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The sodium line-reversal method has been used in some preliminary measurements of flame temperature. Improvements in the method involving a photographic recorder and a means of correcting for the dirtiness of the windows are described. The temperatures so obtained are compared with those calculated from pressure diagrams.
Pumped thermal energy storage (PTES) is a cost-effective method for grid-scale, long-duration electrical energy storage (LDES). PTES uses a heat pump cycle to transfer thermal energy from a low temperature reservoir (LTR) to a high temperature reservoir (HTR), later using a heat engine cycle to reverse the process. A key component of the PTES system, the heat pump compressor, represents a significant driver to the cost, performance and operating characteristics of the PTES system. For grid-scale charging (>100 MW), traditional compressor scaling charts indicate that the operating conditions needed would be best served by a multistage axial compressor. While frame gas turbine compressors at these power ratings exist and operate at higher pressure ratios than needed for the CO2 PTES system, the inlet pressure and fluid density of this application exceed experience values. A conceptual design of a large-scale CO2 axial compressor was completed, including mean-line estimates of the compressor performance at full power conditions. The results of the conceptual design were used to refine the PTES cycle design, and updated operating conditions provided for further aero design optimization. A roughly one third scale of the first three stages of the 100 MW compressor was designed, built, and tested at relevant PTES charging conditions. Results of this test program are compared against high-resolution computational fluid dynamics models.
Rohr Industries, Inc. specializes in the manufacturing of nacelles, thrust reversers, and other engine components, has made use of the COSMIC program in nacelle work for aircraft like the McDonnell Douglas DC-10, the French/German Airbus and the Boeing 727, 737 and 747 jetliners. They also manufacture complete nacelles for military and business aircraft and are supplying nacelle components for the Boeing 757. The nacelle requires research and development for each type of airplane because of complex airflows around the engine inlet and high pressures on the nacelle skin. The use of the computer program defines the airflow field around turbofan engine nacelle inlet and cowls. Pressures on the nacelle skin are estimated for various flight conditions and structural integrity of the nacelle design. Rohr estimates that use of this program saved six man-months of programmer time necessary to develop alternative software.
The main focus of this study is to apply a computational tool for the flow analysis of the engine that has been tested with ice crystal ingestion in the Propulsion Systems Laboratory (PSL) of NASA Glenn Research Center. A data point was selected for analysis during which the engine experienced a full roll back event due to the ice accretion on the blades and flow path of the low pressure compressor. The computational tool consists of the Numerical Propulsion System Simulation (NPSS) engine system thermodynamic cycle code, and an Euler-based compressor flow analysis code, that has an ice particle melt estimation code with the capability of determining the rate of sublimation, melting, and evaporation through the compressor blade rows. Decreasing the performance characteristics of the low pressure compressor (LPC) within the NPSS cycle analysis resulted in matching the overall engine performance parameters measured during testing at data points in short time intervals through the progression of the roll back event. Detailed analysis of the fan-core and LPC with the compressor flow analysis code simulated the effects of ice accretion by increasing the aerodynamic blockage and pressure losses through the low pressure compressor until achieving a match with the NPSS cycle analysis results, at each scan. With the additional blockages and losses in the LPC, the compressor flow analysis code results were able to numerically reproduce the performance that was determined by the NPSS cycle analysis, which was in agreement with the PSL engine test data. The compressor flow analysis indicated that the blockage due to ice accretion in the LPC exit guide vane stators caused the exit guide vane (EGV) to be nearly choked, significantly reducing the air flow rate into the core. This caused the LPC to eventually be in stall due to increasing levels of diffusion in the rotors and high incidence angles in the inlet guide vane (IGV) and EGV stators. The flow analysis indicating compressor stall is substantiated by the video images of the IGV taken during the PSL test, which showed water on the surface of the IGV flowing upstream out of the engine, indicating flow reversal, which is characteristic of a stalled compressor.
The main focus of this study is to apply a computational tool for the flow analysis of the engine that has been tested with ice crystal ingestion in the Propulsion Systems Laboratory (PSL) of NASA Glenn Research Center. A data point was selected for analysis during which the engine experienced a full roll back event due to the ice accretion on the blades and flow path of the low pressure compressor. The computational tool consists of the Numerical Propulsion System Simulation (NPSS) engine system thermodynamic cycle code, and an Euler-based compressor flow analysis code, that has an ice particle melt estimation code with the capability of determining the rate of sublimation, melting, and evaporation through the compressor blade rows. Decreasing the performance characteristics of the low pressure compressor (LPC) within the NPSS cycle analysis resulted in matching the overall engine performance parameters measured during testing at data points in short time intervals through the progression of the roll back event. Detailed analysis of the fan-core and LPC with the compressor flow analysis code simulated the effects of ice accretion by increasing the aerodynamic blockage and pressure losses through the low pressure compressor until achieving a match with the NPSS cycle analysis results, at each scan. With the additional blockages and losses in the LPC, the compressor flow analysis code results were able to numerically reproduce the performance that was determined by the NPSS cycle analysis, which was in agreement with the PSL engine test data. The compressor flow analysis indicated that the blockage due to ice accretion in the LPC exit guide vane stators caused the exit guide vane (EGV) to be nearly choked, significantly reducing the air flow rate into the core. This caused the LPC to eventually be in stall due to increasing levels of diffusion in the rotors and high incidence angles in the inlet guide vane (IGV) and EGV stators. The flow analysis indicating compressor stall is substantiated by the video images of the IGV taken during the PSL test, which showed water on the surface of the IGV flowing upstream out of the engine, indicating flow reversal, which is characteristic of a stalled compressor.
The changes in thrust minus drag performance as well as longitudinal and directional stability and control characteristics of a single-engine jet aircraft attributable to an in-flight thrust reverser of the blocker-deflector door type were investigated in a 16-foot transonic wind tunnel. The longitudinal and directional stability data are presented. Test conditions simulated landing approach conditions as well as high speed maneuvering such as may be required for combat or steep descent from high altitude.
A full-size variable-pitch fan engine was tested in the Ames 40 by 80 foot wind tunnel to determine the effect of forward velocity and crosswind on reverse-thrust performance. Two flight-type inlet configurations were tested, and a flared fan nozzle was installed as an inlet for reverse-thrust operation. Steady-state reverse-thrust performance was obtained up to 54 m/s (105 knots). An abrupt decrease in reverse thrust occurred at about 30 m/s (60 knots). Reverse thrust was established following forward-to-reverse thrust transients both statically and with forward velocities only up to 30 m/s.
Inlet mouthpiece with variable geometry improves low-speed performance of inlet (or ejector) of jet engines by use of reversed cowl-flap mechanism. Flaps can be adjusted mechanically or system can be operated with pressure taps set so pressure on inlet face is always smaller than pressure on back of inlet.
The performance impact of flow reversal at the inlet of an airbreathing rotating detonation engine (RDE) is investigated using 2 and 3-dimensional computational fluid dynamic (CFD) simulations. Flow reversal, or backflow, occurs in RDE inlets in the high-pressure region directly behind the rotating detonation front. This is also where most of the engine thrust or pressure gain is produced. The amount of backflow relative to throughflow depends on the inlet design. For the present work, a simple annular ‘slit’ design is used. The simulations are idealized in several ways, including that fuel and air are premixed, but prevented from reacting when within the inlet region. The results indicate that even with idealizations, the impact of inlet backflow on pressure gain can be substantial. The simulations also reveal an intriguing instability that develops in certain configurations. The mass flow rate into the RDE begins to oscillate at a regular frequency that is substantially less than the detonation rotational frequency. This is accompanied by oscillations in the detonation height. The oscillation amplitude grows over time until the detonation ultimately fails. Both the performance and instability results emphasize the need for carefully designed RDE inlets that provide low loss when flow is in the forward direction, but high resistance when the flow is reversed. Development of such high-diodicity inlets is critical to achieving pressure gain in airbreathing RDE’s.
The performance impact of flow reversal at the inlet of an airbreathing rotating detonation engine (RDE) is investigated using 2 and 3-dimensional computational fluid dynamic (CFD) simulations. Flow reversal, or backflow, occurs in RDE inlets in the high-pressure region directly behind the rotating detonation front. This is also where most of the engine thrust or pressure gain is produced. The amount of backflow relative to throughflow depends on the inlet design. For the present work, a simple annular ‘slit’ design is used. The simulations are idealized in several ways, including that fuel and air are premixed, but prevented from reacting when within the inlet region. The results indicate that even with idealizations, the impact of inlet backflow on pressure gain can be substantial. The simulations also reveal an intriguing instability that develops in certain configurations. The mass flow rate into the RDE begins to oscillate at a regular frequency that is substantially less than the detonation rotational frequency. This is accompanied by oscillations in the detonation height. The oscillation amplitude grows over time until the detonation ultimately fails. Both the performance and instability results emphasize the need for carefully designed RDE inlets that provide low loss when flow is in the forward direction, but high resistance when the flow is reversed. Development of such high-diodicity inlets is critical to achieving pressure gain in airbreathing RDE’s.
The paper describes an advanced composite fairing designed to reduce drag on DC-9 nacelles as a part of the NASA Engine Component Improvement Program. This fairing is the aft enclosure for the thrust reverser actuator system on JT8D engine nacelles and is subjected to a 500 F exhaust flow during the reverse thrust. A reduced-drag configuration was developed by using in-flight tuft surveys for flow visualization in order to identify areas with low-quality flow, and then modifying the aerodynamic lines to improve the flow. A fabrication method for molding the part in an autoclave was developed; this material system is suitable for 500 F. The resultant composite fairing reduces the overall aircraft drag 1% with a weight reduction of 40% when compared with a metal component.
A two-dimensional convergent-divergent exhaust nozzle designed and fabricated by Pratt and Whitney Aircraft was installed on a PW1128 turbofan engine and tested during thrust reverser operation in an altitude facility at NASA Lewis Research Center. A unique collection system was used to capture the thrust reverser exhaust gas and transport it to the primary exhaust collector. Tests were conducted at three flight conditions with varying amounts of thrust reverse at each condition. Some reverser exhaust gas spillage by the collection system was encountered but engine performance was unaffected at all flight conditions tested. Based on the results of this test program, the feasibility of altitude testing of advanced multi-function exhaust nozzle systems has been demonstrated.
A two-dimensional convergent-divergent exhaust nozzle designed and fabricated by Pratt and Whitney Aircraft was installed on a PW1128 turbofan engine and tested during thrust reverser operation in an altitude facility at NASA Lewis Research Center. A unique collection system was used to capture the thrust reverser exhaust gas and transport it to the primary exhaust collector. Tests were conducted at three flight conditions with varying amounts of thrust reverse at each condition. Some reverser exhaust gas spillage by the collection system was encountered but engine performance was unaffected at all flight conditions tested. Based on the results of this test program, the feasibility of altitude testing of advanced multifunction exhaust nozzle systems has been demonstrated.
Two-dimensional metallic transition metal dichalcogenides offer high electrical conductivity and large surface areas for electrocatalysis, yet their inherent basal planes are catalytically inert. Here, we present an atomic-scale vacancy engineering strategy to activate the basal surfaces of metallic WSe 2 for reversible oxygen electrocatalysis. This approach, based on intentionally designed substitutional metal doping, promotes the spontaneous formation of selenium vacancies while preserving the metallic 1 T′ phase, thereby creating highly reactive and oxygen-affinitive sites. Density functional theory calculations reveal that these vacancy-mediated metal complexes dramatically lower the energy barriers for initial oxygen adsorption, enabling dissociative oxygen adsorption. Operando and ex-situ spectroscopic analyses confirm that vacancy-mediated metal complexes transform into dynamic Se/W-oxide intermediates under operating conditions. Se/W-oxides on the surface experimentally and theoretically prove electrocatalytic activity and reversibility. Applying this strategy in lithium–oxygen batteries, the basal-plane activated WSe 2 shows high discharge capacities (9868 mA h g −1 , corresponding to 3947 mA h g$^{-1}_{cathode}$), impressive cycle retention over 550 cycles at 1000 mA h g −1 , and outstanding rate–capability over a wide current–density range (100–3000 mA g −1 ) during 256 cycles.
Results from experimental work on model scale cascade reversers with cold airflow are presented. Sound power level directivity and spectral characteristics for cascade reversers are reported. Effect of cascade exit area ratio, vane profile shape, and emission arc are discussed. Model equivalent diameters varied from 3 to 5 inches, pressure ratios range from 1.15 to 3.0. Depending on the reverser type, acoustic power was proportional to the 4 1/2 to 6th power of ideal jet velocity. Reverser noise peaked at higher frequency and was more omnidirectional than nozzle-alone jet noise. Appreciable reduction in sideline noise was obtained from plane shields. Airfoil-vaned cascades were the most aerodynamically efficient and least noisy reversers. Scaling of cascade reverser data to example aircraft engines showed all cascades above the 95 PNdB sideline goal from STOL aircraft. However, the airfoil-vaned reverser has a good potential for meeting this goal for high-bypass (low pressure ratio) exhausts.
The Advanced Aircraft Flight Simulator is equipped with a motorized mechanism that simulates a two engine throttle control system that can be operated via a computer driven performance management system or manually by the pilots. The throttle control system incorporates features to simulate normal engine operations and thrust reverse and vary the force feel to meet a variety of research needs. While additional testing to integrate the work required is principally now in software design, since the mechanical aspects function correctly. The mechanism is an important part of the flight control system and provides the capability to conduct human factors research of flight crews with advanced aircraft systems under various flight conditions such as go arounds, coupled instrument flight rule approaches, normal and ground operations and emergencies that would or would not normally be experienced in actual flight.
A removable slider shoe for mounting on a translating sleeve of a jet engine cowl and · duct. The sleeve in-eludes a slider which moves in a track on the cowl when the engine is in a reverse thrust mode in order to expose airflow turning vanes. The movable shoe is mounted on the slider.
A flight investigation was made to determine the effect of various vertical-tail modifications and of some combinations of these modifications on the directional stability and control characteristics of a propeller-driven fighter airplane. Six different vertical-tail configurations were investigated to determine the lateral-directional oscillation characteristics, the sideslip characteristics, the yaw due to ailerons in rudder-fixed rolls from turns and pull-outs, the trim changes due to speed changes, and the tim changes due to power changes. Results of the tests showed that increasing the aspect ratio of the vertical tail by 40 percent while increasing the area by only 12 percent approximately doubled the directional stability of the airplane. The pilots considered the directional characteristics of the airplane unsatisfactory with original vertical tail but satisfactory with the enlarged vertical tail. The ventral and dorsal fins tested had little effect on the directional stability of the airplane but were effective in eliminating rudder-force reversals in high-engine-power sideslips.