulic Multiplexer; 2) Active Combustor Control; 3) a Variable Displacement Vane Pump (VDVP); and 4) Intelligent Engine Control. The VDVP and Intelligent Engine Control elements were cancelled due to funding constraints and are reported here only to the state they progressed. The Hydraulic Multiplexing element developed and tested a prototype which improves reliability by combining the functionality of up to 16 solenoids and servo-valves into one component with a single electrically powered force motor. The Active Combustor Control element developed intelligent staging and control strategies for low emission combustors. This included development and tests of a Controlled Pressure Fuel Nozzle for fuel sequencing, a Fuel Multiplexer for individual fuel cup metering, and model-based control logic. Both the Hydraulic Multiplexer and Controlled Pressure Fuel Nozzle system were cleared for engine test. The Fuel Multiplexer was cleared for combustor rig test which must be followed by an engine test to achieve full maturation.

Thrust Augmenting Ejectors

DTIC Science & Technology

1975-06-01

approximately four . Presently referred to as the hypermixing nozzle (See Figure 37). d. Alternating exit nozzle - same as c, but the elemental aspect ratio...UL 0 Reynolds number, defined as R . In addition, due to the double symmetry about the y and z axes, the analysis will treat only one quadrant of the...in the duct of diffuser. Then from the conservation of mass A~x Atx) where A(x) g(x) tl(x) is the streamwise area distribution in one quadrant

Efficiency arcjet thruster with controlled arc startup and steady state attachment

NASA Technical Reports Server (NTRS)

Smith, William W. (Inventor); Knowles, Steven C. (Inventor)

1989-01-01

An improved efficiency arcjet thruster has a constrictor and electrically-conductive nozzle anode defining an arc chamber, and an electrically-conductive rod having a tip spaced upstream from the constrictor and defining a cathode spaced from the anode by a gap generally coextensive with the arc chamber. An electrical potential is applied to the anode and cathode to generate an electrical arc in the arc chamber from the cathode to anode. Catalytically decomposed hydrazine is supplied to the arc chamber with generation of the arc so as to produce thermal heating and expansion thereof through the nozzle. The constrictor can have a electrically insulative portion disposed between the cathode tip and the nozzle anode, and an electrically-conductive anode extension disposed along the insulative portion so as to define an auxiliary gap with the cathode tip substantially smaller than the gap defined between the cathode and nozzle anode for facilitating startup of arc generation. The constrictor can also include an electrically-conductive electrode with a variable electrical potential to vary the shape of the arc generated in the arc chamber. Also, the cathode is mounted for axial movement such that the gap between its tip and the nozzle anode can be varied to facilitate a generally nonerosive generation of the electrical arc at startup and reliable steady state operation. Further, the arc chamber can have a nonparallel subsonic-to-supersonic transition configuration, or alternatively solely a nonparallel supersonic configuration, for improved arc attachment.

High Pressure Water Stripping Using Multi-Orifice Nozzles

NASA Technical Reports Server (NTRS)

Hoppe, David

1999-01-01

The use of multi-orifice rotary nozzles greatly increases the speed and stripping effectiveness of high pressure water blasting systems, but also greatly increases the complexity of selecting and optimizing the operating parameters. The rotational speed of the nozzle must be coupled with its transverse velocity as it passes across the surface of the substrate being stripped. The radial and angular positions of each orifice must be included in the analysis of the nozzle configuration. Orifices at the outer edge of the nozzle head move at a faster rate than the orifices located near the center. The energy transmitted to the surface from the impact force of the water stream from an outer orifice is therefore spread over a larger area than energy from an inner orifice. Utilizing a larger diameter orifice in the outer radial positions increases the total energy transmitted from the outer orifice to compensate for the wider distribution of energy. The total flow rate from the combination of all orifices must be monitored and should be kept below the pump capacity while choosing orifice to insert in each position. The energy distribution from the orifice pattern is further complicated since the rotary path of all the orifices in the nozzle head pass through the center section. All orifices contribute to the stripping in the center of the path while only the outer most orifice contributes to the stripping at the edge of the nozzle. Additional orifices contribute to the stripping from the outer edge toward the center section. With all these parameters to configure and each parameter change affecting the others, a computer model was developed to track and coordinate these parameters. The computer simulation graphically indicates the cumulative affect from each parameter selected. The result from the proper choices in parameters is a well designed, highly efficient stripping system. A poorly chosen set of parameters will cause the nozzle to strip aggressively in some areas while leaving the coating untouched in adjacent sections. The high pressure water stripping system can be set to extremely aggressive conditions allowing stripping of hard to remove adhesives, paint systems, and even cladding and chromate conversion coatings. The energy force can also be reduced to strip coatings from thin aluminum substrates without causing any damage or deterioration to the substrate's surface. High pressure water stripping of aerospace components has thus proven to be an efficient and cost effective method for cleaning and removing coatings.

Study of methods of improving the performance of the Langley Research Center Transonic Dynamics Tunnel (TDT)

NASA Technical Reports Server (NTRS)

1973-01-01

A study has been made of possible ways to improve the performance of the Langley Research Center's Transonic Dynamics Tunnel (TDT). The major effort was directed toward obtaining increased dynamic pressure in the Mach number range from 0.8 to 1.2, but methods to increase Mach number capability were also considered. Methods studied for increasing dynamic pressure capability were higher total pressure, auxiliary suction, reducing circuit losses, reduced test medium temperature, smaller test section and higher molecular weight test medium. Increased Mach number methods investigated were nozzle block inserts, variable geometry nozzle, changes in test section wall configuration, and auxiliary suction.

Theoretical performance of liquid hydrogen and liquid fluorine as a rocket propellant

NASA Technical Reports Server (NTRS)

Gordon, Sanford; Huff, Vearl N

1953-01-01

Theoretical values of performance parameters for liquid hydrogen and liquid fluorine as a rocket propellant were calculated on the assumption of equilibrium composition during the expansion process for a wide range of fuel-oxidant and expansion ratios. The parameters included were specific impulse, combustion-chamber temperature, nozzle-exit temperature, equilibrium composition, mean molecular weight, characteristic velocity, coefficient of thrust, ration of nozzle-exit area to throat area, specific heat at constant pressure, coefficient of viscosity, and coefficient of thermal conductivity. The maximum value of specific impulse was 364.6 pound-seconds per pound for a chamber pressure of 300 pounds per square inch absolute (20.41 atm) and an exit pressure of 1 atmosphere.

Theoretical performance of liquid ammonia and liquid fluorine as a rocket propellant

NASA Technical Reports Server (NTRS)

Gordon, Sanford; Huff, Vearl N

1953-01-01

Theoretical values of performance parameters for liquid ammonia and liquid fluorine as a rocket propellant were calculated on the assumption of equilibrium composition during the expansion process for a wide range of fuel-oxidant and expansion ratios. The parameters included were specific impulse, combustion chamber temperature, nozzle-exit temperature, equilibrium composition, mean molecular weight, characteristic velocity, coefficient of thrust, ratio of nozzle-exit area to throat area, specific heat at constant pressure, coefficient of viscosity, and coefficient of thermal conductivity. The maximum value of specific impulse was 311.5 pound-seconds per pound for a chamber pressure of 300 pounds per square inch absolute (20.41 atm) and an exit pressure of 1 atmosphere.

RANS Analyses of Turbofan Nozzles with Internal Wedge Deflectors for Noise Reduction

NASA Technical Reports Server (NTRS)

DeBonis, James R.

2009-01-01

Computational fluid dynamics (CFD) was used to evaluate the flow field and thrust performance of a promising concept for reducing the noise at take-off of dual-stream turbofan nozzles. The concept, offset stream technology, reduces the jet noise observed on the ground by diverting (offsetting) a portion of the fan flow below the core flow, thickening and lengthening this layer between the high-velocity core flow and the ground observers. In this study a wedge placed in the internal fan stream is used as the diverter. Wind, a Reynolds averaged Navier-Stokes (RANS) code, was used to analyze the flow field of the exhaust plume and to calculate nozzle performance. Results showed that the wedge diverts all of the fan flow to the lower side of the nozzle, and the turbulent kinetic energy on the observer side of the nozzle is reduced. This reduction in turbulent kinetic energy should correspond to a reduction in noise. However, because all of the fan flow is diverted, the upper portion of the core flow is exposed to the freestream, and the turbulent kinetic energy on the upper side of the nozzle is increased, creating an unintended noise source. The blockage due to the wedge reduces the fan mass flow proportional to its blockage, and the overall thrust is consequently reduced. The CFD predictions are in very good agreement with experimental flow field data, demonstrating that RANS CFD can accurately predict the velocity and turbulent kinetic energy fields. While this initial design of a large scale wedge nozzle did not meet noise reduction or thrust goals, this study identified areas for improvement and demonstrated that RANS CFD can be used to improve the concept.

Static investigation of a two-dimensional convergent-divergent exhaust nozzle with multiaxis thrust-vectoring capability

NASA Technical Reports Server (NTRS)

Taylor, John G.

1990-01-01

An investigation was conducted in the Static Test Facility of the NASA Langley 16-Foot Transonic Tunnel to determine the internal performance of two-dimensional convergent-divergent nozzles designed to have simultaneous pitch and yaw thrust vectoring capability. This concept utilized divergent flap rotation of thrust vectoring in the pitch plane and deflection of flat yaw flaps hinged at the end of the sidewalls for yaw thrust vectoring. The hinge location of the yaw flaps was varied at four positions from the nozzle exit plane to the throat plane. The yaw flaps were designed to contain the flow laterally independent of power setting. In order to eliminate any physical interference between the yaw flap deflected into the exhaust stream and the divergent flaps, the downstream corners of both upper and lower divergent flaps were cut off to allow for up to 30 deg of yaw flap deflection. The impact of varying the nozzle pitch vector angle, throat area, yaw flap hinge location, yaw flap length, and yaw flap deflection angle on nozzle internal performance characteristics, was studied. High-pressure air was used to simulate jet exhaust at nozzle pressure ratios up to 7.0. Static results indicate that configurations with the yaw flap hinge located upstream of the exit plane provide relatively high levels of thrust vectoring efficiency without causing large losses in resultant thrust ratio. Therefore, these configurations represent a viable concept for providing simultaneous pitch and yaw thrust vectoring.

Apollo Contour Rocket Nozzle in the Propulsion Systems Laboratory

NASA Image and Video Library

1964-07-21

Bill Harrison and Bud Meilander check the setup of an Apollo Contour rocket nozzle in the Propulsion Systems Laboratory at the National Aeronautics and Space Administration (NASA) Lewis Research Center. The Propulsion Systems Laboratory contained two 14-foot diameter test chambers that could simulate conditions found at very high altitudes. The facility was used in the 1960s to study complex rocket engines such as the Pratt and Whitney RL-10 and rocket components such as the Apollo Contour nozzle, seen here. Meilander oversaw the facility’s mechanics and the installation of test articles into the chambers. Harrison was head of the Supersonic Tunnels Branch in the Test Installations Division. Researchers sought to determine the impulse value of the storable propellant mix, classify and improve the internal engine performance, and compare the results with analytical tools. A special setup was installed in the chamber that included a device to measure the thrust load and a calibration stand. Both cylindrical and conical combustion chambers were examined with the conical large area ratio nozzles. In addition, two contour nozzles were tested, one based on the Apollo Service Propulsion System and the other on the Air Force’s Titan transtage engine. Three types of injectors were investigated, including a Lewis-designed model that produced 98-percent efficiency. It was determined that combustion instability did not affect the nozzle performance. Although much valuable information was obtained during the tests, attempts to improve the engine performance were not successful.

Wind-US Results for the AIAA 2nd Propulsion Aerodynamics Workshop

NASA Technical Reports Server (NTRS)

Dippold, Vance III; Foster, Lancert; Mankbadi, Mina

2014-01-01

This presentation contains Wind-US results presented at the 2nd Propulsion Aerodynamics Workshop. The workshop was organized by the American Institute of Aeronautics and Astronautics, Air Breathing Propulsion Systems Integration Technical Committee with the purpose of assessing the accuracy of computational fluid dynamics for air breathing propulsion applications. Attendees included representatives from government, industry, academia, and commercial software companies. Participants were encouraged to explore and discuss all aspects of the simulation process including the effects of mesh type and refinement, solver numerical schemes, and turbulence modeling. The first set of challenge cases involved computing the thrust and discharge coefficients for a 25deg conical nozzle for a range of nozzle pressure ratios between 1.4 and 7.0. Participants were also asked to simulate two cases in which the 25deg conical nozzle was bifurcated by a solid plate, resulting in vortex shedding (NPR=1.6) and shifted plume shock (NPR=4.0). A second set of nozzle cases involved computing the discharge and thrust coefficients for a convergent dual stream nozzle for a range of subsonic nozzle pressure ratios. The workshop committee also compared the plume mixing of these cases across various codes and models. The final test case was a serpentine inlet diffuser with an outlet to inlet area ratio of 1.52 and an offset of 1.34 times the inlet diameter. Boundary layer profiles, wall static pressure, and total pressure at downstream rake locations were examined.

Flow-structure interaction effects on a jet emanating from a flexible nozzle

PubMed Central

Murugappan, S.; Gutmark, E. J.; Lakhamraju, R. R.; Khosla, S.

2008-01-01

In recent years, a wide variety of applications have been found for the use of pulsed jets in the area of flow control. The goal of the current study was to identify the flow field and mixing characteristics associated with an incompressible elongated jet emitted from a flexible nozzle. The shape of the nozzle was that of a high aspect ratio jet deforming from a fully opened to a completely closed configuration. The jet was characterized by a pulsatile flow that was self-excited by the motion of the flexible tube. The frequency of excitation was found to be between 150 and 175 Hz and the Strouhal number (nondimensional frequency) varied from 0.17 to 0.45. The jet flow was dominated by vortices that were shed from the nozzle with an axis parallel to the major axis. The vortices in the near field were quasi-two-dimensional so that measurements performed at the center plane represented the dynamics of the entire vortex. The nozzle excited two different modes depending on the tension applied to the flexible nozzle and the volumetric flow through it. The first was a flapping mode, which was associated with alternate shedding of vortices. This caused strong steering of the jet to one side or the other. The second mode was a symmetric mode that was associated with the formation of counter-rotating vortex pairs. Turbulence and jet spread in the measured planes were much larger in the first mode than the second one. PMID:19547723

Nonideal isentropic gas flow through converging-diverging nozzles

NASA Technical Reports Server (NTRS)

Bober, W.; Chow, W. L.

1990-01-01

A method for treating nonideal gas flows through converging-diverging nozzles is described. The method incorporates the Redlich-Kwong equation of state. The Runge-Kutta method is used to obtain a solution. Numerical results were obtained for methane gas. Typical plots of pressure, temperature, and area ratios as functions of Mach number are given. From the plots, it can be seen that there exists a range of reservoir conditions that require the gas to be treated as nonideal if an accurate solution is to be obtained.

Modified RS2101 rocket engine study program

NASA Technical Reports Server (NTRS)

1971-01-01

The purpose of the program is to perform design studies and analyses to determine the effects of incorporating a 60:1 expansion area ratio nozzle extension, extended firing time, and modified operating conditions and environments on the MM'71 rocket engine assembly. An injector-to-thrust chamber seal study was conducted to define potential solutions for leakage past this joint. The results and recommendations evolving from the engine thermal analyses, the injector-to-thrust chamber seal studies, and the nozzle extension joint stress analyses are presented.

Engine Cycle Analysis for a Particle Bed Reactor Nuclear Rocket

DTIC Science & Technology

1991-03-01

0 DTIC USERS UNCLASSIFIED 22a. NAME OF RESPONSIBLE INDIVIDUAL ZZb. TELEPHONE (Include Area Code) 22c. OFFICE SYMBOL Lt Timothy J . Lawrence 805-275...Cycle with 2000 MW PBR and Uncooled Nozzle J : Output for Bleed Cycle with 2000 MW PBR and Cooled Nozzle K: Output for Expander Cycle with 2000 MW PBR L...Mars with carbon dioxide, the primary component of the Martian atmosphere. Carbon dioxide would delivera smaller ! j , but its use would eliminate the

Carbon-Carbon Nozzle Extension Development in Support of In-Space and Upper-Stage Liquid Rocket Engines

NASA Technical Reports Server (NTRS)

Gradl, Paul R.; Valentine, Peter G.

2017-01-01

Upper stage and in-space liquid rocket engines are optimized for performance through the use of high area ratio nozzles to fully expand combustion gases to low exit pressures, increasing exhaust velocities. Due to the large size of such nozzles, and the related engine performance requirements, carbon-carbon (C-C) composite nozzle extensions are being considered to reduce weight impacts. Currently, the state-of-the-art is represented by the metallic and foreign composite nozzle extensions limited to approximately 2000 degrees F. used on the Atlas V, Delta IV, Falcon 9, and Ariane 5 launch vehicles. NASA and industry partners are working towards advancing the domestic supply chain for C-C composite nozzle extensions. These development efforts are primarily being conducted through the NASA Small Business Innovation Research (SBIR) program in addition to other low level internal research efforts. This has allowed for the initial material development and characterization, subscale hardware fabrication, and completion of hot-fire testing in relevant environments. NASA and industry partners have designed, fabricated and hot-fire tested several subscale domestically produced C-C extensions to advance the material and coatings fabrication technology for use with a variety of liquid rocket and scramjet engines. Testing at NASA's Marshall Space Flight Center (MSFC) evaluated heritage and state-of-the-art C-C materials and coatings, demonstrating the initial capabilities of the high temperature materials and their fabrication methods. This paper discusses the initial material development, design and fabrication of the subscale carbon-carbon nozzle extensions, provides an overview of the test campaign, presents results of the hot fire testing, and discusses potential follow-on development work. The follow on work includes the fabrication of ultra-high temperature materials, larger C-C nozzle extensions, material characterization, sub-element testing and hot-fire testing at larger scale.

Analytical study of striated nozzle flow with small radius of curvature ratio throats

NASA Technical Reports Server (NTRS)

Norton, D. J.; White, R. E.

1972-01-01

An analytical method was developed which is capable of estimating the chamber and throat conditions in a nozzle with a low radius of curvature throat. The method was programmed using standard FORTRAN 4 language and includes chemical equilibrium calculation subprograms (modified NASA Lewis program CEC71) as an integral part. The method determines detailed and gross rocket characteristics in the presence of striated flows and gives detailed results for the motor chamber and throat plane with as many as 20 discrete zones. The method employs a simultaneous solution of the mass, momentum, and energy equations and allows propellant types, 0/F ratios, propellant distribution, nozzle geometry, and injection schemes to be varied so to predict spatial velocity, density, pressure, and other thermodynamic variable distributions in the chamber as well as the throat. Results for small radius of curvature have shown good comparison to experimental results. Both gaseous and liquid injection may be considered with frozen or equilibrium flow calculations.

Experimental and Theoretical Studies of Axisymmetric Free Jets

NASA Technical Reports Server (NTRS)

Love, Eugene S.; Grigsby, Carl E.; Lee, Louise P.; Woodling, Mildred J.

1959-01-01

Some experimental and theoretical studies have been made of axisymmetric free jets exhausting from sonic and supersonic nozzles into still air and into supersonic streams with a view toward problems associated with propulsive jets and the investigation of these problems. For jets exhausting into still air, consideration is given to the effects of jet Mach number, nozzle divergence angle, and jet static pressure ratio upon jet structure, jet wavelength, and the shape and curvature of the jet boundary. Studies of the effects of the ratio of specific heats of the jets are included are observations pertaining to jet noise and jet simulation. For jets exhausting into supersonic streams, an attempt has been made to present primarily theoretical certain jet interference effects and in formulating experimental studies. The primary variables considered are jet Mach number, free stream Mach number, jet static pressure ratio, ratio of specific heats of the jet, nozzle exit angle, and boattail angle. The simulation problem and the case of a hypothetical hypersonic vehicle are examined, A few experimental observations are included.

Summary of Fluidic Thrust Vectoring Research Conducted at NASA Langley Research Center

NASA Technical Reports Server (NTRS)

Deere, Karen A.

2003-01-01

Interest in low-observable aircraft and in lowering an aircraft's exhaust system weight sparked decades of research for fixed geometry exhaust nozzles. The desire for such integrated exhaust nozzles was the catalyst for new fluidic control techniques; including throat area control, expansion control, and thrust-vector angle control. This paper summarizes a variety of fluidic thrust vectoring concepts that have been tested both experimentally and computationally at NASA Langley Research Center. The nozzle concepts are divided into three categories according to the method used for fluidic thrust vectoring: the shock vector control method, the throat shifting method, and the counterflow method. This paper explains the thrust vectoring mechanism for each fluidic method, provides examples of configurations tested for each method, and discusses the advantages and disadvantages of each method.

Mach 5 electroformed nickel nozzle refurbishment FNAS investigation of ultra-smooth surfaces

NASA Technical Reports Server (NTRS)

Rood, Robert; Griffith, Charles; Engelhaupt, Darell

1992-01-01

This task is in support of the Quiet Hypersonic Wind Tunnel effort currently in effect at NASA Langley Research Center, VA. A laminar flow wind tunnel nozzle has been previously fabricated by electroforming pure nickel over a machined and polished two piece stainless steel mandrel. The mandrel was removed leaving the replicate nozzle surface. The nozzle was then pressed into a heavy stainless steel jacket for mounting features and rigidity. The original nickel surface was a replication of the polished mandrel but had degraded due to oxidation. The inside surface requirements are very stringent in order to achieve laminar or quiet flow at the desired pressure and temperature for the specific design of Mach 5. The throat area of the axisymmetric device must have a surface finish with no defects greater than 16 microinches. This requires an rms average background of about four microinches or better for inspection purposes. A coating of nickel-phosphorous alloy was applied by catalytic deposition and then the inside of the nozzle was polished retaining dimensional and surface finish tolerances as specified per drawings supplied. Since the unit is not an optical component, conventional optical inspection methods for surface finish and figure are not readily achieved. Measurements were made using surface profilometry.

Dual-throat thruster thermal model

NASA Technical Reports Server (NTRS)

Ewen, R. L.; Obrien, C. J.; Matthews, L. W.

1986-01-01

The dual-throat engine is one of the dual nozzle engine concepts studied for advanced space transportation applications. It provides a thrust change and an in-flight area ratio change through the use of two concentric combustors with their throats arranged in series. Test results are presented for a dual throat thruster burning gaseous oxygen and hydrogen at primary (inner) chamber pressures from 380 to 680 psia. Heat flux profiles were obtained from calorimetric cooling channels in the inner nozzle, outer or secondary chamber and the tip of the inner nozzle. Data were obtained for two nozzle spacings over a chamber pressure ratio (secondary/primary) range of 0.45 to 0.83 with both chambers firing (Mode I). Fluxes near the end of the inner nozzle were significantly higher than in Mode II when only the inner chamber was fired, due to the flow separation and recirculation caused by the back pressure imposed by the secondary chamber. As the pressure ratio increased, these heat fluxes increased and the region of high heat flux relative to Mode II extended farther upstream. The use of the gaseous hydrogen bleed flow in the secondary chamber to control heat fluxes in the primary plume attachment region was investigated in Mode II testing. A thermal model of a dual throat thruster was developed and upgraded using the experimental data.

Experimental Altitude Performance of JP-4 Fuel and Liquid-Oxygen Rocket Engine with an Area Ratio of 48

NASA Technical Reports Server (NTRS)

Fortini, Anthony; Hendrix, Charles D.; Huff, Vearl N.

1959-01-01

The performance for four altitudes (sea-level, 51,000, 65,000, and 70,000 ft) of a rocket engine having a nozzle area ratio of 48.39 and using JP-4 fuel and liquid oxygen as a propellant was evaluated experimentally by use of a 1000-pound-thrust engine operating at a chamber pressure of 600 pounds per square inch absolute. The altitude environment was obtained by a rocket-ejector system which utilized the rocket exhaust gases as the pumping fluid of the ejector. Also, an engine having a nozzle area ratio of 5.49 designed for sea level was tested at sea-level conditions. The following table lists values from faired experimental curves at an oxidant-fuel ratio of 2.3 for various approximate altitudes.

Variable Specific Impulse Magnetoplasma Rocket Engine

NASA Technical Reports Server (NTRS)

Chang-Diaz, Franklin R. (Inventor)

2002-01-01

An engine is disclosed, including a controllable output plasma generator, a controllable heater for selectably raising a temperature of the plasma connected to an outlet of the plasma generator, and a nozzle connected to an outlet of the heater, through which heated plasma is discharged to provide thrust. In one embodiment, the source of plasma is a helicon generator. In one embodiment, the heater is an ion cyclotron resonator. In one embodiment, the nozzle is a radially diverging magnetic field disposed on a discharge side of the heater so that helically travelling particles in the beater exit the heater at high axial velocity. A particular embodiment includes control circuits for selectably directing a portion of radio frequency power from an RF generator to the helicon generator and to the cyclotron resonator so that the thrust output and the specific impulse of the engine can be selectively controlled. A method of propelling a vehicle is also disclosed. The method includes generating a plasma, heating said plasma, and discharging the heated plasma through a nozzle. In one embodiment, the nozzle is a diverging magnetic field. In this embodiment, the heating is performed by applying a radio frequency electro magnetic field to the plasma at the ion cyclotron frequency in an axially polarized DC magnetic field.

Influence of Reduced Mass Flow Rate and Chamber Backpressure on Swirl Injector Fluid Mechanics

NASA Technical Reports Server (NTRS)

Kenny, R Jeremy; Hulka, James R.

2008-01-01

Industry interest in variable-thrust liquid rocket engines places a demand on engine injector technology to operate over a wide range of liquid mass flow rates and chamber backpressures. One injection technology of current interest for variable thrust applications is an injector design with swirled fluids. Current swirl injector design methodologies do not take into account how swirl injector design parameters respond to elevated chamber backpressures at less than design mass flow rates. The current work was created to improve state-of-the-art swirl injector design methods in this area. The specific objective was to study the effects of elevated chamber backpressure and off-design mass flow rates on swirl injector fluid mechanics. Using a backpressure chamber with optical access, water was flowed through a swirl injector at various combinations of chamber backpressure and mass flow rates. The film thickness profile down the swirl injector nozzle section was measured through a transparent nozzle section of the injector. High speed video showed measurable increases in the film thickness profile with application of chamber backpressure and mass flow rates less than design. At prescribed combinations of chamber backpressure and injected mass flow rate, a discrete change in the film thickness profile was observed. Measured injector discharge coefficient values showed different trends with increasing chamber backpressure at low mass flow rates as opposed to near-design mass flow rates. Downstream spray angles showed classic changes in morphology as the mass flow rate was decreased below the design value. Increasing chamber backpressure decreased the spray angle at any injection mass flow rate. Experimental measurements and discussion of these results are reported in this paper.

End-effects-regime in full scale and lab scale rocket nozzles

NASA Astrophysics Data System (ADS)

Rojo, Raymundo; Tinney, Charles; Baars, Woutijn; Ruf, Joseph

2014-11-01

Modern rockets utilize a thrust-optimized parabolic-contour design for their nozzles for its high performance and reliability. However, the evolving internal flow structures within these high area ratio rocket nozzles during start up generate a powerful amount of vibro-acoustic loads that act on the launch vehicle. Modern rockets must be designed to accommodate for these heavy loads or else risk a catastrophic failure. This study quantifies a particular moment referred to as the ``end-effects regime,'' or the largest source of vibro-acoustic loading during start-up [Nave & Coffey, AIAA Paper 1973-1284]. Measurements from full scale ignitions are compared with aerodynamically scaled representations in a fully anechoic chamber. Laboratory scale data is then matched with both static and dynamic wall pressure measurements to capture the associating shock structures within the nozzle. The event generated during the ``end-effects regime'' was successfully reproduced in the both the lab-scale models, and was characterized in terms of its mean, variance and skewness, as well as the spectral properties of the signal obtained by way of time-frequency analyses.

Analysis and classification of topographic flow steering and inferred geomorphic processes as a function of discharge in a mountain river

NASA Astrophysics Data System (ADS)

Gore, J.; Pasternack, G. B.; Wiener, J.

2016-12-01

Process-based river classification tends to be done at reach to catchment scales assuming channels are uniform and thus differentiated by the simple specific stream power metric. In fact, mountain rivers are highly variable at subreach scales to the point that local topographic steering may be the dominant control on geomorphic processes. This study presents a new framework for characterizing how stage-dependent topographic steering varies continuously down a river, leading to a classification of subreach landforms on the basis of the geomorphic mechanism of flow convergence routing. The two remote mountain river segments were located in the 3480-km2 Yuba River, with the upper South Yuba having a substantial sediment supply from legacy hydraulic gold mining and the mainstem Yuba downstream of New Bullards Bar Dam having a restricted sediment supply. Meter-scale DEMs were produced for both cases using airborne LiDAR and survey data. DEMs were slope detrended to focus the analysis on cross-sectional variability. DEMs were then heavily smoothed to allow for automated tracing of the valley centerline, and then cross-sectional rectangles were spaced every 5 m. The average width (W) and detrended bed elevation (Z) of the wetted area was computed from the DEM for each raster for 6-7 different river stages. Both width and cross-sectionally averaged bed elevation were standardized. The product of these two variables was computed as a measure of cross-sectional area, and is termed the geomorphic covariance (Czw) series when plotted along each river corridor. Cwz was then used to classify each cross-section as one of five distinct landform types: nozzle, wide bar, normal channel, constricted pool, and oversized pool- with this classification varying with discharge such that a section could, for example, function as a nozzle during low flow but an oversized pool at high flow, or any other combination. Longitudinal profiles of bed elevation, width, covariance, and landform type were analyzed for their stage-dependent patterns to understand their geomorphic significance and to contrast the two rivers. This new method may be the first example of a hierarchical, process-based classification at the subreach scale in which one mechanism is assessed for how it varies not only in space, but as a function of discharge.

Mixing of Supersonic Streams

NASA Technical Reports Server (NTRS)

Hawk, C. W.; Landrum, D. B.; Muller, S.; Turner, M.; Parkinson, D.

1998-01-01

The Strutjet approach to Rocket Based Combined Cycle (RBCC) propulsion depends upon fuel-rich flows from the rocket nozzles and turbine exhaust products mixing with the ingested air for successful operation in the ramjet and scramjet modes. It is desirable to delay this mixing process in the air-augmented mode of operation present during low speed flight. A model of the Strutjet device has been built and is undergoing test to investigate the mixing of the streams as a function of distance from the Strutjet exit plane during simulated low speed flight conditions. Cold flow testing of a 1/6 scale Strutjet model is underway and nearing completion. Planar Laser Induced Fluorescence (PLIF) diagnostic methods are being employed to observe the mixing of the turbine exhaust gas with the gases from both the primary rockets and the ingested air simulating low speed, air augmented operation of the RBCC. The ratio of the pressure in the turbine exhaust duct to that in the rocket nozzle wall at the point of their intersection is the independent variable in these experiments. Tests were accomplished at values of 1.0, 1.5 and 2.0 for this parameter. Qualitative results illustrate the development of the mixing zone from the exit plane of the model to a distance of about 10 rocket nozzle exit diameters downstream. These data show the mixing to be confined in the vertical plane for all cases, The lateral expansion is more pronounced at a pressure ratio of 1.0 and suggests that mixing with the ingested flow would be likely beginning at a distance of 7 nozzle exit diameters downstream of the nozzle exit plane.

Theoretical Performance of Liquid Hydrogen with Liquid Oxygen as a Rocket Propellant

NASA Technical Reports Server (NTRS)

Gordon, Sanford; McBride, Bonnie J.

1959-01-01

Theoretical rocket performance for both equilibrium and frozen composition during expansion was calculated for the propellant combination liquid hydrogen and liquid oxygen at four chamber pressures (60, 150, 300, and 600 lb/sq in. abs) and a wide range of pressure ratios (1 to 4000) and oxidant-fuel ratios (1.190 to 39.683). Data are given to estimate performance parameters at chamber pressures other than those for which data are tabulated. The parameters included are specific impulse, specific impulse in vacuum, combustion-chamber temperature, nozzle-exit temperature, molecular weight, molecular-weight derivatives, characteristic velocity, coefficient of thrust, ratio of nozzle-exit area to throat area, specific heat at constant pressure, isentropic exponent, viscosity, thermal conductivity, Mach number, and equilibrium gas compositions.

Effect of empennage location on twin-engine afterbody-nozzle aerodynamic characteristics at Mach Numbers from 0.6 to 1.2. [wind tunnel tests

NASA Technical Reports Server (NTRS)

Leavitt, L. D.

1983-01-01

The Langley 16-foot transonic tunnel was used to determine the effects of several empennage and afterbody parameters on the aft-end aerodynamic characteristics of a twin-engine fighter-type configuration. Model variables were as follows: horizontal tail axial location and incidence, vertical tail axial location and configuration (twin- versus single-tail arrangements), tail booms, and nozzle power setting. Tests were conducted over a Mach number range from 0.6 to 1.2 and over an angle-of-attack from -2 deg to 10 deg. Jet total-pressure ratio was varied from jet off to approximately 10.0.

Methane Dual Expander Aerospike Nozzle Rocket Engine

DTIC Science & Technology

2012-03-22

include O/F ratio, thrust, and engine geometry. After thousands of iterations over the design space , the selected MDEAN engine concept has 349 s of...35 Table 7: Fluid Property Table Supported Parameters...44 Table 8: Fluid Property Input Data Independent Variable Ranges. ................................. 46 Table 9

Recent Progress on the VASIMR

NASA Technical Reports Server (NTRS)

ChangDiaz, Franklin R.; Squire, J. P.; Ilin, A. V.; Jacobson, V. T.; Glover, T. W.; Baity, F. W.; Carter, M. D.; Goulding, R. H.; Breizman, B. N.

1999-01-01

Experimental and theoretical studies on the Variable Specific Impulse Magnetoplasma Rocket (VASIMR) have continued through a NASA led collaborative program involving several research groups. In the experimental area, performance characterization of the VASIMR helicon plasma source has been obtained over a portion of the parameter space, with helium and hydrogen propellant. Density (10(exp 18) - 10(exp 19)/ cubic meter) and temperature (5 eV) were measured at moderate degree of ionization in two separate experimental devices. Helicon design improvement and optimization will be discussed. Experiments with the ion cyclotron resonance heating (ICRH) subsection have begun and preliminary results will be discussed. Theoretical picture and integrated numerical simulation continue to be refined to account for the main physics elements of the VASIMR, including RF absorption and particle acceleration with subsequent detachment in the magnetic nozzle.

Control Design for an Advanced Geared Turbofan Engine

NASA Technical Reports Server (NTRS)

Chapman, Jeffryes W.; Litt, Jonathan S.

2017-01-01

This paper describes the design process for the control system of an advanced geared turbofan engine. This process is applied to a simulation that is representative of a 30,000 lbf thrust class concept engine with two main spools, ultra-high bypass ratio, and a variable area fan nozzle. Control system requirements constrain the non-linear engine model as it operates throughout its flight envelope of sea level to 40,000 ft and from 0 to 0.8 Mach. The control architecture selected for this project was developed from literature and reflects a configuration that utilizes a proportional integral controller integrated with sets of limiters that enable the engine to operate safely throughout its flight envelope. Simulation results show the overall system meets performance requirements without exceeding system operational limits.

Cylindrical diffuser performance using a truncated plug nozzle

NASA Technical Reports Server (NTRS)

Galanga, F. L.; Mueller, T. J.

1976-01-01

Cylindrical diffuser performance for a truncated plug nozzle without external flow was tested in a blowdown wind tunnel. The nozzle was designed for an exit Mach number of 1.9 and the plug was conical in shape from the throat and converged to the axis of symmetry at an angle of 10 degrees. The diffuser section was fashioned into two 13.97 cm lengths to facilitate boring of the duct diameter and to allow for testing of two different duct lengths. A slotted hypotube was installed in the base of the diffuser to measure pressure distribution down the centerline of the diffuser. The data obtained included: the typical centerline and sidewall pressure ratio variation along the diffuser, cell pressure ratio vs overall pressure ratio for long and short diffusers and a comparison of minimum experimental cell pressure ratio vs area ratio.

Debris control design achievements of the booster separation motors

NASA Technical Reports Server (NTRS)

Smith, G. W.; Chase, C. A.

1985-01-01

The stringent debris control requirements imposed on the design of the Space Shuttle booster separation motor are described along with the verification program implemented to ensure compliance with debris control objectives. The principal areas emphasized in the design and development of the Booster Separation Motor (BSM) relative to debris control were the propellant formulation and nozzle closures which protect the motors from aerodynamic heating and moisture. A description of the motor design requirements, the propellant formulation and verification program, and the nozzle closures design and verification are presented.

Rainfall simulators - innovations seeking rainfall uniformity and automatic flow rate measurements

NASA Astrophysics Data System (ADS)

Bauer, Miroslav; Kavka, Petr; Strouhal, Luděk; Dostál, Tomáš; Krása, Josef

2016-04-01

Field rainfall simulators are used worldwide for many experimental purposes, such as runoff generation and soil erosion research. At CTU in Prague a laboratory simulator with swinging nozzles VeeJet has been operated since 2001. Since 2012 an additional terrain simulator is being used with 4 fixed FullJet 40WSQ nozzles with 2,4 m spacing and operating over two simultaneously sprinkled experimental plots sizing 8x2 and 1x1 m. In parallel to other research projects a specific problem was solved: improving rainfall spatial uniformity and overall intensity and surface runoff measurements. These fundamental variables significantly affect investigated processes as well as resulting water balance of the plot, therefore they need to be determined as accurately as possible. Although the original nozzles setting produced (commonly used) Christiansen uniformity index CU over 80 %, detailed measurements proved this index insufficient and showed many unrequired rainfall extremes within the plot. Moreover the number of rainfall intensity scenarios was limited and some of them required problematic multi-pressure operation of the water distribution system. Therefore the simulator was subjected to many substantial changes in 2015. Innovations ranged from pump intensification to control unit upgrade. As essential change was considered increase in number of nozzles to 9 in total and reducing their spacing to 1,2 m. However new uniformity measurements did not bring any significant improvement. Tested scenarios showed equal standard deviations of interpolated intensity rasters and equal or slightly lower CU index. Imperfections of sprinkling nozzles were found to be the limiting factor. Still many other benefits were brought with the new setup. Whole experimental plot 10x2 m is better covered with the rainfall while the water consumption is retained. Nozzles are triggered in triplets, which enables more rainfall intensity scenarios. Water distribution system is more stable due to single pressure operating mode, which is ensured by the pressure probe controlled electromagnetic valve. Previous experiments implied the need of automatic continuous measurements of selected variables. To this end the control unit was equipped with a datalogger. In a several seconds time step it collects the values of water pressure, nozzle-valves operation, control point rainfall intensity from a tipping bucket rain gauge, topsoil moisture from several Theta ML2x probes and most recently the plot outlet runoff rate. For a continuous runoff rate measurement a 0,4-foot HS-flume was constructed and equipped with S18U ultrasonic sensor. Assemble setting was optimised both in flow rate laboratory flume and in laboratory rainfall simulator. Namely the rating curves for particular flume bottom slopes were derived. Employment of the flume in the terrain is scheduled for the experimental season 2016, but laboratory results already show sufficient measurement accuracy and are promising in terms of experimental campaigns simplification. The abovementioned activities have been supported by the research grants SGS14/180/OHK1/3T/11, QJ1530181, QJ1520265 and QJ1330118.

Numerical Evaluation of an Ejector-Enhanced Resonant Pulse Combustor with a Poppet Inlet Valve and a Converging Exhaust Nozzle

NASA Technical Reports Server (NTRS)

Yungster, Shaye; Paxson, Daniel E.; Perkins, Hugh D.

2016-01-01

A computational investigation of a pressure-gain combustor system for gas turbine applications is presented. The system consists of a valved pulse combustor and an ejector, housed within a shroud. The study focuses on two enhancements to previous models, related to the valve and ejector components. First, a new poppet inlet valve system is investigated, replacing the previously used reed valve configuration. Secondly, a new computational approach to approximating the effects of choked turbine inlet guide vanes present immediately downstream of the Ejector-Enhanced Resonant Pulse Combustor (EERPC) is investigated. Instead of specifying a back pressure at the EERPC exit boundary (as was done in previous studies) the new model adds a converging-diverging (CD) nozzle at the exit of the EERPC. The throat area of the CD nozzle can be adjusted to obtain the desired back pressure level and total mass flow rate. The results presented indicate that the new poppet valve configuration performs nearly as well as the original reed valve system, and that the addition of the CD nozzle is an effective method to approximate the exit boundary effects of a turbine present downstream of the EERPC. Furthermore, it is shown that the more acoustically reflective boundary imposed by a nozzle as compared to a constant pressure surface does not significantly affect operation or performance.

An Integration of the Turbojet and Single-Throat Ramjet

NASA Technical Reports Server (NTRS)

Trefny, C. J.; Benson, T. J.

1995-01-01

A turbine-engine-based hybrid propulsion system is described. Turbojet engines are integrated with a single-throat ramjet so as to minimize variable geometry and eliminate redundant propulsion components. The result is a simple, lightweight system that is operable from takeoff to high Mach numbers. Non-afterburning turbojets are mounted within the ramjet duct. They exhaust through a converging-diverging (C-D) nozzle into a common ramjet burner section. At low speed the ejector effect of the C-D nozzle aerodynamically isolates the relatively high pressure turbojet exhaust stream from the ramjet duct. As the Mach number increases, and the turbojet pressure ratio diminishes, the system is biased naturally toward ramjet operation. The common ramjet burner is fueled with hydrogen and thermally choked, thus avoiding the weight and complexity of a variable geometry, split-flow exhaust system. The mixed-compression supersonic inlet and subsonic diffuser are also common to both the turbojet and ramjet cycles. As the compressor face total temperature limit is approached, a two-position flap within the inlet is actuated, which closes off the turbojet inlet and provides increased internal contraction for ramjet operation. Similar actuation of the turbojet C-D nozzle flap completes the enclosure of the turbojet. Performance of the hybrid system is compared herein to that of the discrete turbojet and ramjet engines from takeoff to Mach 6. The specific impulse of the hybrid system falls below that of the non-integrated turbojet and ramjet because of ejector and Rayleigh losses. Unlike the discrete turbojet or ramjet however, the hybrid system produces thrust over the entire Mach number range. An alternate mode of operation for takeoff and low speed is also described. In this mode the C-D nozzle flap is deflected to a third position, which closes off the ramjet duct and eliminates the ejector total pressure loss.

Space Shuttle Main Engine structural analysis and data reduction/evaluation. Volume 6: Primary nozzle diffuser analysis

NASA Technical Reports Server (NTRS)

Foley, Michael J.

1989-01-01

The primary nozzle diffuser routes fuel from the main fuel valve on the Space Shuttle Main Engine (SSME) to the nozzle coolant inlet mainfold, main combustion chamber coolant inlet mainfold, chamber coolant valve, and the augmented spark igniters. The diffuser also includes the fuel system purge check valve connection. A static stress analysis was performed on the diffuser because no detailed analysis was done on this part in the past. Structural concerns were in the area of the welds because approximately 10 percent are in areas inaccessible by X-ray testing devices. Flow dynamics and thermodynamics were not included in the analysis load case. Constant internal pressure at maximum SSME power was used instead. A three-dimensional, finite element method was generated using ANSYS version 4.3A on the Lockheed VAX 11/785 computer to perform the stress computations. IDEAS Supertab on a Sun 3/60 computer was used to create the finite element model. Rocketdyne drawing number RS009156 was used for the model interpretation. The flight diffuser is denoted as -101. A description of the model, boundary conditions/load case, material properties, structural analysis/results, and a summary are included for documentation.

Acoustic testing of a 1.5 pressure ratio low tip speed fan with a serrated rotor (QEP fan B scale model). [reduction of engine noise

NASA Technical Reports Server (NTRS)

Kazin, S. B.; Paas, J. E.; Minzner, W. R.

1973-01-01

A scale model of the bypass flow region of a 1.5 pressure ratio, single stage, low tip speed fan was tested with a serrated rotor leading edge to determine its effects on noise generation. The serrated rotor was produced by cutting teeth into the leading edge of the nominal rotor blades. The effects of speed and exhaust nozzle area on the scale models noise characteristics were investigated with both the nominal rotor and serrated rotor. Acoustic results indicate the serrations reduced front quadrant PNL's at takeoff power. In particular, the 200 foot (61.0 m) sideline noise was reduced from 3 to 4 PNdb at 40 deg for nominal and large nozzle operation. However, the rear quadrant maximum sideline PNL's were increased 1.5 to 3 PNdb at approach thust and up to 2 PNdb at takeoff thust with these serrated rotor blades. The configuration with the serrated rotor produced the lowest maximum 200 foot (61.0 m) sideline PNL for any given thust when the large nozzle (116% of design area) was employed.

Performance characteristics of two multiaxis thrust-vectoring nozzles at Mach numbers up to 1.28

NASA Technical Reports Server (NTRS)

Wing, David J.; Capone, Francis J.

1993-01-01

The thrust-vectoring axisymmetric (VA) nozzle and a spherical convergent flap (SCF) thrust-vectoring nozzle were tested along with a baseline nonvectoring axisymmetric (NVA) nozzle in the Langley 16-Foot Transonic Tunnel at Mach numbers from 0 to 1.28 and nozzle pressure ratios from 1 to 8. Test parameters included geometric yaw vector angle and unvectored divergent flap length. No pitch vectoring was studied. Nozzle drag, thrust minus drag, yaw thrust vector angle, discharge coefficient, and static thrust performance were measured and analyzed, as well as external static pressure distributions. The NVA nozzle and the VA nozzle displayed higher static thrust performance than the SCF nozzle throughout the nozzle pressure ratio (NPR) range tested. The NVA nozzle had higher overall thrust minus drag than the other nozzles throughout the NPR and Mach number ranges tested. The SCF nozzle had the lowest jet-on nozzle drag of the three nozzles throughout the test conditions. The SCF nozzle provided yaw thrust angles that were equal to the geometric angle and constant with NPR. The VA nozzle achieved yaw thrust vector angles that were significantly higher than the geometric angle but not constant with NPR. Nozzle drag generally increased with increases in thrust vectoring for all the nozzles tested.

Relationship of stage mensuration data to the performance of new and used cascade impactors.

PubMed

Roberts, Daryl L; Romay, Francisco J

2005-01-01

Cascade impaction is a standard test method for characterizing the quality of inhalable drug products. The sizes of the nozzles on each stage of the impactor are the critical dimensions for the performance of the impactor. Compendial reference methods call for periodic measurement of the size of the nozzles on each stage, a procedure known as stage mensuration. There is however currently no guidance on acceptable mensuration criteria. We aim to remedy this situation by providing a sound basis for understanding and using mensuration data, be it for acceptance criteria for new impactors or for the setting of mensuration tolerances for in-use impactors. We first show that multi-nozzle impactor stages behave as if all of the nozzles are equal in size to an effective diameter, , that is composed of the area-mean and areamedian diameters, W* and , calculated directly from the individual nozzle diameters for all nozzles on a given stage (equation 1): W= (W*)(2/3) x (W)(1/3) (1). Hence, the effective diameter provides an intuitive and technically sound basis for setting acceptance criteria for new and in-use impactors. We tabulate these criteria for the Mark II eight-stage Andersen cascade impactor and the Next Generation Pharmaceutical Impactor in a manner similar to the tables of critical impactor dimensions published in EP Supplement 5.1 and in USP 28. For two different impactors or for one impactor measured at two different times (e.g., at manufacture and in use), we find that the D50 values of a given stage are related to the effective diameters by D(50,2)/D(50,1)= (W(2)/W(1))(3/2) (2). Using the stage mensuration data for new, as-manufactured NGIs, we compare the D(50 )values of the first 125 as-manufactured NGIs with those of the archivally calibrated NGI. We further establish that the archivally calibrated NGI has D(50) values within 0.3% of an entirely perfect, hypothetical NGI with all nozzles equal to the nominal nozzle diameters. We also apply the equations to a specific mensurated impactor to show that a used impactor with some nozzles outside of the original manufacturing specifications can have the same aerodynamic performance as a new impactor.

Evaluation of an automatic-timed insecticide application system for backyard mosquito control.

PubMed

Cilek, J E; Hallmon, C F; Johnson, R

2008-12-01

Several manufacturers and pest management companies have begun to market and install outdoor automatically timed insecticide application systems that claim to provide an envelope of protection against host-seeking mosquitoes within a defined area, e.g., residential backyards. A typical system consists of a multi-gallon reservoir attached to a continuous loop of plastic tubing with multiple single spray head nozzles. Nozzles are usually placed along the perimeter of a backyard in landscaping or other areas suitable for mosquito harborage. This array is then connected to a programmable electric pump set to automatically apply an insecticide at predetermined intervals. An operational field study was conducted to evaluate this technology using previously installed MistAway systems at.3 residences in northwestern Florida. This system applied a mist-like application of 0.05% AI synergized pyrethrins for 45 sec at dawn and again at dusk in each backyard. Twice-weekly collections from ABC suction light traps, baited with carbon dioxide, were used as the evaluation tool. Female mosquitoes from treatment backyards were compared with trap collections from 3 backyards without automatic misting systems used as controls. We found that weekly mosquito reduction was highly variable and ranged from 98% to 14% during the 35-wk study. Because the primary method of reduction by these application systems was not well understood, a MistAway system was installed in an outdoor simulated residential backyard to determine exposure pathway under controlled conditions with field cage and excised-leaf bioassays. Using laboratory-reared females of Aedes albopictus and Culex quinquefasciatus in those assays, we found that reduction by the MistAway system was primarily achieved by direct exposure of the mosquitoes to the insecticide application and not from residual deposits on treated vegetation.

Predictive GT-Power Simulation for VNT Matching on a 1.6 L Turbocharged GDI Engine

EPA Science Inventory

The thermal efficiency benefits of low-pressure (LP) exhaust gas recirculation (EGR) in spark-ignition engine combustion are well known. One of the greatest barriers facing adoption of LP-EGR for high power-density applications is the challenge of boosting. Variable nozzle turbin...

Fuel nozzle assembly

DOEpatents

Johnson, Thomas Edward [Greer, SC; Ziminsky, Willy Steve [Simpsonville, SC; Lacey, Benjamin Paul [Greer, SC; York, William David [Greer, SC; Stevenson, Christian Xavier [Inman, SC

2011-08-30

A fuel nozzle assembly is provided. The assembly includes an outer nozzle body having a first end and a second end and at least one inner nozzle tube having a first end and a second end. One of the nozzle body or nozzle tube includes a fuel plenum and a fuel passage extending therefrom, while the other of the nozzle body or nozzle tube includes a fuel injection hole slidably aligned with the fuel passage to form a fuel flow path therebetween at an interface between the body and the tube. The nozzle body and the nozzle tube are fixed against relative movement at the first ends of the nozzle body and nozzle tube, enabling the fuel flow path to close at the interface due to thermal growth after a flame enters the nozzle tube.

Gaseous hydrogen/oxygen injector performance characterization

NASA Technical Reports Server (NTRS)

Degroot, W. A.; Tsuei, H. H.

1994-01-01

Results are presented of spontaneous Raman scattering measurements in the combustion chamber of a 110 N thrust class gaseous hydrogen/oxygen rocket. Temperature, oxygen number density, and water number density profiles at the injector exit plane are presented. These measurements are used as input profiles to a full Navier-Stokes computational fluid dynamics (CFD) code. Predictions of this code while using the measured profiles are compared with predictions while using assumed uniform injector profiles. Axial and radial velocity profiles derived from both sets of predictions are compared with Rayleigh scattering measurements in the exit plane of a 33:1 area ratio nozzle. Temperature and number density Raman scattering measurements at the exit plane of a test rocket with a 1:1.36 area ratio nozzle are also compared with results from both sets of predictions.

Model aerodynamic test results for two variable cycle engine coannular exhaust systems at simulated takeoff and cruise conditions. Comprehensive data report. Volume 2: Tabulated aeroynamic data book 1

NASA Technical Reports Server (NTRS)

Nelson, D. P.

1981-01-01

Tabulated data from wind tunnel tests conducted to evaluate the aerodynamic performance of an advanced coannular exhaust nozzle for a future supersonic propulsion system are presented. Tests were conducted with two test configurations: (1) a short flap mechanism for fan stream control with an isentropic contoured flow splitter, and (2) an iris fan nozzle with a conical flow splitter. Both designs feature a translating primary plug and an auxiliary inlet ejector. Tests were conducted at takeoff and simulated cruise conditions. Data were acquired at Mach numbers of 0, 0.36, 0.9, and 2.0 for a wide range of nozzle operating conditions. At simulated supersonic cruise, both configurations demonstrated good performance, comparable to levels assumed in earlier advanced supersonic propulsion studies. However, at subsonic cruise, both configurations exhibited performance that was 6 to 7.5 percent less than the study assumptions. At takeoff conditions, the iris configuration performance approached the assumed levels, while the short flap design was 4 to 6 percent less. Data are provided through test run 25.

Measuring air core characteristics of a pressure-swirl atomizer via a transparent acrylic nozzle at various Reynolds numbers

DOE Office of Scientific and Technical Information (OSTI.GOV)

Lee, Eun J.; Oh, Sang Youp; Kim, Ho Y.

2010-11-15

Because of thermal fluid-property dependence, atomization stability (or flow regime) can change even at fixed operating conditions when subject to temperature change. Particularly at low temperatures, fuel's high viscosity can prevent a pressure-swirl (or simplex) atomizer from sustaining a centrifugal-driven air core within the fuel injector. During disruption of the air core inside an injector, spray characteristics outside the nozzle reflect a highly unstable, nonlinear mode where air core length, Sauter mean diameter (SMD), cone angle, and discharge coefficient variability. To better understand injector performance, these characteristics of the pressure-swirl atomizer were experimentally investigated and data were correlated to Reynoldsmore » numbers (Re). Using a transparent acrylic nozzle, the air core length, SMD, cone angle, and discharge coefficient are observed as a function of Re. The critical Reynolds numbers that distinguish the transition from unstable mode to transitional mode and eventually to a stable mode are reported. The working fluids are diesel and a kerosene-based fuel, referred to as bunker-A. (author)« less

Magnetic Nozzle and Plasma Detachment Experiment

NASA Technical Reports Server (NTRS)

Chavers, Gregory; Dobson, Chris; Jones, Jonathan; Martin, Adam; Bengtson, Roger D.; Briezman, Boris; Arefiev, Alexey; Cassibry, Jason; Shuttpelz, Branwen; Deline, Christopher

2006-01-01

High power plasma propulsion can move large payloads for orbit transfer (such as the ISS), lunar missions, and beyond with large savings in fuel consumption owing to the high specific impulse. At high power, lifetime of the thruster becomes an issue. Electrodeless devices with magnetically guided plasma offer the advantage of long life since magnetic fields confine the plasma radially and keep it from impacting the material surfaces. For decades, concerns have been raised about the plasma remaining attached to the magnetic field and returning to the vehicle along the closed magnetic field lines. Recent analysis suggests that this may not be an issue of the magnetic field is properly shaped in the nozzle region and the plasma has sufficient energy density to stretch the magnetic field downstream. An experiment was performed to test the theory regarding the Magneto-hydrodynamic (MHD) detachment scenario. Data from this experiment will be presented. The Variable Specific Impulse Magnetoplasma Rocket (VASIMR) being developed by the Ad Astra Rocket Company uses a magnetic nozzle as described above. The VASIMR is also a leading candidate for exploiting an electric propulsion test platform being considered for the ISS.

Rock Cutting Depth Model Based on Kinetic Energy of Abrasive Waterjet

NASA Astrophysics Data System (ADS)

Oh, Tae-Min; Cho, Gye-Chun

2016-03-01

Abrasive waterjets are widely used in the fields of civil and mechanical engineering for cutting a great variety of hard materials including rocks, metals, and other materials. Cutting depth is an important index to estimate operating time and cost, but it is very difficult to predict because there are a number of influential variables (e.g., energy, geometry, material, and nozzle system parameters). In this study, the cutting depth is correlated to the maximum kinetic energy expressed in terms of energy (i.e., water pressure, water flow rate, abrasive feed rate, and traverse speed), geometry (i.e., standoff distance), material (i.e., α and β), and nozzle system parameters (i.e., nozzle size, shape, and jet diffusion level). The maximum kinetic energy cutting depth model is verified with experimental test data that are obtained using one type of hard granite specimen for various parameters. The results show a unique curve for a specific rock type in a power function between cutting depth and maximum kinetic energy. The cutting depth model developed here can be very useful for estimating the process time when cutting rock using an abrasive waterjet.

Feedback Control of a Morphing Chevron for Takeoff and Cruise Noise Reduction

NASA Technical Reports Server (NTRS)

Cabell, Randolph H.; Schiller, Noah H.; Mabe, James H.; Ruggeri, Robert T.; Butler, G. W.

2004-01-01

Noise from commercial high-bypass ratio turbofan engines is generated by turbulent mixing of the hot jet exhaust, fan stream, and ambient air. Serrated aerodynamic devices, known as chevrons, along the trailing edges of a jet engine primary and secondary exhaust nozzle have been shown to reduce jet noise at takeoff and shock-cell noise at cruise conditions. Their optimum shape is a finely tuned compromise between noise-benefit and thrust-loss. The design of a full scale Variable Geometry Chevron (VGC) fan-nozzle incorporating Shape Memory Alloy (SMA) actuators is described in a companion paper. This paper describes the development and testing of a proportional-integral control system that regulates the heating of the SMA actuators to control the VGC s tip immersion. The VGC and control system were tested under representative flow conditions in Boeing s Nozzle Test Facility (NTF). Results from the NTF test which demonstrate controllable immersion of the VGC are described. The paper also describes the correlation between strains and temperatures on the chevron with a photogrammetric measurement of the chevron's tip immersion.

Static and wind tunnel model tests for the development of externally blown flap noise reduction techniques

NASA Technical Reports Server (NTRS)

Pennock, A. P.; Swift, G.; Marbert, J. A.

1975-01-01

Externally blown flap models were tested for noise and performance at one-fifth scale in a static facility and at one-tenth scale in a large acoustically-treated wind tunnel. The static tests covered two flap designs, conical and ejector nozzles, third-flap noise-reduction treatments, internal blowing, and flap/nozzle geometry variations. The wind tunnel variables were triple-slotted or single-slotted flaps, sweep angle, and solid or perforated third flap. The static test program showed the following noise reductions at takeoff: 1.5 PNdB due to treating the third flap; 0.5 PNdB due to blowing from the third flap; 6 PNdB at flyover and 4.5 PNdB in the critical sideline plane (30 deg elevation) due to installation of the ejector nozzle. The wind tunnel program showed a reduction of 2 PNdB in the sideline plane due to a forward speed of 43.8 m/s (85 kn). The best combination of noise reduction concepts reduced the sideline noise of the reference aircraft at constant field length by 4 PNdB.

Thrust reverser design studies for an over-the-wing STOL transport

NASA Technical Reports Server (NTRS)

Ammer, R. C.; Sowers, H. D.

1977-01-01

Aerodynamic and acoustics analytical studies were conducted to evaluate three thrust reverser designs for potential use on commercial over-the-wing STOL transports. The concepts were: (1) integral D nozzle/target reverser, (2) integral D nozzle/top arc cascade reverser, and (3) post exit target reverser integral with wing. Aerodynamic flowpaths and kinematic arrangements for each concept were established to provide a 50% thrust reversal capability. Analytical aircraft stopping distance/noise trade studies conducted concurrently with flow path design showed that these high efficiency reverser concepts are employed at substantially reduced power settings to meet noise goals of 100 PNdB on a 152.4 m sideline and still meet 609.6 m landing runway length requirements. From an overall installation standpoint, only the integral D nozzle/target reverser concept was found to penalize nacelle cruise performance; for this concept a larger nacelle diameter was required to match engine cycle effective area demand in reverse thrust.

Measurement of Vibrational Non-Equilibrium in a Supersonic Freestream Using Dual-Pump CARS

NASA Technical Reports Server (NTRS)

Cutler, Andrew D.; Magnotti, Gaetano; Cantu, Luca M. L.; Gallo, Emanuela C. A.; Danehy, Paul M.; Burle, Rob; Rockwell, Robert; Goyne, Christopher; McDaniel, James

2012-01-01

Measurements have been conducted at the University of Virginia Supersonic Combustion Facility of the flow in a constant area duct downstream of a Mach 2 nozzle, where the airflow has first been heated to approximately 1200 K. Dual-pump CARS was used to acquire rotational and vibrational temperatures of N2 and O2 at two planes in the duct at different downstream distances from the nozzle exit. Wall static pressures in the nozzle are also reported. With a flow of clean air, the vibrational temperature of N2 freezes at close to the heater stagnation temperature, while the O2 vibrational temperature is about 1000 K. The results are well predicted by computational fluid mechanics models employing separate "lumped" vibrational and translational/rotational temperatures. Experimental results are also reported for a few percent steam addition to the air and the effect of the steam is to bring the flow to thermal equilibrium.

Effect of the nozzle tip’s geometrical shape on electrospray deposition of organic thin films

NASA Astrophysics Data System (ADS)

Ueda, Hiroyuki; Takeuchi, Keita; Kikuchi, Akihiko

2017-04-01

Electrospray deposition (ESD) is a favorable wet fabrication technique for organic thin films. We investigated the effects of the nozzle tip’s geometrical shape on the spraying properties of an organic solution used for ESD. Five types of cylindrical metal nozzles with zero (flat end) to four protrusions at the tips were prepared for depositing a solution of a small-molecule compound, tris(8-hydroxyquinolinato)aluminum (Alq3) solution. We confirmed that the diameter of the deposited droplets and their size dispersion decreased with an increase in the number of protrusions. The area occupation ratio of small droplets with a diameter smaller than 2 µm increased from 21 to 83% as the number of protrusions was increased from zero to four. The surface roughness root mean square of 60-nm-thick Alq3 films substantially improved from 32.5 to 6.8 nm with increasing number of protrusions.

Mixing Enhancement by Tabs in Round Supersonic Jets

NASA Technical Reports Server (NTRS)

Seiner, John M.; Grosch, C. E.

1998-01-01

The objective of this study was to analyze jet plume mass flow entrainment rates associated with the introduction of counter-rotating streamwise vorticity by prism shaped devices (tabs) located at the lip of the nozzle. We have examined the resulting mixing process through coordinated experimental tests and numerical simulations of the supersonic flow from a model axisymmetric nozzle. In the numerical simulations, the total induced vorticity was held constant while varying the distribution of counter-rotating vorticity around the nozzle lip training edge. In the experiment, the number of tabs applied was varied while holding the total projected area constant. Evaluations were also conducted on initial vortex strength. The results of this work show that the initial growth rate of the jet shear layer is increasingly enhanced as more tabs are added, but that the lowest tab count results in the largest entrained mass flow. The numerical simulations confirm these results.

Measured opening characteristics of an electromagnetically opened diaphragm for the Langley expansion tunnel

NASA Technical Reports Server (NTRS)

Moore, J. A.

1976-01-01

Results from an experimental study of the opening characteristics of an electromagnetically opened, 15.24 cm diameter diaphragm are presented. This diaphragm consists of a polyester film bonded to a preformed wire and is opened by passing a current pulse (capacitor discharge) through the wire. The diaphragm separates the acceleration section of the expansion tunnel from the nozzle so that the nozzle may be at a lower pressure than the acceleration section prior to a test. Opening times and cleanness of the opened area were examined for dependence on diaphragm thickness, on wire diameter, on technique of bonding the wire to the diaphragm, and on voltage and energy level of the energy source. Time histories of the pitot pressure measured at the expansion-tunnel nozzle entrance location are presented for (1) no diaphragm, (2) a flow-opened diaphragm, and (3) an electromagnetically opened diaphragm.

Charged particle concepts for fog dispersion

NASA Technical Reports Server (NTRS)

Frost, W.; Collins, F. G.; Koepf, D.

1981-01-01

Charged particle techniques hold promise for dispersing warm fog in the terminal area of commercial airports. This report focuses on features of the charged particle technique which require further study. The basic physical principles of the technique and the major verification experiments carried out in the past are described. The fundamentals of the nozzle operation are given. The nozzle characteristics and the theory of particle charging in the nozzle are discussed, including information from extensive literature on electrostatic precipitation relative to environmental pollution control and a description of some preliminary reported analyses on the jet characteristics and interaction with neighboring jets. The equation governing the transfer of water substances and of electrical charge is given together with a brief description of several semi-empirical, mathematical expressions necessary for the governing equations. The necessary ingredients of a field experiment to verify the system once a prototype is built are described.

Method and apparatus for improved melt flow during continuous strip casting

DOEpatents

Follstaedt, Donald W.; King, Edward L.; Schneider, Ken C.

1991-11-12

The continuous casting of metal strip using the melt overflow process is improved by controlling the weir conditions in the nozzle to provide a more uniform flow of molten metal across the width of the nozzle and reducing the tendency for freezing of metal along the interface with refractory surfaces. A weir design having a sloped rear wall and tapered sidewalls and critical gap controls beneath the weir has resulted in the drastic reduction in edge tearing and a significant improvement in strip uniformity. The floor of the container vessel is preferably sloped and the gap between the nozzle and the rotating substrate is critically controlled. The resulting flow patterns observed with the improved casting process have reduced thermal gradients in the bath, contained surface slag and eliminated undesirable solidification near the discharge area by increasing the flow rates at those points.

Method and apparatus for improved melt flow during continuous strip casting

DOEpatents

Follstaedt, D.W.; King, E.L.; Schneider, K.C.

1991-11-12

The continuous casting of metal strip using the melt overflow process is improved by controlling the weir conditions in the nozzle to provide a more uniform flow of molten metal across the width of the nozzle and reducing the tendency for freezing of metal along the interface with refractory surfaces. A weir design having a sloped rear wall and tapered sidewalls and critical gap controls beneath the weir has resulted in the drastic reduction in edge tearing and a significant improvement in strip uniformity. The floor of the container vessel is preferably sloped and the gap between the nozzle and the rotating substrate is critically controlled. The resulting flow patterns observed with the improved casting process have reduced thermal gradients in the bath, contained surface slag and eliminated undesirable solidification near the discharge area by increasing the flow rates at those points. 8 figures.

Noncontact Capacitive Clearance Control System For Laser Cutting Machines

NASA Astrophysics Data System (ADS)

Topkaya, Ahmet; Schmall, Karl-Heinz; Majoli, Ralf

1989-03-01

For a continuous high quality laser cut, it is necessary among other things to position the focal point of the laser beam correctly. This means that a constant clearance between the cutting head and the workpiece with a tolerance of +/- 0.Imm must he ensured. When cutting corrugated automobile bodysheet for example, a good quality cut can only be achieved with automatic clearance control. In the following, a method of automatic clearance control is described using the assistance of a noncontact capacitive sensor system. The copper nozzle of the laser cutting head acts as the electrode of the clearance sensor. The nozzle electrode and the workpiece build a small variable capacitance depending on the clearance. A change of clearance also changes the capacitance, which in turn influences a high frequency oscillator circuit. This shift in frequency is then converted into an analogue DC signal, which can be used to operate a servo motor control for the positioning of the laser cutting head in a closed loop servo system. Laser cutting heads with clearance sensor nozzles of different shapes, suited fur most applications in the industry, with focal lengths from 2.5" to 5" have been developed. They are capable to cut metal sheet from 0.2 to 12 mm of thickness, using CO2-lasers with output power up to 2.5 kW. For special applications involving difficult workpiece topographies in automobile production applications special "trunk" nozzles have been developed. For 5-axis cutting machines and robots, new laser cutting heads with integrated nozzle sensors in combination with a high dynamic Z-axis motor drive are in a stage of development.

Mixing of Supersonic Streams

NASA Technical Reports Server (NTRS)

Hawk, C. W.; Landrum, D. B.; Muller, S.; Turner, M.; Parkinson, D.

1998-01-01

The Strutjet approach to Rocket Based Combined Cycle (RBCC) propulsion depends upon fuel-rich flows from the rocket nozzles and turbine exhaust products mixing with the ingested air for successful operation in the ramjet and scramjet modes. It is desirable to delay this mixing process in the air-augmented mode of operation present during low speed flight. A model of the Strutjet device has been built and is undergoing test to investigate the mixing of the streams as a function of distance from the Strutjet exit plane during simulated low speed flight conditions. Cold flow testing of a 1/6 scale Strutjet model is underway and nearing completion. Planar Laser Induced Fluorescence (PLIF) diagnostic methods are being employed to observe the mixing of the turbine exhaust gas with the gases from both the primary rockets and the ingested air simulating low speed, air augmented operation of the RBCC. The ratio of the pressure in the turbine exhaust duct to that in the rocket nozzle wall at the point of their intersection is the independent variable in these experiments. Tests were accomplished at values of 1.0, 1.5 and 2.0 for this parameter. Qualitative results illustrate the development of the mixing zone from the exit plane of the model to a distance of about 19 equivalent rocket nozzle exit diameters downstream. These data show the mixing to be confined in the vertical plane for all cases, The lateral expansion is more pronounced at a pressure ratio of 1.0 and suggests that mixing with the ingested flow would be likely beginning at a distance of 7 nozzle exit diameters downstream of the nozzle exit plane.

Mixing of Supersonic Jets in a RBCC Strutjet Propulsion System

NASA Technical Reports Server (NTRS)

Muller, S.; Hawk, Clark W.; Bakker, P. G.; Parkinson, D.; Turner, M.

1998-01-01

The Strutjet approach to Rocket Based Combined Cycle (RBCC) propulsion depends upon fuel-rich flows from the rocket nozzles and turbine exhaust products mixing with the ingested air for successful operation in the ramjet and scramjet modes. It is desirable to delay this mixing process in the air-augmented mode of operation present during take-off and low speed flight. A scale model of the Strutjet device was built and tested to investigate the mixing of the streams as a function of distance from the Strut exit plane in simulated sea level take-off conditions. The Planar Laser Induced Fluorescence (PLIF) diagnostic method has been employed to observe the mixing of the turbine exhaust gas with the gases from both the primary rockets and the ingested air. The ratio of the pressure in the turbine exhaust to that in the rocket nozzle wall at the point where the two jets meet, is the independent variable in these experiments. Tests were accomplished at values of 1.0 (the original design point), 1.5 and 2.0 for this parameter at 8 locations downstream of the rocket nozzle exit. The results illustrate the development of the mixing zone from the exit plane of the strut to a distance of about 18 equivalent rocket nozzle exit diameters downstream (18"). These images show the turbine exhaust to be confined until a short distance downstream. The expansion into the ingested air is more pronounced at a pressure ratio of 1.0 and 1.5 and shows that mixing with this air would likely begin at a distance of 2" downstream of the nozzle exit plane. Of the pressure ratios tested in this research, 2.0 is the best value for delaying the mixing at the operating conditions considered.

The Effects of Blowing Over Various Trailing-edge Flaps on an NACA 0006 Airfoil Section, Comparisons with Various Types of Flaps on other Airfoil Sections, and an Analysis of Flow and Power Relationships for Blowing Systems

NASA Technical Reports Server (NTRS)

Dods, J. B., Jr.; Watson, E. C.

1976-01-01

The results are presented of a two-dimensional investigation conducted to determine the effect of blowing over various types of trailing-edge flaps on a wing having the NACA 0006 airfoil section and a drooped-nose flap. The position and profile of the trailing-edge flap, the nozzle height, and the location of the flap with respect to the nozzle were found to be important variables. Data from many investigations were used to make an evaluation of the effects of blowing on lift. An analysis was made of flow and power relationships for blowing systems.

A tandem mirror hybrid plume plasma propulsion facility

NASA Technical Reports Server (NTRS)

Yang, T. F.; Krueger, W. A.; Peng, S.; Urbahn, J.; Chang-Diaz, F. R.

1988-01-01

This paper discusses a novel concept in electrodeless plasma propulsion, in which the materials problems are ameliorated by an electrodeless magnetic confinement scheme borrowed from the tandem mirror approach to controlled thermonuclear fusion. The concept also features a two-stage magnetic nozzle with an annular hypersonic coaxial gas injector near the throat. The nozzle produces hybrid plume by the coaxial injection of hypersonic neutral gas, and the gas layer thus formed protects the material walls from the hot plasma and, through increased collisions, helps detach it from the diverging magnetic field. The tandem mirror plasma propulsion facility is capable of delivering a variable I(sp). The results of numerical simulation of this concept are presented together with those from an experimental tandem-mirror plasma propulsion device.

Effect of aviation fuel type and fuel injection conditions on the spray characteristics of pressure swirl and hybrid air blast fuel injectors

NASA Astrophysics Data System (ADS)

Feddema, Rick

Feddema, Rick T. M.S.M.E., Purdue University, December 2013. Effect of Aviation Fuel Type and Fuel Injection Conditions on the Spray Characteristics of Pressure Swirl and Hybrid Air Blast Fuel Injectors. Major Professor: Dr. Paul E. Sojka, School of Mechanical Engineering Spray performance of pressure swirl and hybrid air blast fuel injectors are central to combustion stability, combustor heat management, and pollutant formation in aviation gas turbine engines. Next generation aviation gas turbine engines will optimize spray atomization characteristics of the fuel injector in order to achieve engine efficiency and emissions requirements. Fuel injector spray atomization performance is affected by the type of fuel injector, fuel liquid properties, fuel injection pressure, fuel injection temperature, and ambient pressure. Performance of pressure swirl atomizer and hybrid air blast nozzle type fuel injectors are compared in this study. Aviation jet fuels, JP-8, Jet A, JP-5, and JP-10 and their effect on fuel injector performance is investigated. Fuel injector set conditions involving fuel injector pressure, fuel temperature and ambient pressure are varied in order to compare each fuel type. One objective of this thesis is to contribute spray patternation measurements to the body of existing drop size data in the literature. Fuel droplet size tends to increase with decreasing fuel injection pressure, decreasing fuel injection temperature and increasing ambient injection pressure. The differences between fuel types at particular set conditions occur due to differences in liquid properties between fuels. Liquid viscosity and surface tension are identified to be fuel-specific properties that affect the drop size of the fuel. An open aspect of current research that this paper addresses is how much the type of aviation jet fuel affects spray atomization characteristics. Conventional aviation fuel specifications are becoming more important with new interest in alternative fuels. Optical patternation data and line of sight laser diffraction data show that there is significant difference between jet fuels. Particularly at low fuel injection pressures (0.345 MPa) and cold temperatures (-40 C), the patternation data shows that the total surface area in the spray at 38.1 mm from the pressure swirl injector for the JP-10 fuel type is one-sixth the amount of the JP-8. Finally, this study compares the atomizer performance of a pressure swirl nozzle to a hybrid air blast nozzle. The total surface area for both the hybrid air blast nozzle and the pressure swirl nozzle show a similar decline in atomization performance at low fuel injection pressures and cold temperatures. However, the optical patternator radial profile data and the line of sight laser diffraction data show that the droplet size and spray distribution data are less affected by injection conditions and fuel type in the hybrid air blast nozzle, than they are in the pressure swirl nozzle. One explanation is that the aerodynamic forces associated with the swirler on the hybrid air blast nozzle control the distribution droplets in the spray. This is in contrast to the pressure swirl nozzle droplet distribution that is controlled by internal geometry and droplet ballistics.

Experimental aerodynamic and acoustic model testing of the Variable Cycle Engine (VCE) testbed coannular exhaust nozzle system: Comprehensive data report

NASA Technical Reports Server (NTRS)

Nelson, D. P.; Morris, P. M.

1980-01-01

The component detail design drawings of the one sixth scale model of the variable cycle engine testbed demonstrator exhaust syatem tested are presented. Also provided are the basic acoustic and aerodynamic data acquired during the experimental model tests. The model drawings, an index to the acoustic data, an index to the aerodynamic data, tabulated and graphical acoustic data, and the tabulated aerodynamic data and graphs are discussed.

Results of Aero/Acoustic Tests and Analytical Studies of a Two-Dimensional Eight-Lobe Mixer-Ejector Exhaust Nozzle at Takeoff Conditions

NASA Technical Reports Server (NTRS)

Harrington, Douglas (Technical Monitor); Schweiger, P.; Stern, A.; Gamble, E.; Barber, T.; Chiappetta, L.; LaBarre, R.; Salikuddin, M.; Shin, H.; Majjigi, R.

2005-01-01

Hot flow aero-acoustic tests were conducted with Pratt & Whitney's High-Speed Civil Transport (HSCT) Mixer-Ejector Exhaust Nozzles by General Electric Aircraft Engines (GEAE) in the GEAE Anechoic Freejet Noise Facility (Cell 41) located in Evendale, Ohio. The tests evaluated the impact of various geometric and design parameters on the noise generated by a two-dimensional (2-D) shrouded, 8-lobed, mixer-ejector exhaust nozzle. The shrouded mixer-ejector provides noise suppression by mixing relatively low energy ambient air with the hot, high-speed primary exhaust jet. Additional attenuation was obtained by lining the shroud internal walls with acoustic panels, which absorb acoustic energy generated during the mixing process. Two mixer designs were investigated, the high mixing "vortical" and aligned flow "axial", along with variations in the shroud internal mixing area ratios and shroud length. The shrouds were tested as hardwall or lined with acoustic panels packed with a bulk absorber. A total of 21 model configurations at 1:11.47 scale were tested. The models were tested over a range of primary nozzle pressure ratios and primary exhaust temperatures representative of typical HSCT aero thermodynamic cycles. Static as well as flight simulated data were acquired during testing. A round convergent unshrouded nozzle was tested to provide an acoustic baseline for comparison to the test configurations. Comparisons were made to previous test results obtained with this hardware at NASA Glenn's 9- by 15-foot low-speed wind tunnel (LSWT). Laser velocimetry was used to investigate external as well as ejector internal velocity profiles for comparison to computational predictions. Ejector interior wall static pressure data were also obtained. A significant reduction in exhaust system noise was demonstrated with the 2-D shrouded nozzle designs.

Spray drift reduction under Southern European conditions: a pilot study in the Ecopest Project in Greece.

PubMed

Kasiotis, Konstantinos M; Glass, C Richard; Tsakirakis, Angelos N; Machera, Kyriaki

2014-05-01

The objective of this work was to generate spray drift data from pesticide application in the field comparing spray drift from traditional equipment with emerging, anti-drift technologies. The applications were carried out in the Kopais area in central Greece. Currently few data exist as regards to pesticide spray drift in Southern European conditions. This work details the data for ground and airborne deposition of spray drift using the methodology developed in the UK by the Food and Environment Research Agency (FERA). Three trials were performed in two days using sunset yellow dye which deposited on dosimeters placed at specific distances from the edge of the sprayer boom. The application was carried out with a tractor mounted boom sprayer, which was of local manufacture, as were the nozzles of Trial I, being flat fan brass nozzles. For Trials II and III anti-drift nozzles were used. The boom sprayers were used with the settings as employed by the farmers for the routine pesticide applications. The results of this work indicate that drift was significantly reduced when anti-drift nozzles were utilized. Copyright © 2014 Elsevier B.V. All rights reserved.

Insights into the Behavior of Potential Structural Failures Originating from Localized High Stress Regions in Configurations Relevant to Solid Rocket Motor Nozzles

NASA Technical Reports Server (NTRS)

McCutcheon, David Matthew

2017-01-01

During the structural certification effort for the Space Launch System solid rocket booster nozzle, it was identified that no consistent method for addressing local negative margins of safety in non-metallic materials had been developed. Relevant areas included bond-line terminations and geometric features in the composite nozzle liners. In order to gain understanding, analog test specimens were designed that very closely mimic the conditions in the actual full scale hardware. Different locations in the nozzle were represented by different analog specimen designs. This paper describes those tests and corresponding results. Finite element analysis results for the tests are presented. Strain gage correlation of the analysis to the test results is addressed. Furthermore, finite fracture mechanics (a coupled stress and energy failure criterion) is utilized to predict the observed crack pop-in loads for the different configurations. The finite fracture mechanics predictions are found to be within a 10% error relative to the average measured pop-in load for each of four configurations. Initiation locations, arrest behaviors, and resistances to further post-arrest crack propagation are also discussed.

Modifications to the nozzle test chamber to extend nozzle static-test capability

NASA Technical Reports Server (NTRS)

Keyes, J. W.

1985-01-01

The nozzle test chamber was modified to provide a high-pressure-ratio nozzle static-test capability. Experiments were conducted to determine the range of the ratio of nozzle total pressure to chamber pressure and to make direct nozzle thrust measurements using a three-component strain-gage force balance. Pressure ratios from 3 to 285 were measured with several axisymmetric nozzles at a nozzle total pressure of 15 to 190 psia. Devices for measuring system mass flow were calibrated using standard axisymmetric convergent choked nozzles. System mass-flow rates up to 10 lbm/sec are measured. The measured thrust results of these nozzles are in good agreement with one-dimensional theoretical predictions for convergent nozzles.

The role of modern control theory in the design of controls for aircraft turbine engines

NASA Technical Reports Server (NTRS)

Zeller, J.; Lehtinen, B.; Merrill, W.

1982-01-01

The development, applications, and current research in modern control theory (MCT) are reviewed, noting the importance for fuel-efficient operation of turbines with variable inlet guide vanes, compressor stators, and exhaust nozzle area. The evolution of multivariable propulsion control design is examined, noting a basis in a matrix formulation of the differential equations defining the process, leading to state space formulations. Reports and papers which appeared from 1970-1982 which dealt with problems in MCT applications to turbine engine control design are outlined, including works on linear quadratic regulator methods, frequency domain methods, identification, estimation, and model reduction, detection, isolation, and accommodation, and state space control, adaptive control, and optimization approaches. Finally, NASA programs in frequency domain design, sensor failure detection, computer-aided control design, and plant modeling are explored

Procedures for generation and reduction of linear models of a turbofan engine

NASA Technical Reports Server (NTRS)

Seldner, K.; Cwynar, D. S.

1978-01-01

A real time hybrid simulation of the Pratt & Whitney F100-PW-F100 turbofan engine was used for linear-model generation. The linear models were used to analyze the effect of disturbances about an operating point on the dynamic performance of the engine. A procedure that disturbs, samples, and records the state and control variables was developed. For large systems, such as the F100 engine, the state vector is large and may contain high-frequency information not required for control. This, reducing the full-state to a reduced-order model may be a practicable approach to simplifying the control design. A reduction technique was developed to generate reduced-order models. Selected linear and nonlinear output responses to exhaust-nozzle area and main-burner fuel flow disturbances are presented for comparison.

Convoluted nozzle design for the RL10 derivative 2B engine

NASA Technical Reports Server (NTRS)

1985-01-01

The convoluted nozzle is a conventional refractory metal nozzle extension that is formed with a portion of the nozzle convoluted to show the extendible nozzle within the length of the rocket engine. The convoluted nozzle (CN) was deployed by a system of four gas driven actuators. For spacecraft applications the optimum CN may be self-deployed by internal pressure retained, during deployment, by a jettisonable exit closure. The convoluted nozzle is included in a study of extendible nozzles for the RL10 Engine Derivative 2B for use in an early orbit transfer vehicle (OTV). Four extendible nozzle configurations for the RL10-2B engine were evaluated. Three configurations of the two position nozzle were studied including a hydrogen dump cooled metal nozzle and radiation cooled nozzles of refractory metal and carbon/carbon composite construction respectively.

Figuring of plano-elliptical neutron focusing mirror by local wet etching.

PubMed

Yamamura, Kazuya; Nagano, Mikinori; Takai, Hiroyuki; Zettsu, Nobuyuki; Yamazaki, Dai; Maruyama, Ryuji; Soyama, Kazuhiko; Shimada, Shoichi

2009-04-13

Local wet etching technique was proposed to fabricate high-performance aspherical mirrors. In this process, only the limited area facing to the small nozzle is removed by etching on objective surface. The desired objective shape is deterministically fabricated by performing the numerically controlled scanning of the nozzle head. Using the technique, a plano-elliptical mirror to focus the neutron beam was successfully fabricated with the figure accuracy of less than 0.5 microm and the focusing gain of 6. The strong and thin focused neutron beam is expected to be a useful tool for the analyses of various material properties.

The prediction of noise and installation effects of high-subsonic dual-stream jets in flight

NASA Astrophysics Data System (ADS)

Saxena, Swati

Both military and civil aircraft in service generate high levels of noise. One of the major contributors to this noise generated from the aircraft is the jet engine exhaust. This makes the study of jet noise and methods to reduce jet noise an active research area with the aim of designing quieter military and commercial aircraft. The current stringent aircraft noise regulations imposed by the Federal Aviation Administration (FAA) and other international agencies, have further raised the need to perform accurate jet noise calculations for more reliable estimation of the jet noise sources. The main aim of the present research is to perform jet noise simulations of single and dual-stream jets with engineering accuracy and assess forward flight effects on the jet noise. Installation effects such as caused by the pylon are also studied using a simplified pylon nozzle configuration. Due to advances in computational power, it has become possible to perform turbulent flow simulations of high speed jets, which leads to more accurate noise predictions. In the present research, a hybrid unsteady RANS-LES parallel multi-block structured grid solver called EAGLEJet is written to perform the nozzle flow calculations. The far-field noise calculation is performed using solutions to the Ffowcs Williams and Hawkings equation. The present calculations use meshes with 5 to 11 million grid points and require about three weeks of computing time with about 100 processors. A baseline single stream convergent nozzle and a dual-stream coaxial convergent nozzle are used for the flow and noise analysis. Calculations for the convergent nozzle are performed at a high subsonic jet Mach number of Mj = 0.9, which is similar to the operating conditions for commercial aircraft engines. A parallel flow gives the flight effect, which is simulated with a co-flow Mach number, Mcf varying from 0.0 to 0.28. The grid resolution effects, statistical properties of the turbulence and the heated jet effects ( TTR = 2.7) are studied and related to the noise characteristics of the jet. Both flow and noise predictions show good agreement with PIV and microphone measurements. The potential core lengths and nozzle wall boundary characteristics are studied to understand the differences between the numerical potential core lengths as compared to experiments. The flight velocity exponent, m is calculated from the noise reduction in overall sound pressure levels (OASPL, dB) and relative velocity (V j -- Vcf) at all jet inlet (angular) angles. The variation of the exponent, m at lower (50° to 90°) and higher aft inlet angles (120° to 150°) is studied and compared with available measurements. Previous studies have shown a different variation of the exponent with inlet angles while the current numerical data match well with recent experiments conducted on the same nozzle geometry. Today, turbofans are the most efficient engines in service used in almost all major commercial aircraft. Turbofans have a dual-stream exhaust nozzle with primary and secondary flow whose flow and noise characteristics are different from that of single stream jets. A Boeing-designed coaxial nozzle, with area ratio of As/Ap = 3.0, is used to study dual-stream jet noise in the present research. In this configuration, the primary nozzle extends beyond the secondary nozzle, which is representative of large turbofan engines in commercial service. The flow calculations are performed at high subsonic Mach numbers in the primary and secondary nozzles (Mpj = 0.85, Msj = 0.95) with heated core flow, TTRp = 2.26 and unheated fan flow, TTRs = 1.0. The co-flow of Mcf = 0.2 is used. The subscript p, s and amb represent the primary (core) nozzle, the secondary (fan) nozzle, and the ambient flow conditions, respectively. The statistical properties in the primary and secondary shear layers are studied and compared with those of the single stream jets. It has been found that the eddy convection velocity is lower in dual-stream jets as compared to the single stream jet operating at a similar jet exit Mach number. The phase velocity is higher in the secondary shear layer as compared to primary shear layer. The noise measurements agree well with the predicted data and noise reduction is observed in the presence of co-flow. The variation of the flight velocity exponent is calculated as a function of nozzle inlet angle. The value of the exponent at higher inlet angles is lower as compared to the single stream jets. This suggests that the noise levels are less affected in the peak noise direction in the presence of co-flow in dual-stream jets as compared to single stream jets. Two reference velocities: primary jet exit velocity Vpj and mixed velocity Vmix are considered which result in different absolute values of the exponents. Scaling of the jet spectra is performed at different inlet angles and good collapse has been obtained between the spectra. The installation effects on jet noise are studied using a simplified pylon structure with a dual-stream nozzle. In the presence of a pylon, the azimuthal symmetry of the nozzle is lost and thus the flow characteristics are different as compared to the baseline nozzle. This will result in different noise characteristics of the installed jet.

Experimental Investigation of Shock-Cell Noise Reduction for Single Stream Nozzles in Simulated Flight

NASA Technical Reports Server (NTRS)

Yamamoto, K.; Brausch, J. F.; Balsa, T. F.; Janardan, B. A.; Knott, P. R.

1984-01-01

Seven single stream model nozzles were tested in the Anechoic Free-Jet Acoustic Test Facility to evaluate the effectiveness of convergent divergent (C-D) flowpaths in the reduction of shock-cell noise under both static and mulated flight conditions. The test nozzles included a baseline convergent circular nozzle, a C-D circular nozzle, a convergent annular plug nozzle, a C-D annular plug nozzle, a convergent multi-element suppressor plug nozzle, and a C-D multi-element suppressor plug nozzle. Diagnostic flow visualization with a shadowgraph and aerodynamic plume measurements with a laser velocimeter were performed with the test nozzles. A theory of shock-cell noise for annular plug nozzles with shock-cells in the vicinity of the plug was developed. The benefit of these C-D nozzles was observed over a broad range of pressure ratiosin the vicinity of their design conditions. At the C-D design condition, the C-D annual nozzle was found to be free of shock-cells on the plug.

Acoustic and aerodynamic performance of a variable-pitch 1.83-meter-(6-ft) diameter 1.20-pressure-ratio fan stage (QF-9)

NASA Technical Reports Server (NTRS)

Glaser, F. W.; Woodward, R. P.; Lucas, J. G.

1977-01-01

Far field noise data and related aerodynamic performance are presented for a variable pitch fan stage having characteristics suitable for low noise, STOL engine application. However, no acoustic suppression material was used in the flow passages. The fan was externally driven by an electric motor. Tests were made at several forward thrust rotor blade pitch angles and one for reverse thrust. Fan speed was varied from 60 to 120 percent of takeoff (design) speed, and exhaust nozzles having areas 92 to 105 percent of design were tested. The fan noise level was at a minimum at the design rotor blade pitch angles of 64 deg for takeoff thrust and at 57 deg for approach (50 percent takeoff thrust). Perceived noise along a 152.4-m sideline reached 100.1 PNdb for the takeoff (design) configuration for a stage pressure ratio of 1.17 and thrust of 57,600 N. For reverse thrust the PNL values were 4 to 5 PNdb above the takeoff values at comparable fan speeds.

Identification of critical process variables affecting particle size following precipitation using a supercritical fluid.

PubMed

Sacha, Gregory A; Schmitt, William J; Nail, Steven L

2006-01-01

The critical processing parameters affecting average particle size, particle size distribution, yield, and level of residual carrier solvent using the supercritical anti-solvent method (SAS) were identified. Carbon dioxide was used as the supercritical fluid. Methylprednisolone acetate was used as the model solute in tetrahydrofuran. Parameters examined included pressure of the supercritical fluid, agitation rate, feed solution flow rate, impeller diameter, and nozzle design. Pressure was identified as the most important process parameter affecting average particle size, either through the effect of pressure on dispersion of the feed solution into the precipitation vessel or through the effect of pressure on solubility of drug in the CO2/organic solvent mixture. Agitation rate, impeller diameter, feed solution flow rate, and nozzle design had significant effects on particle size, which suggests that dispersion of the feed solution is important. Crimped HPLC tubing was the most effective method of introducing feed solution into the precipitation vessel, largely because it resulted in the least amount of clogging during the precipitation. Yields of 82% or greater were consistently produced and were not affected by the processing variables. Similarly, the level of residual solvent was independent of the processing variables and was present at 0.0002% wt/wt THF or less.

The Effect of Variable Gravity on the Cooling Performance of a 16-Nozzle Spray Array

DTIC Science & Technology

2008-09-01

Coolants and Capabilities [1]. spray cooling, have been studied in depth and are currently being used on various types of systems. Mudawar [1] presented many...meniscus. Surface modifications could be made to reduce this effect while enhancing overall operability. 58 Bibliography [1] Mudawar , I., “Assessment of

30 CFR 18.65 - Flame test of hose.

Code of Federal Regulations, 2010 CFR

2010-07-01

... MINING PRODUCTS ELECTRIC MOTOR-DRIVEN MINE EQUIPMENT AND ACCESSORIES Inspections and Tests § 18.65 Flame... variable-speed electric fan and an ASME flow nozzle (16-81/2 inches reduction) to attain constant air velocities at any speed between 50-500 feet a minute. (4) An electric timer or stopwatch to measure the...

A computational model for the prediction of jet entrainment in the vicinity of nozzle boattails (The BOAT code)

NASA Technical Reports Server (NTRS)

Dash, S. M.; Pergament, H. S.

1978-01-01

The basic code structure is discussed, including the overall program flow and a brief description of all subroutines. Instructions on the preparation of input data, definitions of key FORTRAN variables, sample input and output, and a complete listing of the code are presented.

Characteristics and performance of the variable polarity plasma arc welding process used in the Space Shuttle external tank

NASA Technical Reports Server (NTRS)

Hung, R. J.; Lee, C. C.; Liu, J. W.

1990-01-01

Significant advantages of the Variable Polarity Plasma Arc (VPPA) Welding Process include faster welding, fewer repairs, less joint preparation, reduced weldment distortion, and absence of porosity. Flow profiles and power distribution of argon plasma gas as a working fluid to produce plasma arc jet in the VPPA welding process was analyzed. Major loss of heat transfer for flow through the nozzle is convective heat transfer; for the plasma jet flow between the outlet of the nozzle and workpiece is radiative heat transfer; and for the flow through the keyhole of the workpiece is convective heat transfer. The majority of the power absorbed by the keyhole of the workpiece is used for melting the solid metal workpiece into a molten metallic puddle. The crown and root widths and the crown and root heights can be predicted. An algorithm for promoting automatic control of flow parameters and the dimensions of the final product of the welding specification to be used for the VPPA Welding System operated at MSFC are provided.

Dual-Mode Combustor

NASA Technical Reports Server (NTRS)

Trefny, Charles J (Inventor); Dippold, Vance F (Inventor)

2013-01-01

A new dual-mode ramjet combustor used for operation over a wide flight Mach number range is described. Subsonic combustion mode is usable to lower flight Mach numbers than current dual-mode scramjets. High speed mode is characterized by supersonic combustion in a free-jet that traverses the subsonic combustion chamber to a variable nozzle throat. Although a variable combustor exit aperture is required, the need for fuel staging to accommodate the combustion process is eliminated. Local heating from shock-boundary-layer interactions on combustor walls is also eliminated.

A detailed description of the uncertainty analysis for high area ratio rocket nozzle tests at the NASA Lewis Research Center

NASA Technical Reports Server (NTRS)

Davidian, Kenneth J.; Dieck, Ronald H.; Chuang, Isaac

1987-01-01

A preliminary uncertainty analysis was performed for the High Area Ratio Rocket Nozzle test program which took place at the altitude test capsule of the Rocket Engine Test Facility at the NASA Lewis Research Center. Results from the study establish the uncertainty of measured and calculated parameters required for the calculation of rocket engine specific impulse. A generalized description of the uncertainty methodology used is provided. Specific equations and a detailed description of the analysis is presented. Verification of the uncertainty analysis model was performed by comparison with results from the experimental program's data reduction code. Final results include an uncertainty for specific impulse of 1.30 percent. The largest contributors to this uncertainty were calibration errors from the test capsule pressure and thrust measurement devices.

A detailed description of the uncertainty analysis for High Area Ratio Rocket Nozzle tests at the NASA Lewis Research Center

NASA Technical Reports Server (NTRS)

Davidian, Kenneth J.; Dieck, Ronald H.; Chuang, Isaac

1987-01-01

A preliminary uncertainty analysis has been performed for the High Area Ratio Rocket Nozzle test program which took place at the altitude test capsule of the Rocket Engine Test Facility at the NASA Lewis Research Center. Results from the study establish the uncertainty of measured and calculated parameters required for the calculation of rocket engine specific impulse. A generalized description of the uncertainty methodology used is provided. Specific equations and a detailed description of the analysis are presented. Verification of the uncertainty analysis model was performed by comparison with results from the experimental program's data reduction code. Final results include an uncertainty for specific impulse of 1.30 percent. The largest contributors to this uncertainty were calibration errors from the test capsule pressure and thrust measurement devices.

The effect of using a cement gun with a narrow nozzle on cement penetration for total elbow arthroplasty: a cadaveric study.

PubMed

Smith, Geoffrey C S; McCann, Phillip S; Simpson, Danielle; Blewitt, Neil; Amirfeyz, Rouin

2015-02-01

To compare the cement mantle characteristics associated with use of a narrow nozzle cement gun versus the use of a 60-mL catheter tip syringe. Twelve cadaveric distal humeri were cemented with either a cement gun or a syringe without canal occlusion. The humeri were sectioned and photographed. The corticocancellous junction and the outer margin of the cement mantle were analyzed digitally. The corticocancellous junction defined the available area for cement penetration. The outline of the cement mantle defined the actual area of penetration. The ratio of penetration to the available area was recorded for each slice. The mean ratio for each humerus was multiplied by the number of slices in that sample containing cement to calculate a cement index. The cement penetration ratios observed in cross-sections at the same level and the cement index were significantly greater with the use of the cement gun than with the use of the syringe. There was no difference in the number of slices that contained cement. The use of a cement gun with a narrow nozzle improved cement mantle characteristics compared with the use of a syringe when measured in a cadaveric model in the absence of canal occlusion. Improving cement mantle characteristics may decrease the incidence of aseptic loosening after total elbow arythroplasty. Copyright © 2015 American Society for Surgery of the Hand. Published by Elsevier Inc. All rights reserved.

Gas only nozzle

DOEpatents

Bechtel, William Theodore; Fitts, David Orus; DeLeonardo, Guy Wayne

2002-01-01

A diffusion flame nozzle gas tip is provided to convert a dual fuel nozzle to a gas only nozzle. The nozzle tip diverts compressor discharge air from the passage feeding the diffusion nozzle air swirl vanes to a region vacated by removal of the dual fuel components, so that the diverted compressor discharge air can flow to and through effusion holes in the end cap plate of the nozzle tip. In a preferred embodiment, the nozzle gas tip defines a cavity for receiving the compressor discharge air from a peripheral passage of the nozzle for flow through the effusion openings defined in the end cap plate.

Gas only nozzle fuel tip

DOEpatents

Bechtel, William Theodore; Fitts, David Orus; DeLeonardo, Guy Wayne

2002-01-01

A diffusion flame nozzle gas tip is provided to convert a dual fuel nozzle to a gas only nozzle. The nozzle tip diverts compressor discharge air from the passage feeding the diffusion nozzle air swirl vanes to a region vacated by removal of the dual fuel components, so that the diverted compressor discharge air can flow to and through effusion holes in the end cap plate of the nozzle tip. In a preferred embodiment, the nozzle gas tip defines a cavity for receiving the compressor discharge air from a peripheral passage of the nozzle for flow through the effusion openings defined in the end cap plate.

Fundamental Aeronautics Program: Overview of Project Work in Supersonic Cruise Efficiency

NASA Technical Reports Server (NTRS)

Castner, Raymond

2011-01-01

The Supersonics Project, part of NASA?s Fundamental Aeronautics Program, contains a number of technical challenge areas which include sonic boom community response, airport noise, high altitude emissions, cruise efficiency, light weight durable engines/airframes, and integrated multi-discipline system design. This presentation provides an overview of the current (2011) activities in the supersonic cruise efficiency technical challenge, and is focused specifically on propulsion technologies. The intent is to develop and validate high-performance supersonic inlet and nozzle technologies. Additional work is planned for design and analysis tools for highly-integrated low-noise, low-boom applications. If successful, the payoffs include improved technologies and tools for optimized propulsion systems, propulsion technologies for a minimized sonic boom signature, and a balanced approach to meeting efficiency and community noise goals. In this propulsion area, the work is divided into advanced supersonic inlet concepts, advanced supersonic nozzle concepts, low fidelity computational tool development, high fidelity computational tools, and improved sensors and measurement capability. The current work in each area is summarized.

Combustor with multistage internal vortices

DOEpatents

Shang, Jer Yu; Harrington, R.E.

1987-05-01

A fluidized bed combustor is provided with a multistage arrangement of vortex generators in the freeboard area. The vortex generators are provided by nozzle means which extend into the interior of the freeboard for forming vortices within the freeboard areas to enhance the combustion of particulate material entrained in product gases ascending into the freeboard from the fluidized bed. Each of the nozzles are radially inwardly spaced from the combustor walls defining the freeboard to provide for the formation of an essentially vortex-free, vertically extending annulus about the vortices whereby the particulate material centrifuged from the vortices against the inner walls of the combustor is returned through the annulus to the fluidized bed. By adjusting the vortex pattern within the freeboard, a significant portion of the full cross-sectional area of the freeboard except for the peripheral annulus can be contacted with the turbulent vortical flow for removing the particulate material from the gaseous products and also for enhancing the combustion thereof within the freeboard. 2 figs.

Combustor with multistage internal vortices

DOEpatents

Shang, Jer Y.; Harrington, Richard E.

1989-01-01

A fluidized bed combustor is provided with a multistage arrangement of vortex generators in the freeboard area. The vortex generators are provided by nozzle means which extend into the interior of the freeboard for forming vortices within the freeboard area to enhance the combustion of particulate material entrained in product gases ascending into the freeboard from the fluidized bed. Each of the nozzles are radially inwardly spaced from the combustor walls defining the freeboard to provide for the formation of an essentially vortex-free, vertically extending annulus about the vortices whereby the particulate material centrifuged from the vortices against the inner walls of the combustor is returned through the annulus to the fluidized bed. By adjusting the vortex pattern within the freeboard, a significant portion of the full cross-sectional area of the freeboard except for the peripheral annulus can be contacted with the turbulent vortical flow for removing the particulate material from the gaseous products and also for enhancing the combustion thereof within the freeboard.

Fundamental Aeronautics Program: Overview of Propulsion Work in the Supersonic Cruise Efficiency Technical Challenge

NASA Technical Reports Server (NTRS)

Castner, Ray

2012-01-01

The Supersonics Project, part of NASA's Fundamental Aeronautics Program, contains a number of technical challenge areas which include sonic boom community response, airport noise, high altitude emissions, cruise efficiency, light weight durable engines/airframes, and integrated multi-discipline system design. This presentation provides an overview of the current (2012) activities in the supersonic cruise efficiency technical challenge, and is focused specifically on propulsion technologies. The intent is to develop and validate high-performance supersonic inlet and nozzle technologies. Additional work is planned for design and analysis tools for highly-integrated low-noise, low-boom applications. If successful, the payoffs include improved technologies and tools for optimized propulsion systems, propulsion technologies for a minimized sonic boom signature, and a balanced approach to meeting efficiency and community noise goals. In this propulsion area, the work is divided into advanced supersonic inlet concepts, advanced supersonic nozzle concepts, low fidelity computational tool development, high fidelity computational tools, and improved sensors and measurement capability. The current work in each area is summarized.

PARC Analysis of the NASA/GE 2D NRA Mixer/Ejector Nozzle

NASA Technical Reports Server (NTRS)

DeBonis, J. R.

1999-01-01

Interest in developing a new generation supersonic transport has increased in the past several years. Current projections indicate this aircraft would cruise at approximately Mach 2.4, have a range of 5000 nautical miles and carry at least 250 passengers. A large market for such an aircraft will exist in the next century due to a predicted doubling of the demand for long range air transportation by the end of the century and the growing influence of the Pacific Rim nations. Such a proposed aircraft could more than halve the flying time from Los Angeles to Tokyo. However, before a new economically feasible supersonic transport can be built, many key technologies must be developed. Among these technologies is noise suppression. Propulsion systems for a supersonic transport using current technology would exceed acceptable noise levels. All new aircraft must satisfy FAR 36 Stage III noise regulations. The largest area of concern is the noise generated during takeoff. A concerted effort under NASA's High Speed Research (HSR) program has begun to address the problem of noise suppression. One of the most promising concepts being studied in the area of noise suppression is the mixer/ejector nozzle. This study analyzes a typical noise suppressing mixer ejector nozzle at take off conditions, using a Full Navier-Stokes (FNS) computational fluid dynamics (CFD) code.

Design of porous microparticles with single-micron size by novel spray freeze-drying technique using four-fluid nozzle.

PubMed

Niwa, Toshiyuki; Shimabara, Hiroko; Kondo, Masahiro; Danjo, Kazumi

2009-12-01

Spray freeze-drying (SFD) process, which is a novel particle design technique previously developed by authors, has been improved by using four-fluid nozzle (4N) instead of conventional two-fluid nozzle (2N) to expand its application in pharmaceutical industry. Aqueous spray solutions of the drug and the polymeric carrier were separately supplied into 4N, and atomized while colliding with each other at the tip of nozzle. The droplets of mixed solutions were directly immersed into liquid nitrogen and immediately frozen to form a suspension. Then, the iced droplets were lyophilized by freeze-dryer to prepare the composite particles of the drug and carrier. This process has been used in the present study to modify and enhance the dissolution profiles of poorly water-soluble drug, phenytoin. Water-soluble and enteric polymeric carriers in pharmaceutical use were used as a dissolution modifier. The SFD composite particles prepared by using 4N were fully characterized compared to those using 2N from morphological and physicochemical perspectives. It was found that the particles have fine porous structure producing vast specific surface area. Further, phenytoin was completely dispersed as amorphous state in the polymeric matrix with higher carrier ratio than phenytoin:carrier = 1:3. The dissolution of phenytoin from the water-soluble carrier-based particles was greatly enhanced because of large effective surface area and disappearance of crystalline. On the other hand, the release profiles from enteric carrier-based particles showed the typical enteric patterns, that is, delayed in acidic medium and accelerated in neutral pH. The results demonstrated that SFD technique using 4N has potential to develop the novel solubilized formulation for poorly water-soluble APIs.

Pulse Detonation Rocket Magnetohydrodynamic Power Experiment

NASA Technical Reports Server (NTRS)

Litchford, R. J.; Jones, J. E.; Dobson, C. C.; Cole, J. W.; Thompson, B. R.; Plemmons, D. H.; Turner, M. W.

2003-01-01

The production of onboard electrical power by pulse detonation engines is problematic in that they generate no shaft power; however, pulse detonation driven magnetohydrodynamic (MHD) power generation represents one intriguing possibility for attaining self-sustained engine operation and generating large quantities of burst power for onboard electrical systems. To examine this possibility further, a simple heat-sink apparatus was developed for experimentally investigating pulse detonation driven MHD generator concepts. The hydrogen oxygen fired driver was a 90 cm long stainless steel tube having a 4.5 cm square internal cross section and a short Schelkin spiral near the head end to promote rapid formation of a detonation wave. The tube was intermittently filled to atmospheric pressure and seeded with a CsOH/methanol prior to ignition by electrical spark. The driver exhausted through an aluminum nozzle having an area contraction ratio of A*/A(sub zeta) = 1/10 and an area expansion ratio of A(sub zeta)/A* = 3.2 (as limited by available magnet bore size). The nozzle exhausted through a 24-electrode segmented Faraday channel (30.5 cm active length), which was inserted into a 0.6 T permanent magnet assembly. Initial experiments verified proper drive operation with and without the nozzle attachment, and head end pressure and time resolved thrust measurements were acquired. The exhaust jet from the nozzle was interrogated using a polychromatic microwave interferometer yielding an electron number density on the order of 10(exp 12)/cm at the generator entrance. In this case, MHD power generation experiments suffered from severe near-electrode voltage drops and low MHD interaction; i.e., low flow velocity, due to an inherent physical constraint on expansion with the available magnet. Increased scaling, improved seeding techniques, higher magnetic fields, and higher expansion ratios are expected to greatly improve performance.

Intelligent Engine Systems: Acoustics

NASA Technical Reports Server (NTRS)

Wojno, John; Martens, Steve; Simpson, Benjamin

2008-01-01

An extensive study of new fan exhaust nozzle technologies was performed. Three new uniform chevron nozzles were designed, based on extensive CFD analysis. Two new azimuthally varying variants were defined. All five were tested, along with two existing nozzles, on a representative model-scale, medium BPR exhaust nozzle. Substantial acoustic benefits were obtained from the uniform chevron nozzle designs, the best benefit being provided by an existing design. However, one of the azimuthally varying nozzle designs exhibited even better performance than any of the uniform chevron nozzles. In addition to the fan chevron nozzles, a new technology was demonstrated, using devices that enhance mixing when applied to an exhaust nozzle. The acoustic benefits from these devices applied to medium BPR nozzles were similar, and in some cases superior to, those obtained from conventional uniform chevron nozzles. However, none of the low noise technologies provided equivalent acoustic benefits on a model-scale high BPR exhaust nozzle, similar to current large commercial applications. New technologies must be identified to improve the acoustics of state-of-the-art high BPR jet engines.

Heat transfer in rocket engine combustion chambers and regeneratively cooled nozzles

NASA Technical Reports Server (NTRS)

1993-01-01

A conjugate heat transfer computational fluid dynamics (CFD) model to describe regenerative cooling in the main combustion chamber and nozzle and in the injector faceplate region for a launch vehicle class liquid rocket engine was developed. An injector model for sprays which treats the fluid as a variable density, single-phase media was formulated, incorporated into a version of the FDNS code, and used to simulate the injector flow typical of that in the Space Shuttle Main Engine (SSME). Various chamber related heat transfer analyses were made to verify the predictive capability of the conjugate heat transfer analysis provided by the FDNS code. The density based version of the FDNS code with the real fluid property models developed was successful in predicting the streamtube combustion of individual injector elements.

Synthetic optimization of air turbine for dental handpieces.

PubMed

Shi, Z Y; Dong, T

2014-01-01

A synthetic optimization of Pelton air turbine in dental handpieces concerning the power output, compressed air consumption and rotation speed in the mean time is implemented by employing a standard design procedure and variable limitation from practical dentistry. The Pareto optimal solution sets acquired by using the Normalized Normal Constraint method are mainly comprised of two piecewise continuous parts. On the Pareto frontier, the supply air stagnation pressure stalls at the lower boundary of the design space, the rotation speed is a constant value within the recommended range from literature, the blade tip clearance insensitive to while the nozzle radius increases with power output and mass flow rate of compressed air to which the residual geometric dimensions are showing an opposite trend within their respective "pieces" compared to the nozzle radius.

Design and control of a variable geometry turbofan with an independently modulated third stream

NASA Astrophysics Data System (ADS)

Simmons, Ronald J.

Emerging 21st century military missions task engines to deliver the fuel efficiency of a high bypass turbofan while retaining the ability to produce the high specific thrust of a low bypass turbofan. This study explores the possibility of satisfying such competing demands by adding a second independently modulated bypass stream to the basic turbofan architecture. This third stream can be used for a variety of purposes including: providing a cool heat sink for dissipating aircraft heat loads, cooling turbine cooling air, and providing a readily available stream of constant pressure ratio air for lift augmentation. Furthermore, by modulating airflow to the second and third streams, it is possible to continuously match the engine's airflow demand to the inlet's airflow supply thereby reducing spillage and increasing propulsive efficiency. This research begins with a historical perspective of variable cycle engines and shows a logical progression to proposed architectures. Then a novel method for investigating optimal performance is presented which determines most favorable on design variable geometry settings, most beneficial moment to terminate flow holding, and an optimal scheduling of variable features for fuel efficient off design operation. Mission analysis conducted across the three candidate missions verifies that these three stream variable cycles can deliver fuel savings in excess of 30% relative to a year 2000 reference turbofan. This research concludes by evaluating the relative impact of each variable technology on the performance of adaptive engine architectures. The most promising technologies include modulated turbine cooling air, variable high pressure turbine inlet area and variable third stream nozzle throat area. With just these few features it is possible to obtain nearly optimal performance, including 90% or more of the potential fuel savings, with far fewer variable features than are available in the study engine. It is abundantly clear that three stream variable architectures can significantly outperform existing two stream turbofans in both fuel efficiency and at the vehicle system level with only a modest increase in complexity and weight. Such engine architectures should be strongly considered for future military applications.

Nozzle insert for mixed mode fuel injector

DOEpatents

Lawrence, Keith E [Peoria, IL

2006-11-21

A fuel injector includes a homogenous charge nozzle outlet set and a conventional nozzle outlet set controlled respectively, by first and second needle valve members. The homogeneous charged nozzle outlet set is defined by a nozzle insert that is attached to an injector body, which defines the conventional nozzle outlet set. The nozzle insert is a one piece metallic component with a large diameter segment separated from a small diameter segment by an annular engagement surface. One of the needle valve members is guided on an outer surface of the nozzle insert, and the nozzle insert has an interference fit attachment to the injector body.

A performance comparison of two small rocket nozzles

NASA Technical Reports Server (NTRS)

Arrington, Lynn A.; Reed, Brian D.; Rivera, Angel, Jr.

1996-01-01

An experimental study was conducted on two small rockets (110 N thrust class) to directly compare a standard conical nozzle with a bell nozzle optimized for maximum thrust using the Rao method. In large rockets, with throat Reynolds numbers of greater than 1 x 10(exp 5), bell nozzles outperform conical nozzles. In rockets with throat Reynolds numbers below 1 x 10(exp 5), however, test results have been ambiguous. An experimental program was conducted to test two small nozzles at two different fuel film cooling percentages and three different chamber pressures. Test results showed that for the throat Reynolds number range from 2 x 10(exp 4) to 4 x 10(exp 4), the bell nozzle outperformed the conical nozzle. Thrust coefficients for the bell nozzle were approximately 4 to 12 percent higher than those obtained with the conical nozzle. As expected, testing showed that lowering the fuel film cooling increased performance for both nozzle types.

Parametric investigation of single-expansion-ramp nozzles at Mach numbers from 0.60 to 1.20

NASA Technical Reports Server (NTRS)

Capone, Francis J.; Re, Richard J.; Bare, E. Ann

1992-01-01

An investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine the effects of varying six nozzle geometric parameters on the internal and aeropropulsive performance characteristics of single-expansion-ramp nozzles. This investigation was conducted at Mach numbers from 0.60 to 1.20, nozzle pressure ratios from 1.5 to 12, and angles of attack of 0 deg +/- 6 deg. Maximum aeropropulsive performance at a particular Mach number was highly dependent on the operating nozzle pressure ratio. For example, as the nozzle upper ramp length or angle increased, some nozzles had higher performance at a Mach number of 0.90 because of the nozzle design pressure was the same as the operating pressure ratio. Thus, selection of the various nozzle geometric parameters should be based on the mission requirements of the aircraft. A combination of large upper ramp and large lower flap boattail angles produced greater nozzle drag coefficients at Mach number greater than 0.80, primarily from shock-induced separation on the lower flap of the nozzle. A static conditions, the convergent nozzle had high and nearly constant values of resultant thrust ratio over the entire range of nozzle pressure ratios tested. However, these nozzles had much lower aeropropulsive performance than the convergent-divergent nozzle at Mach number greater than 0.60.

29. SATURN ROCKET ENGINE LOCATED ON NORTH SIDE OF STATIC ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

29. SATURN ROCKET ENGINE LOCATED ON NORTH SIDE OF STATIC TEST STAND - DETAILS OF THE EXPANSION NOZZLE. - Marshall Space Flight Center, Saturn Propulsion & Structural Test Facility, East Test Area, Huntsville, Madison County, AL

Introduction of a theoretical splashing degree to assess the performance of low-viscosity oils in filling of capsules.

PubMed

Niederquell, Andreas; Kuentz, Martin

2011-03-01

These days an alternative to soft capsules is liquid-filled hard capsules. Their filling technology was investigated earlier with highly viscous formulations, while hardly any academic research focused on low-viscosity systems. Accordingly, this work addressed the filling of such oils that are splashing during the dosing process. It was aimed to first study capsule filling, using middle-chain triglycerides as reference oil, in order to then evaluate the concept of a new theoretical splashing degree for different oils. A laboratory-scale filling machine was used that included capsule sealing. Thus, the liquid encapsulation by microspray technology was employed to seal the dosage form. As a result of the study with reference oil, the filling volume and the temperature were found to be significant for the rate of leaking capsules. The filling volume was also important for weight variability of the capsules. However, most critical for this variability was the diameter of the filling nozzle. We proposed a power law for the coefficient of weight variability as a function of the nozzle diameter and the obtained exponent agreed with the proposed theory. Subsequently, a comparison of different oils revealed that the relative splashing degree shared a correlation with the coefficient of the capsule weight variability (Pearson product moment correlation of r=0.990). The novel theoretical concept was therefore found to be predictive for weight variability of the filled capsules. Finally, guidance was provided for the process development of liquid-filled capsules using low-viscosity oils. © 2011 American Association of Pharmaceutical Scientists

Flow Separation Side Loads Excitation of Rocket Nozzle FEM

NASA Technical Reports Server (NTRS)

Smalley, Kurt B.; Brown, Andrew; Ruf, Joseph; Gilbert, John

2007-01-01

Modern rocket nozzles are designed to operate over a wide range of altitudes, and are also built with large aspect ratios to enable high efficiencies. Nozzles designed to operate over specific regions of a trajectory are being replaced in modern launch vehicles by those that are designed to operate from earth to orbit. This is happening in parallel with modern manufacturing and wall cooling techniques allowing for larger aspect ratio nozzles to be produced. Such nozzles, though operating over a large range of altitudes and ambient pressures, are typically designed for one specific altitude. Above that altitude the nozzle flow is 'underexpanded' and below that altitude, the nozzle flow is 'overexpanded'. In both conditions the nozzle produces less than the maximum possible thrust at that altitude. Usually the nozzle design altitude is well above sea level, leaving the nozzle flow in an overexpanded state for its start up as well as for its ground testing where, if it is a reusable nozzle such as the Space Shuttle Main Engine (SSME), the nozzle will operate for the majority of its life. Overexpansion in a rocket nozzle presents the critical, and sometimes design driving, problem of flow separation induced side loads. To increase their understanding of nozzle side loads, engineers at MSFC began an investigation in 2000 into the phenomenon through a task entitled "Characterization and Accurate Modeling of Rocket Engine Nozzle Side Loads", led by A. Brown. The stated objective of this study was to develop a methodology to accurately predict the character and magnitude of nozzle side loads. The study included further hot-fire testing of the MC-l engine, cold flow testing of subscale nozzles, CFD analyses of both hot-fire and cold flow nozzle testing, and finite element (fe.) analysis of the MC-1 engine and cold flow tested nozzles. A follow on task included an effort to formulate a simplified methodology for modeling a side load during a two nodal diameter fluid/structure interaction for a single moment in time.

Support pedestals for interconnecting a cover and nozzle band wall in a gas turbine nozzle segment

DOEpatents

Yu, Yufeng Phillip; Itzel, Gary Michael; Webbon, Waylon Willard; Bagepalli, Radhakrishna; Burdgick, Steven Sebastian; Kellock, Iain Robertson

2002-01-01

A gas turbine nozzle segment has outer and inner band portions. Each band portion includes a nozzle wall, a cover and an impingement plate between the cover and nozzle wall defining two cavities on opposite sides of the impingement plate. Cooling steam is supplied to one cavity for flow through the apertures of the impingement plate to cool the nozzle wall. Structural pedestals interconnect the cover and nozzle wall and pass through holes in the impingement plate to reduce localized stress otherwise resulting from a difference in pressure within the chamber of the nozzle segment and the hot gas path and the fixed turbine casing surrounding the nozzle stage. The pedestals may be cast or welded to the cover and nozzle wall.

Filling of High-Concentration Monoclonal Antibody Formulations into Pre-filled Syringes: Investigating Formulation-Nozzle Interactions To Minimize Nozzle Clogging.

PubMed

Shieu, Wendy; Stauch, Oliver B; Maa, Yuh-Fun

2015-01-01

Syringe filling of high-concentration/viscosity monoclonal antibody formulations is a complex process that is not fully understood. This study, which builds on a previous investigation that used a bench-top syringe filling unit to examine formulation drying at the filling nozzle tip and subsequent nozzle clogging, further explores the impact of formulation-nozzle material interactions on formulation drying and nozzle clogging. Syringe-filling nozzles made of glass, stainless steel, or plastic (polypropylene, silicone, and Teflon®), which represent a full range of materials with hydrophilic and hydrophobic properties as quantified by contact angle measurements, were used to fill liquids of different viscosity, including a high-concentration monoclonal antibody formulation. Compared with hydrophilic nozzles, hydrophobic nozzles offered two unique features that discouraged formulation drying and nozzle clogging: (1) the liquid formulation is more likely to be withdrawn into the hydrophobic nozzle under the same suck-back conditions, and (2) the residual liquid film left on the nozzle wall when using high suck-back settings settles to form a liquid plug away from the hydrophobic nozzle tip. Making the tip of the nozzle hydrophobic (silicone-coating on glass and Teflon-coating stainless steel) could achieve the same suck-back performance as plastic nozzles. This study demonstrated that using hydrophobic nozzles are most effective in reducing the risk of nozzle clogging by drying of high-concentration monoclonal antibody formulation during extended nozzle idle time in a large-scale filling facility and environment. Syringe filling is a well-established manufacturing process and has been implemented by numerous contract manufacturing organizations and biopharmaceutical companies. However, its technical details and associated critical process parameters are rarely published. Information on high-concentration/viscosity formulation filling is particularly lacking. This study is the continuation of a previous investigation with a focus on understanding the impact of nozzle material on the suck-back function of liquid formulations. The findings identified the most critical parameter-nozzle material hydrophobicity-in alleviating formulation drying at the nozzle tip and eventually limiting the occurrence of nozzle clogging during the filling process. The outcomes of this study will benefit scientists and engineers who develop pre-filled syringe products by providing a better understanding of high-concentration formulation filling principles and challenges. © PDA, Inc. 2015.

Effect of the Thruster Configurations on a Laser Ignition Microthruster

NASA Astrophysics Data System (ADS)

Koizumi, Hiroyuki; Hamasaki, Kyoichi; Kondo, Ryo; Okada, Keisuke; Nakano, Masakatsu; Arakawa, Yoshihiro

Research and development of small spacecraft have advanced extensively throughout the world and propulsion devices suitable for the small spacecraft, microthruster, is eagerly anticipated. The authors proposed a microthruster using 1—10-mm-size solid propellant. Small pellets of solid propellant are installed in small combustion chambers and ignited by the irradiation of diode laser beam. This thruster is referred as to a laser ignition microthruster. Solid propellant enables large thrust capability and compact propulsion system. To date theories of a solid-propellant rocket have been well established. However, those theories are for a large-size solid propellant and there are a few theories and experiments for a micro-solid rocket of 1—10mm class. This causes the difficulty of the optimum design of a micro-solid rocket. In this study, we have experimentally investigated the effect of thruster configurations on a laser ignition microthruster. The examined parameters are aperture ratio of the nozzle, length of the combustion chamber, area of the nozzle throat, and divergence angle of the nozzle. Specific impulse dependences on those parameters were evaluated. It was found that large fraction of the uncombusted propellant was the main cause of the degrading performance. Decreasing the orifice diameter in the nozzle with a constant open aperture ratio was an effective method to improve this degradation.

Development of Flow and Heat Transfer Models for the Carbon Fiber Rope in Nozzle Joints of the Space Shuttle Reusable Solid Rocket Motor

NASA Technical Reports Server (NTRS)

Wang, Q.; Ewing, M. E.; Mathias, E. C.; Heman, J.; Smith, C.; McCool, Alex (Technical Monitor)

2001-01-01

Methodologies have been developed for modeling both gas dynamics and heat transfer inside the carbon fiber rope (CFR) for applications in the space shuttle reusable solid rocket motor joints. Specifically, the CFR is modeled using an equivalent rectangular duct with a cross-section area, friction factor and heat transfer coefficient such that this duct has the same amount of mass flow rate, pressure drop, and heat transfer rate as the CFR. An equation for the friction factor is derived based on the Darcy-Forschheimer law and the heat transfer coefficient is obtained from pipe flow correlations. The pressure, temperature and velocity of the gas inside the CFR are calculated using the one-dimensional Navier-Stokes equations. Various subscale tests, both cold flow and hot flow, have been carried out to validate and refine this CFR model. In particular, the following three types of testing were used: (1) cold flow in a RSRM nozzle-to-case joint geometry, (2) cold flow in a RSRM nozzle joint No. 2 geometry, and (3) hot flow in a RSRM nozzle joint environment simulator. The predicted pressure and temperature history are compared with experimental measurements. The effects of various input parameters for the model are discussed in detail.

Two-Dimensional Supersonic Nozzle Thrust Vectoring Using Staggered Ramps

NASA Astrophysics Data System (ADS)

Montes, Carlos Fernando

A novel mechanism for vectoring the thrust of a supersonic, air-breathing engine was analyzed numerically using ANSYS Fluent. The mechanism uses two asymmetrically staggered ramps; one placed at the throat, the other positioned at the exit lip of the nozzle. The nozzle was designed using published flow data, isentropic relationships, and piecewise quartic splines. The design was verified numerically and was in fair agreement with the analytical data. Using the steady-state pressure-based solver, along with the realizable kappa - epsilon turbulence model, a total of eighteen simulations were conducted: three ramp lengths at three angles, using two sets of inlet boundary conditions (non-afterburning and afterburning). The vectoring simulations showed that the afterburning flow yields a lower flow deflection distribution, shown by the calculated average deflection angle and area-weighted integrals of the distributions. The data implies that an aircraft can achieve an average thrust vectoring angle of approximately 30° in a given direction with the longest ramp length and largest ramp angle configuration. With increasing ramp angle, the static pressure across the nozzle inlet increased, causing concern for potential negative effects on the engine's turbine. The mechanism, for which a provisional patent application has been filed, will require further work to investigate the maximum possible thrust vectoring angle, including experiments.

A novel method for the production of core-shell microparticles by inverse gelation optimized with artificial intelligent tools.

PubMed

Rodríguez-Dorado, Rosalia; Landín, Mariana; Altai, Ayça; Russo, Paola; Aquino, Rita P; Del Gaudio, Pasquale

2018-03-01

Numerous studies have been focused on hydrophobic compounds encapsulation as oils. In fact, oils can provide numerous health benefits as synergic ingredient combined with other hydrophobic active ingredients. However, stable microparticles for pharmaceutical purposes are difficult to achieve when commonly techniques are used. In this work, sunflower oil was encapsulated in calcium-alginate capsules by prilling technique in co-axial configuration. Core-shell beads were produced by inverse gelation directly at the nozzle using a w/o emulsion containing aqueous calcium chloride solution in sunflower oil pumped through the inner nozzle while an aqueous alginate solution, coming out from the annular nozzle, produced the beads shell. To optimize process parameters artificial intelligence tools were proposed to optimize the numerous prilling process variables. Homogeneous and spherical microcapsules with narrow size distribution and a thin alginate shell were obtained when the parameters as w/o constituents, polymer concentrations, flow rates and frequency of vibration were optimized by two commercial software, FormRules® and INForm®, which implement neurofuzzy logic and Artificial Neural Networks together with genetic algorithms, respectively. This technique constitutes an innovative approach for hydrophobic compounds microencapsulation. Copyright © 2018 Elsevier B.V. All rights reserved.

Simulation of Liquid Injection Thrust Vector Control for Mars Ascent Vehicle

NASA Technical Reports Server (NTRS)

Gudenkauf, Jared

2017-01-01

The Jet Propulsion Laboratory is currently in the initial design phase for a potential Mars Ascent Vehicle; which will be landed on Mars, stay on the surface for period of time, collect samples from the Mars 2020 rover, and then lift these samples into orbit around Mars. The engineers at JPL have down selected to a hybrid wax-based fuel rocket using a liquid oxidizer based on nitrogen tetroxide, or a Mixed Oxide of Nitrogen. To lower the gross lift-off mass of the vehicle the thrust vector control system will use liquid injection of the oxidizer to deflect the thrust of the main nozzle instead of using a gimbaled nozzle. The disadvantage of going with the liquid injection system is the low technology readiness level with a hybrid rocket. Presented in this paper is an effort to simulate the Mars Ascent Vehicle hybrid rocket nozzle and liquid injection thrust vector control system using the computational fluid dynamic flow solver Loci/Chem. This effort also includes determining the sensitivity of the thrust vector control system to a number of different design variables for the injection ports; including axial location, number of adjacent ports, injection angle, and distance between the ports.

Experimental Characterization of Plasma Detachment from Magnetic Nozzles

NASA Astrophysics Data System (ADS)

Olsen, Christopher Scott

Magnetic nozzles, like Laval nozzles, are observed in several natural systems and have application in areas such as electric propulsion and plasma processing. Plasma flowing through these nozzles is inherently tied to the field lines and must separate for momentum redirection or particle transport to occur. Plasma detachment and associated mechanisms from a magnetic nozzle are investigated. Experimental results are presented from the plume of the VASIMRRTM VX-200 device flowing along an axisymmetric magnetic nozzle and operated at two ion energies to explore momentum dependent detachment. The argon plume expanded into a 150m3 vacuum chamber where the background pressure was low enough that charge-exchange mean-free-paths were longer than experiment scale lengths. This magnetic nozzle system is demonstrated to hydrodynamically scale up to astrophysical plasmas, particularly the solar chromosphere, implying general relevance to many systems. Plasma parameters were mapped over a large spatial range using measurements from multiple plasma diagnostics. The data show that the plume does not follow the magnetic field lines. A mapped integration of the ion flux shows the plume may be divided into three regions where 1) the plume briefly follows the magnetic flux, 2) diverges quadratically before 3) expanding with linear trajectories. Transitioning from region 1→2, the ion flux departs from the magnetic flux suggesting ion detachment. An instability forms in region 2 driving an oscillating electric field that causes ions to expand before enhancing electron cross-field transport through anomalous resistivity. Transitioning from region 2→3 the electric field dissipates, the trajectories linearize, and the plume effectively detaches. A delineation of sub-to-super Alfvenic flow aligns well with the inflection points of the linearization without a change in magnetic topology. The detachment process is best described as a two part process: First, ions detach by a breakdown of the magnetic moment when the quantity |v/fcLB| becomes of order unity. Second, the turbulent electric field enhances electron transport up to a factor of 4+/-1 above collisional diffusion; electron cross-field velocities approximate that of the ions and depart on more centralized field lines. Electrons are believed to detach by breakdown of magnetic moment further downstream in the weaker magnetic field.

Numerical Modeling of a Vortex Stabilized Arcjet. Ph.D. Thesis, 1991 Final Report

NASA Technical Reports Server (NTRS)

Pawlas, Gary E.

1992-01-01

Arcjet thrusters are being actively considered for use in Earth orbit maneuvering applications. Experimental studies are currently the chief means of determining an optimal thruster configuration. Earlier numerical studies have failed to include all of the effects found in typical arcjets including complex geometries, viscosity, and swirling flow. Arcjet geometries are large area ratio converging nozzles with centerbodies in the subsonic portion of the nozzle. The nozzle walls serve as the anode while the centerbody functions as the cathode. Viscous effects are important because the Reynolds number, based on the throat radius, is typically less than 1,000. Experimental studies have shown that a swirl or circumferential velocity component stabilizes a constricted arc. This dissertation describes the equations governing flow through a constricted arcjet thruster. An assumption that the flowfield is in local thermodynamic equilibrium leads to a single fluid plasma temperature model. An order of magnitude analysis reveals the governing fluid mechanics equations are uncoupled from the electromagnetic field equations. A numerical method is developed to solve the governing fluid mechanics equations, the Thin Layer Navier-Stokes equations. A coordinate transformation is employed in deriving the governing equations to simplify the application of boundary conditions in complex geometries. An axisymmetric formulation is employed to include the swirl velocity component as well as the axial and radial velocity components. The numerical method is an implicit finite-volume technique and allows for large time steps to reach a converged steady-state solution. The inviscid fluxes are flux-split, and Gauss-Seidel line relaxation is used to accelerate convergence. Converging-diverging nozzles with exit-to-throat area ratios up to 100:1 and annular nozzles were examined. Quantities examined included Mach number and static wall pressure distributions, and oblique shock structures. As the level of swirl and viscosity in the flowfield increased the mass flow rate and thrust decreased. The technique was used to predict the flow through a typical arcjet thruster geometry. Results indicate swirl and viscosity play an important role in the complex geometry of an arcjet.

Numerical modeling of a vortex stabilized arcjet

NASA Astrophysics Data System (ADS)

Pawlas, Gary E.

1992-11-01

Arcjet thrusters are being actively considered for use in Earth orbit maneuvering applications. Experimental studies are currently the chief means of determining an optimal thruster configuration. Earlier numerical studies have failed to include all of the effects found in typical arcjets including complex geometries, viscosity, and swirling flow. Arcjet geometries are large area ratio converging nozzles with centerbodies in the subsonic portion of the nozzle. The nozzle walls serve as the anode while the centerbody functions as the cathode. Viscous effects are important because the Reynolds number, based on the throat radius, is typically less than 1,000. Experimental studies have shown that a swirl or circumferential velocity component stabilizes a constricted arc. This dissertation describes the equations governing flow through a constricted arcjet thruster. An assumption that the flowfield is in local thermodynamic equilibrium leads to a single fluid plasma temperature model. An order of magnitude analysis reveals the governing fluid mechanics equations are uncoupled from the electromagnetic field equations. A numerical method is developed to solve the governing fluid mechanics equations, the Thin Layer Navier-Stokes equations. A coordinate transformation is employed in deriving the governing equations to simplify the application of boundary conditions in complex geometries. An axisymmetric formulation is employed to include the swirl velocity component as well as the axial and radial velocity components. The numerical method is an implicit finite-volume technique and allows for large time steps to reach a converged steady-state solution. The inviscid fluxes are flux-split, and Gauss-Seidel line relaxation is used to accelerate convergence. Converging-diverging nozzles with exit-to-throat area ratios up to 100:1 and annular nozzles were examined. Quantities examined included Mach number and static wall pressure distributions, and oblique shock structures. As the level of swirl and viscosity in the flowfield increased the mass flow rate and thrust decreased.

Fluid Structure Interaction in a Cold Flow Test and Transient CFD Analysis of Out-of-Round Nozzles

NASA Technical Reports Server (NTRS)

Ruf, Joseph; Brown, Andrew; McDaniels, David; Wang, Ten-See

2010-01-01

This viewgraph presentation describes two nozzle fluid flow interactions. They include: 1) Cold flow nozzle tests with fluid-structure interaction at nozzle separated flow; and 2) CFD analysis for nozzle flow and side loads of nozzle extensions with various out-of-round cases.

Static investigation of two STOL nozzle concepts with pitch thrust-vectoring capability

NASA Technical Reports Server (NTRS)

Mason, M. L.; Burley, J. R., II

1986-01-01

A static investigation of the internal performance of two short take-off and landing (STOL) nozzle concepts with pitch thrust-vectoring capability has been conducted. An axisymmetric nozzle concept and a nonaxisymmetric nozzle concept were tested at dry and afterburning power settings. The axisymmetric concept consisted of a circular approach duct with a convergent-divergent nozzle. Pitch thrust vectoring was accomplished by vectoring the approach duct without changing the nozzle geometry. The nonaxisymmetric concept consisted of a two dimensional convergent-divergent nozzle. Pitch thrust vectoring was implemented by blocking the nozzle exit and deflecting a door in the lower nozzle flap. The test nozzle pressure ratio was varied up to 10.0, depending on model geometry. Results indicate that both pitch vectoring concepts produced resultant pitch vector angles which were nearly equal to the geometric pitch deflection angles. The axisymmetric nozzle concept had only small thrust losses at the largest pitch deflection angle of 70 deg., but the two-dimensional convergent-divergent nozzle concept had large performance losses at both of the two pitch deflection angles tested, 60 deg. and 70 deg.

Aeroacoustic Improvements to Fluidic Chevron Nozzles

NASA Technical Reports Server (NTRS)

Henderson, Brenda; Kinzie, Kevin; Whitmire, Julia; Abeysinghe, Amal

2006-01-01

Fluidic chevrons use injected air near the trailing edge of a nozzle to emulate mixing and jet noise reduction characteristics of mechanical chevrons. While previous investigations of "first generation" fluidic chevron nozzles showed only marginal improvements in effective perceived noise levels when compared to nozzles without injection, significant improvements in noise reduction characteristics were achieved through redesigned "second generation" nozzles on a bypass ratio 5 model system. The second-generation core nozzles had improved injection passage contours, external nozzle contour lines, and nozzle trailing edges. The new fluidic chevrons resulted in reduced overall sound pressure levels over that of the baseline nozzle for all observation angles. Injection ports with steep injection angles produced lower overall sound pressure levels than those produced by shallow injection angles. The reductions in overall sound pressure levels were the result of noise reductions at low frequencies. In contrast to the first-generation nozzles, only marginal increases in high frequency noise over that of the baseline nozzle were observed for the second-generation nozzles. The effective perceived noise levels of the new fluidic chevrons are shown to approach those of the core mechanical chevrons.

Small scale noise and wind tunnel tests of upper surface blowing nozzle flap concepts. Volume 1. Aerodynamic test results

NASA Technical Reports Server (NTRS)

Renselaer, D. J.; Nishida, R. S.; Wilkin, C. A.

1975-01-01

The results and analyses of aerodynamic and acoustic studies conducted on the small scale noise and wind tunnel tests of upper surface blowing nozzle flap concepts are presented. Various types of nozzle flap concepts were tested. These are an upper surface blowing concept with a multiple slot arrangement with seven slots (seven slotted nozzle), an upper surface blowing type with a large nozzle exit at approximately mid-chord location in conjunction with a powered trailing edge flap with multiple slots (split flow or partially slotted nozzle). In addition, aerodynamic tests were continued on a similar multi-slotted nozzle flap, but with 14 slots. All three types of nozzle flap concepts tested appear to be about equal in overall aerodynamic performance but with the split flow nozzle somewhat better than the other two nozzle flaps in the landing approach mode. All nozzle flaps can be deflected to a large angle to increase drag without significant loss in lift. The nozzle flap concepts appear to be viable aerodynamic drag modulation devices for landing.

Dynamic loads on twin jet exhaust nozzles due to shock noise

NASA Technical Reports Server (NTRS)

Norum, T. D.; Shearin, J. G.

1986-01-01

Acoustic near field data were collected with model single and twin jet nozzles to determine if closely spaced nozzles produce higher acoustic loading than do single nozzles. The tests were spurred by structural failure of the B-1 exhaust nozzle external flaps and similar damage on the F-15. The test was performed using two 5/8 in. ID pipes machined and placed side-by-side to mimic B-1 nozzles. A microphone mounted on the internozzle fairing measured acoustic levels near the nozzle exit plane. The nozzles oscillated significantly more than did a single nozzle over a wide range of nozzle pressure ratios. Acoustic levels in the dual jets exceeded single jet noise by as much as 20 dB, making acoustic resonance a definite candidate for structural damage in the twin jet configuration.

Nozzle Side Load Testing and Analysis at Marshall Space Flight Center

NASA Technical Reports Server (NTRS)

Ruf, Joseph H.; McDaniels, David M.; Brown, Andrew M.

2009-01-01

Realistic estimates of nozzle side loads, the off-axis forces that develop during engine start and shutdown, are important in the design cycle of a rocket engine. The estimated magnitude of the nozzle side loads has a large impact on the design of the nozzle shell and the engine s thrust vector control system. In 2004 Marshall Space Flight Center (MSFC) began developing a capability to quantify the relative magnitude of side loads caused by different types of nozzle contours. The MSFC Nozzle Test Facility was modified to measure nozzle side loads during simulated nozzle start. Side load results from cold flow tests on two nozzle test articles, one with a truncated ideal contour and one with a parabolic contour are provided. The experimental approach, nozzle contour designs and wall static pressures are also discussed

The Data Acquisition and Control Systems of the Jet Noise Laboratory at the NASA Langley Research Center

NASA Technical Reports Server (NTRS)

Jansen, B. J., Jr.

1998-01-01

The features of the data acquisition and control systems of the NASA Langley Research Center's Jet Noise Laboratory are presented. The Jet Noise Laboratory is a facility that simulates realistic mixed flow turbofan jet engine nozzle exhaust systems in simulated flight. The system is capable of acquiring data for a complete take-off assessment of noise and nozzle performance. This paper describes the development of an integrated system to control and measure the behavior of model jet nozzles featuring dual independent high pressure combusting air streams with wind tunnel flow. The acquisition and control system is capable of simultaneous measurement of forces, moments, static and dynamic model pressures and temperatures, and jet noise. The design concepts for the coordination of the control computers and multiple data acquisition computers and instruments are discussed. The control system design and implementation are explained, describing the features, equipment, and the experiences of using a primarily Personal Computer based system. Areas for future development are examined.

High speed jet noise research at NASA Lewis

NASA Astrophysics Data System (ADS)

Krejsa, Eugene A.; Cooper, B. A.; Kim, C. M.; Khavaran, Abbas

1992-04-01

The source noise portion of the High Speed Research Program at NASA LeRC is focused on jet noise reduction. A number of jet noise reduction concepts are being investigated. These include two concepts, the Pratt & Whitney ejector suppressor nozzle and the General Electric (GE) 2D-CD mixer ejector nozzle, that rely on ejectors to entrain significant amounts of ambient air to mix with the engine exhaust to reduce the final exhaust velocity. Another concept, the GE 'Flade Nozzle' uses fan bypass air at takeoff to reduce the mixed exhaust velocity and to create a fluid shield around a mixer suppressor. Additional concepts are being investigated at Georgia Tech Research Institute and at NASA LeRC. These will be discussed in more detail in later figures. Analytical methods for jet noise prediction are also being developed. Efforts in this area include upgrades to the GE MGB jet mixing noise prediction procedure, evaluation of shock noise prediction procedures, and efforts to predict jet noise directly from the unsteady Navier-Stokes equation.

Investigation of the Rocket Induced Flow Field in a Rectangular Duct

NASA Technical Reports Server (NTRS)

Landrum, D. Brian; Thames, Mignon; Parkinson, Doug; Gautney, Serena; Hawk, Clark

1999-01-01

Several tests were performed on a one-sixth scale Rocket Based Combined Cycle (RBCC) engine model at the University of Alabama in Huntsville. The UAH RBCC facility consists of a rectangular duct with a vertical strut mounted in the center. The scaled strut consists of two supersonic rocket nozzles with an embedded vertical turbine between the rocket nozzles. The tests included mass flow, flow visualization and horizontal pressure traverses. The mass flow test indicated a c:hoked condition when the rocket chamber pressure is between 200 psi and 300 psi. The flow visualization tests narrowed the rocket chamber pressure range from, 250 psi to 300 psi. Also, from this t.est, an assumption of a minimum

Nozzle extension design status report

NASA Technical Reports Server (NTRS)

Classen, L. B.

1972-01-01

Twenty possible concepts of a possible nozzle/nozzle extension interface were originated. Not all of the concepts were considered worthy of analysis time. Six of them were thermally analyzed and three were stress analyzed. These analyses were done to determine which of the concepts would have the best chance of succeeding, that is, they were a screening process which was to allow rating of one concept against another. This was done because adequate material properties to determine absolute stress levels were not available at the time of the analyses. Through all of the concepts still exhibit some areas of negative margin of safety, concept no. 1 shows good promise that, with slight modifications, it could have all positive margins of safety. Another significant question, regarding these designs, has to do with the Grafoil seals and insulators. Some additional data was just recently received on Grafoil properties, but it was too late to incorporate in the analyses. The new data were not significantly different from the properties which were used.

High speed jet noise research at NASA Lewis

NASA Technical Reports Server (NTRS)

Krejsa, Eugene A.; Cooper, B. A.; Kim, C. M.; Khavaran, Abbas

1992-01-01

The source noise portion of the High Speed Research Program at NASA LeRC is focused on jet noise reduction. A number of jet noise reduction concepts are being investigated. These include two concepts, the Pratt & Whitney ejector suppressor nozzle and the General Electric (GE) 2D-CD mixer ejector nozzle, that rely on ejectors to entrain significant amounts of ambient air to mix with the engine exhaust to reduce the final exhaust velocity. Another concept, the GE 'Flade Nozzle' uses fan bypass air at takeoff to reduce the mixed exhaust velocity and to create a fluid shield around a mixer suppressor. Additional concepts are being investigated at Georgia Tech Research Institute and at NASA LeRC. These will be discussed in more detail in later figures. Analytical methods for jet noise prediction are also being developed. Efforts in this area include upgrades to the GE MGB jet mixing noise prediction procedure, evaluation of shock noise prediction procedures, and efforts to predict jet noise directly from the unsteady Navier-Stokes equation.

Model aerodynamic test results for two variable cycle engine coannular exhaust systems at simulated takeoff and cruise conditions. Comprehensive data report. Volume 1: Design layouts

NASA Technical Reports Server (NTRS)

Nelson, D. P.

1981-01-01

The design layouts and detailed design drawings of coannular exhaust nozzle models for a supersonic propulsion system are presented. The layout drawings show the assembly of the component parts for each configuration. A listing of the component parts is also given.

Development Status of the NASA MC-1 (Fastrac) Engine

NASA Technical Reports Server (NTRS)

Ballard, Richard O.; Olive, Tim; Turner, James E. (Technical Monitor)

2000-01-01

The MC-1 (formerly known as the Fastrac 60K) Engine is being developed for the X-34 technology demonstrator vehicle. It is a pump-fed liquid rocket engine with fixed thrust operating at one rated power level of 60,000 lbf vacuum thrust using a 15:1 area ratio nozzle (slightly higher for the 30:1 flight nozzle). Engine system development testing of the MC-1 has been ongoing since 24 Oct 1998. To date, 48 tests have been conducted on three engines using three separate test stands. This paper will provide some details of the engine, the tests conducted, and the lessons learned to date.

Dimpled/grooved face on a fuel injection nozzle body for flame stabilization and related method

DOEpatents

Uhm, Jong Ho; Johnson, Thomas Edward; Kim, Kwanwoo; Zuo, Baifang

2013-08-20

A fuel injection head for a fuel nozzle used in a gas turbine combustor includes a substantially hollow body formed with an upstream end face, a downstream end face and a peripheral wall extending therebetween. A plurality of pre-mix tubes or passages extend axially through the hollow body with inlets at the upstream end face and outlets at the downstream end face. An exterior surface of the downstream end face is formed with three-dimensional surface features that increase a total surface area of the exterior surface as compared to a substantially flat, planar downstream end face.

Basic features of boron isotope separation by SILARC method in the two-step iterative static model

NASA Astrophysics Data System (ADS)

Lyakhov, K. A.; Lee, H. J.

2013-05-01

In this paper we develop a new static model for boron isotope separation by the laser assisted retardation of condensation method (SILARC) on the basis of model proposed by Jeff Eerkens. Our model is thought to be adequate to so-called two-step iterative scheme for isotope separation. This rather simple model helps to understand combined action on boron separation by SILARC method of all important parameters and relations between them. These parameters include carrier gas, molar fraction of BCl3 molecules in carrier gas, laser pulse intensity, gas pulse duration, gas pressure and temperature in reservoir and irradiation cells, optimal irradiation cell and skimmer chamber volumes, and optimal nozzle throughput. A method for finding optimal values of these parameters based on some objective function global minimum search was suggested. It turns out that minimum of this objective function is directly related to the minimum of total energy consumed, and total setup volume. Relations between nozzle throat area, IC volume, laser intensity, number of nozzles, number of vacuum pumps, and required isotope production rate were derived. Two types of industrial scale irradiation cells are compared. The first one has one large throughput slit nozzle, while the second one has numerous small nozzles arranged in parallel arrays for better overlap with laser beam. It is shown that the last one outperforms the former one significantly. It is argued that NO2 is the best carrier gas for boron isotope separation from the point of view of energy efficiency and Ar from the point of view of setup compactness.

Hook nozzle arrangement for supporting airfoil vanes

DOEpatents

Shaffer, James E.; Norton, Paul F.

1996-01-01

A gas turbine engine's nozzle structure includes a nozzle support ring, a plurality of shroud segments, and a plurality of airfoil vanes. The plurality of shroud segments are distributed around the nozzle support ring. Each airfoil vane is connected to a corresponding shroud segment so that the airfoil vanes are also distributed around the nozzle support ring. Each shroud segment has a hook engaging the nozzle support ring so that the shroud segments and corresponding airfoil vanes are supported by the nozzle support ring. The nozzle support ring, the shroud segments, and the airfoil vanes may be ceramic.

Hook nozzle arrangement for supporting airfoil vanes

DOEpatents

Shaffer, J.E.; Norton, P.F.

1996-02-20

A gas turbine engine`s nozzle structure includes a nozzle support ring, a plurality of shroud segments, and a plurality of airfoil vanes. The plurality of shroud segments are distributed around the nozzle support ring. Each airfoil vane is connected to a corresponding shroud segment so that the airfoil vanes are also distributed around the nozzle support ring. Each shroud segment has a hook engaging the nozzle support ring so that the shroud segments and corresponding airfoil vanes are supported by the nozzle support ring. The nozzle support ring, the shroud segments, and the airfoil vanes may be ceramic. 8 figs.

Pressure Regulator With Internal Ejector Circulation Pump, Flow and Pressure Measurement Porting, and Fuel Cell System Integration Options

NASA Technical Reports Server (NTRS)

Vasquez, Arturo

2011-01-01

An advanced reactant pressure regulator with an internal ejector reactant circulation pump has been developed to support NASA's future fuel cell power systems needs. These needs include reliable and safe operation in variable-gravity environments, and for exploration activities with both manned and un manned vehicles. This product was developed for use in Proton Exchange Membrane Fuel Cell (PEMFC) power plant reactant circulation systems, but the design could also be applied to other fuel cell system types, (e.g., solid-oxide or alkaline) or for other gas pressure regulation and circulation needs. The regulator design includes porting for measurement of flow and pressure at key points in the system, and also includes several fuel cell system integration options. NASA has recognized ejectors as a viable alternative to mechanical pumps for use in spacecraft fuel cell power systems. The ejector motive force is provided by a variable, high-pressure supply gas that travels through the ejector s jet nozzle, whereby the pressure energy of the fluid stream is converted to kinetic energy in the gas jet. The ejector can produce circulation-to-consumption-flow ratios that are relatively high (2-3 times), and this phenomenon can potentially (with proper consideration of the remainder of the fuel cell system s design) be used to provide completely for reactant pre-humidification and product water removal in a fuel cell system. Specifically, a custom pressure regulator has been developed that includes: (1) an ejector reactant circulation pump (with interchangeable jet nozzles and mixer sections, gas-tight sliding and static seals in required locations, and internal fluid porting for pressure-sensing at the regulator's control elements) and (2) internal fluid porting to allow for flow rate and system pressure measurements. The fluid porting also allows for inclusion of purge, relief, and vacuum-breaker check valves on the regulator assembly. In addition, this regulator could also be used with NASA's advanced nonflow-through fuel cell power systems by simply incorporating a jet nozzle with an appropriate nozzle diameter.

High mass throughput particle generation using multiple nozzle spraying

DOE Office of Scientific and Technical Information (OSTI.GOV)

Pui, David Y. H.; Chen, Da-Ren

Spraying apparatus and methods that employ multiple nozzle structures for producing multiple sprays of particles, e.g., nanoparticles, for various applications, e.g., pharmaceuticals, are provided. For example, an electrospray dispensing device may include a plurality of nozzle structures, wherein each nozzle structure is separated from adjacent nozzle structures by an internozzle distance. Sprays of particles are established from the nozzle structures by creating a nonuniform electrical field between the nozzle structures and an electrode electrically isolated therefrom.

High mass throughput particle generation using multiple nozzle spraying

DOEpatents

Pui, David Y.H.; Chen, Da-Ren

2004-07-20

Spraying apparatus and methods that employ multiple nozzle structures for producing multiple sprays of particles, e.g., nanoparticles, for various applications, e.g., pharmaceuticals, are provided. For example, an electrospray dispensing device may include a plurality of nozzle structures, wherein each nozzle structure is separated from adjacent nozzle structures by an internozzle distance. Sprays of particles are established from the nozzle structures by creating a nonuniform electrical field between the nozzle structures and an electrode electrically isolated therefrom.

High mass throughput particle generation using multiple nozzle spraying

DOEpatents

Pui, David Y. H. [Plymouth, MN; Chen, Da-Ren [Creve Coeur, MO

2009-03-03

Spraying apparatus and methods that employ multiple nozzle structures for producing multiple sprays of particles, e.g., nanoparticles, for various applications, e.g., pharmaceuticals, are provided. For example, an electrospray dispensing device may include a plurality of nozzle structures, wherein each nozzle structure is separated from adjacent nozzle structures by an internozzle distance. Sprays of particles are established from the nozzle structures by creating a nonuniform electrical field between the nozzle structures and an electrode electrically isolated therefrom.

Internal performance of two nozzles utilizing gimbal concepts for thrust vectoring

NASA Technical Reports Server (NTRS)

Berrier, Bobby L.; Taylor, John G.

1990-01-01

The internal performance of an axisymmetric convergent-divergent nozzle and a nonaxisymmetric convergent-divergent nozzle, both of which utilized a gimbal type mechanism for thrust vectoring was evaluated in the Static Test Facility of the Langley 16-Foot Transonic Tunnel. The nonaxisymmetric nozzle used the gimbal concept for yaw thrust vectoring only; pitch thrust vectoring was accomplished by simultaneous deflection of the upper and lower divergent flaps. The model geometric parameters investigated were pitch vector angle for the axisymmetric nozzle and pitch vector angle, yaw vector angle, nozzle throat aspect ratio, and nozzle expansion ratio for the nonaxisymmetric nozzle. All tests were conducted with no external flow, and nozzle pressure ratio was varied from 2.0 to approximately 12.0.

Turbomachine combustor nozzle including a monolithic nozzle component and method of forming the same

DOE Office of Scientific and Technical Information (OSTI.GOV)

Stoia, Lucas John; Melton, Patrick Benedict; Johnson, Thomas Edward

A turbomachine combustor nozzle includes a monolithic nozzle component having a plate element and a plurality of nozzle elements. Each of the plurality of nozzle elements includes a first end extending from the plate element to a second end. The plate element and plurality of nozzle elements are formed as a unitary component. A plate member is joined with the nozzle component. The plate member includes an outer edge that defines first and second surfaces and a plurality of openings extending between the first and second surfaces. The plurality of openings are configured and disposed to register with and receivemore » the second end of corresponding ones of the plurality of nozzle elements.« less

The TICTOP nozzle: a new nozzle contouring concept

NASA Astrophysics Data System (ADS)

Frey, Manuel; Makowka, Konrad; Aichner, Thomas

2017-06-01

Currently, mainly two types of nozzle contouring methods are applied in space propulsion: the truncated ideal contour (TIC) and the thrust-optimized parabola (TOP). This article presents a new nozzle contouring method called TICTOP, combining elements of TIC and TOP design. The resulting nozzle is shock-free as the TIC and therefore does not induce restricted shock separation leading to excessive side-loads. Simultaneously, the TICTOP nozzle will allow higher nozzle wall exit pressures and hence give a better separation margin than is the case for a TIC. Hence, this new nozzle type combines the good properties of TIC and TOP nozzles and eliminates their drawbacks. It is especially suited for first stage application in launchers where flow separation and side-loads are design drivers.

F-15/nonaxisymmetric nozzle system integration study support program

NASA Technical Reports Server (NTRS)

Stevens, H. L.

1978-01-01

Nozzle and cooling methods were defined and analyzed to provide a viable system for demonstration 2-D nozzle technology on the F-15 aircraft. Two candidate cooling systems applied to each nozzle were evaluated. The F-100 engine mount and case modifications requirements were analyzed and the actuation and control system requirements for two dimensional nozzles were defined. Nozzle performance changes relative to the axisymmetric baseline nozzle were evaluated and performance and weight characteristics for axisymmetric reference configurations were estimated. The infrared radiation characteristics of these nozzles installed on the F-100 engine were predicted. A full scale development plan with associated costs to carry the F100 engine/two-dimensional (2-D) nozzle through flight tests was defined.

Computational study of performance characteristics for truncated conical aerospike nozzles

NASA Astrophysics Data System (ADS)

Nair, Prasanth P.; Suryan, Abhilash; Kim, Heuy Dong

2017-12-01

Aerospike nozzles are advanced rocket nozzles that can maintain its aerodynamic efficiency over a wide range of altitudes. It belongs to class of altitude compensating nozzles. A vehicle with an aerospike nozzle uses less fuel at low altitudes due to its altitude adaptability, where most missions have the greatest need for thrust. Aerospike nozzles are better suited to Single Stage to Orbit (SSTO) missions compared to conventional nozzles. In the current study, the flow through 20% and 40% aerospike nozzle is analyzed in detail using computational fluid dynamics technique. Steady state analysis with implicit formulation is carried out. Reynolds averaged Navier-Stokes equations are solved with the Spalart-Allmaras turbulence model. The results are compared with experimental results from previous work. The transition from open wake to closed wake happens in lower Nozzle Pressure Ratio for 20% as compared to 40% aerospike nozzle.

Two stroke engine exhaust emissions separator

DOEpatents

Turner, Terry D.; Wilding, Bruce M.; McKellar, Michael G.; Raterman, Kevin T.

2003-04-22

A separator for substantially resolving at least one component of a process stream, such as from the exhaust of an internal combustion engine. The separator includes a body defining a chamber therein. A nozzle housing is located proximate the chamber. An exhaust inlet is in communication with the nozzle housing and the chamber. A nozzle assembly is positioned in the nozzle housing and includes a nozzle moveable within and relative to the nozzle housing. The nozzle includes at least one passage formed therethrough such that a process stream entering the exhaust inlet connection passes through the passage formed in the nozzle and imparts a substantially rotational flow to the process stream as it enters the chamber. A positioning member is configured to position the nozzle relative to the nozzle housing in response to changes in process stream pressure thereby adjusting flowrate of said process stream entering into the chamber.

Two stroke engine exhaust emissions separator

DOEpatents

Turner, Terry D.; Wilding, Bruce M.; McKellar, Michael G.; Raterman, Kevin T.

2002-01-01

A separator for substantially resolving at least one component of a process stream, such as from the exhaust of an internal combustion engine. The separator includes a body defining a chamber therein. A nozzle housing is located proximate the chamber. An exhaust inlet is in communication with the nozzle housing and the chamber. A nozzle assembly is positioned in the nozzle housing and includes a nozzle moveable within and relative to the nozzle housing. The nozzle includes at least one passage formed therethrough such that a process stream entering the exhaust inlet connection passes through the passage formed in the nozzle, which imparts a substantially rotational flow to the process stream as it enters the chamber. A positioning member is configured to position the nozzle relative to the nozzle housing in response to changes in process stream pressure to adjust flowrate of said process stream entering into the chamber.

Parametric study of solar thermal rocket nozzle performance

NASA Technical Reports Server (NTRS)

Pearson, J. Boise; Landrum, D. Brian; Hawk, Clark W.

1995-01-01

This paper details a numerical investigation of performance losses in low-thrust solar thermal rocket nozzles. The effects of nozzle geometry on three types of losses were studied; finite rate dissociation-recombination kinetic losses, two dimensional axisymmetric divergence losses, and compressible viscous boundary layer losses. Short nozzle lengths and supersonic flow produce short residence times in the nozzle and a nearly frozen flow, resulting in large kinetic losses. Variations in geometry have a minimal effect on kinetic losses. Divergence losses are relatively small, and careful shaping of the nozzle can nearly eliminate them. The boundary layer in these small nozzles can grow to a major fraction of nozzle radius, and cause large losses. These losses are attributed to viscous drag on the nozzle walls and flow blockage by the boundary layer, especially in the throat region. Careful shaping of the nozzle can produce a significant reduction in viscous losses.

Control Design for an Advanced Geared Turbofan Engine

NASA Technical Reports Server (NTRS)

Chapman, Jeffryes W.; Litt, Jonathan S.

2017-01-01

This paper describes the design process for the control system of an advanced geared turbofan engine. This process is applied to a simulation that is representative of a 30,000 pound-force thrust class concept engine with two main spools, ultra-high bypass ratio, and a variable area fan nozzle. Control system requirements constrain the non-linear engine model as it operates throughout its flight envelope of sea level to 40,000 feet and from 0 to 0.8 Mach. The purpose of this paper is to review the engine control design process for an advanced turbofan engine configuration. The control architecture selected for this project was developed from literature and reflects a configuration that utilizes a proportional integral controller with sets of limiters that enable the engine to operate safely throughout its flight envelope. Simulation results show the overall system meets performance requirements without exceeding operational limits.

Static internal performance of an axisymmetric nozzle with multiaxis thrust-vectoring capability

NASA Technical Reports Server (NTRS)

Carson, George T., Jr.; Capone, Francis J.

1991-01-01

An investigation was conducted in the static test facility of the Langley 16 Foot Transonic Tunnel in order to determine the internal performance characteristics of a multiaxis thrust vectoring axisymmetric nozzle. Thrust vectoring for this nozzle was achieved by deflection of only the divergent section of this nozzle. The effects of nozzle power setting and divergent flap length were studied at nozzle deflection angles of 0 to 30 at nozzle pressure ratios up to 8.0.

Endoscope system with plasma flushing and coaxial round jet nozzle for off-pump cardiac surgery.

PubMed

Horiuchi, Tetsuya; Masamune, Ken; Iwase, Yuki; Ymashita, Hiromasa; Tsukihara, Hiroyuki; Motomura, Noboru; Ohta, Yuji; Dohi, Takeyoshi

2011-07-01

To develop a new endoscope for performing simple surgical tasks inside the blood-filled cardiac atrium/chamber, that is, "off-pump" cardiac surgeries. We developed the endoscope system with plasma flushing and coaxial round jet nozzle. The "plasma flushing" system was invented to observe the interior of the blood-filled heart by displacing blood cells in front of the endoscope tip. However, some areas could not be observed with simple flushing of the liquid because the flushed liquid mixed with blood. Further, a large amount of liquid had to be flushed, which posed a risk of cardiac damage caused by excess volume. Therefore, to safely capture high-resolution images of the interior of the heart, an endoscope with a coaxial round jet nozzle through which plasma is flushed has been developed. And to reduce the volume of flushed liquid, the synchronization system of heartbeat and the endoscope system with plasma flushing has been developed. We conducted an in vivo experiment to determine whether we could observe intracardiac tissues in swine without the use of a heart-lung machine. As a result, we successfully observed intracardiac tissues without using a heart-lung machine. By using a coaxial nozzle, we could even observe the tricuspid valve. Moreover, we were able to save up to 30% of the flushed liquid by replacing the original system with a synchronization system. And we evaluated the performance of the endoscope with the coaxial round jet nozzle by conducting fluid analysis and an in vitro experiment. We successfully observed intracardiac tissues without using a heart-lung machine. By using a coaxial nozzle, we could even observe the tricuspid valve. And by replacing an original system to a synchronization system, we were able to save up to 30% of the flushed liquid. As a follow-up study, we plan to create a surgical flexible device for valve disease that can grasp, staple, and repair cardiac valves by endoscopic visualization.