Thermal Modeling of a Novel Air-Cooled Temperature Swing Adsorption Compressor (AC-TSAC)
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Hannah Alpert↗
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Modeling the Integrated Thermal Amine Scrubber and Air Cooled-Temperature Swing Adsorbent Compressor
As part of a larger CO2 removal architecture study, a model was developed in Aspen Custom Modeler of the Thermal Amine Scrubber (TAS), an ISS flight experiment intended to scrub CO2 from the cabin air. In order to minimize mass losses, it was determined the TAS should work in conjunction with the Air Cooled-Temperature Swing Adsorption Compressor (AC-TSAC) to reduce the CO2 for O2 production. First the TAS removes CO2 from the cabin atmosphere, then the AC-TSAC compresses the CO2 and feeds it in a steady steam to a Sabatier or Bosch process. This document details an effort to add a model of the AC-TSAC to the existing TAS model to determine if the two hardware systems will have deleterious effects on each other’s performance. It was found the AC-TSAC provided sufficient vacuum for the TAS’s desorption to occur, but not the same level of vacuum as space vacuum. This lower quality vacuum led to a reduction in the total cyclical uptake capacity of the TAS and thus a lower CO2 removal under the same conditions. The TAS did not remove enough CO2 to fill the AC-TSAC bed to capacity during the production phase which caused the AC-TSAC to run out of CO2 during production. This renders the addition of the AC-TSAC pointless. It is believed these effects can be overcome with a redesign of one or both systems.
TAS↗
Altitude-Wind-Tunnel Investigation of the 19B-2, 19B-8 and 19XB-1 Jet- Propulsion Engines: Analysis of Compressor Performance - 4
Investigations were conducted in the Cleveland altitude wind tunnel to determine the performance and operational characteristics of the 19B-2, 19B-8, and 19XS-1 turbojet engines. One objective was to determine the effect of altitude, flight Mach number, and tail-pipe-nozzle area on the performance characteristics of the six-stage and ten-stage axial-flow compressors of the 19B-8 and 19XB-1 engines, respectively, The data were obtained over a range of simulated altitudes and flight Mach numbers. At each simulated flight condition the engine was run over its full operable range of speeds. Performance characteristics of the 19B-8 and 19XB-1 compressors for the range of operation obtainable in the turboJet-engine installation are presented. Compressor characteristics are presented as functions of air flow corrected to sea-level conditions, compressor Mach number, and compressor load coefficient. For the range of compressor operation investigated, changes in Reynolds number had no measurable effect on the relations among compressor Mach number, corrected air flow, compressor load coefficient, compressor pressure ratio, and compressor efficiency. The operating lines for the 19B-8 compressor lay on the low-air-flow side of the region of maximum compressor efficiency; the 19B-8 compressor operated at higher average pressure coefficients per stage and produced a lower over-all pressure ratio than did the 19XB-1 compressor.
Dietz, Robert O.↗
Damage-Tolerant Fan Casings for Jet Engines
All turbofan engines work on the same principle. A large fan at the front of the engine draws air in. A portion of the air enters the compressor, but a greater portion passes on the outside of the engine this is called bypass air. The air that enters the compressor then passes through several stages of rotating fan blades that compress the air more, and then it passes into the combustor. In the combustor, fuel is injected into the airstream, and the fuel-air mixture is ignited. The hot gasses produced expand rapidly to the rear, and the engine reacts by moving forward. If there is a flaw in the system, such as an unexpected obstruction, the fan blade can break, spin off, and harm other engine components. Fan casings, therefore, need to be strong enough to contain errant blades and damage-tolerant to withstand the punishment of a loose blade-turned-projectile. NASA has spearheaded research into improving jet engine fan casings, ultimately discovering a cost-effective approach to manufacturing damage-tolerant fan cases that also boast significant weight reduction. In an aircraft, weight reduction translates directly into fuel burn savings, increased payload, and greater aircraft range. This technology increases safety and structural integrity; is an attractive, viable option for engine manufacturers, because of the low-cost manufacturing; and it is a practical alternative for customers, as it has the added cost saving benefits of the weight reduction.
Source record↗
Centrifugal Compressor Surge Margin Improved With Diffuser Hub Surface Air Injection
Aerodynamic stability is an important parameter in the design of compressors for aircraft gas turbine engines. Compression system instabilities can cause compressor surge, which may lead to the loss of an aircraft. As a result, engine designers include a margin of safety between the operating line of the engine and the stability limit line of the compressor. The margin of safety is typically referred to as "surge margin." Achieving the highest possible level of surge margin while meeting design point performance objectives is the goal of the compressor designer. However, performance goals often must be compromised in order to achieve adequate levels of surge margin. Techniques to improve surge margin will permit more aggressive compressor designs. Centrifugal compressor surge margin improvement was demonstrated at the NASA Glenn Research Center by injecting air into the vaned diffuser of a 4:1-pressure-ratio centrifugal compressor. Tests were performed using injector nozzles located on the diffuser hub surface of a vane-island diffuser in the vaneless region between the impeller trailing edge and the diffuser-vane leading edge. The nozzle flow path and discharge shape were designed to produce an air stream that remained tangent to the hub surface as it traveled into the diffuser passage. Injector nozzles were located near the leading edge of 23 of the 24 diffuser vanes. One passage did not contain an injector so that instrumentation located in that passage would be preserved. Several orientations of the injected stream relative to the diffuser vane leading edge were tested over a range of injected flow rates. Only steady flow (nonpulsed) air injection was tested. At 100 percent of the design speed, a 15-percent improvement in the baseline surge margin was achieved with a nozzle orientation that produced a jet that was bisected by the diffuser vane leading edge. Other orientations also improved the baseline surge margin. Tests were conducted at speeds below the design speed, and similar results were obtained. In most cases, the greatest improvement in surge margin occurred at fairly low levels of injected flow rate. Externally supplied injection air was used in these experiments. However, the injected flow rates that provided the greatest benefit could be produced using injection air that is recirculating between the diffuser discharge and nozzles located in the diffuser vaneless region. Future experiments will evaluate the effectiveness of recirculating air injection.
Skoch, Gary J.↗
Method of experimentally determining radial distributions of velocity through axial-flow compressor
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AIR FLOW - COMPRESSORS, AXIAL↗
Design of floated shoe close clearance seals for supersonic jet engine compressors.
Floated shoe close running air seals for supersonic jet engine compressors, discussing axial and radial seals design
Cheng, H.↗
Evaluation of centrifugal compressor performance with water injection
The effects of water injection on a compressor are presented. To determine the effects of varying water-air ratio, the compressor was operated at a constant equivalent impeller speed over a range of water-air ratios and weight flows. Operation over a range of weight flows at one water-air ratio and two inlet air temperatures was carried out to obtain an indication of the effects of varying inlet air temperature. Beyond a water-air ratio of 0.03 there was no increase in maximum air-weight flow, a negligible rise in peak total-pressure ratio, and a decrease in peak adiabatic efficiency. An increase in inlet air temperature resulted in an increase in the magnitude of evaporation. An analysis of data indicated that the magnitude of evaporation within the compressor impeller was small.
COMPRESSORS - RADIAL FLOW↗
8- by 6-Foot Supersonic Wind Tunnel Compressor Inspected
The NASA Glenn Research Center's 8- by 6-Foot Supersonic Wind Tunnel (8 6 SWT) is NASA's only transonic propulsion wind tunnel. The test section speed range is between Mach 0.25 and 2.0. The 9- by 15-Foot Low-Speed Wind Tunnel (9 15 LWST), which has a speed range from 0 to 175 mph, is housed in the return leg of the 8 6 SWT and uses the same compressor. The 8 6 SWT uses a large, seven-stage axial flow compressor to drive the air through the tunnel. The compressor is 17 ft in diameter and is rated at 1600 m3 (56,600 ft3) of air/sec. It is driven by three electric motors with a combined horsepower of 87,000. A close examination of this compressor was performed in 2001, the first time since February of 1966.
Krupar, Martin J.↗
Compressor Stall Recovery Through Tip Injection Assessed
Aerodynamic stability is a fundamental limit in the compressor design process. The development of robust techniques for increasing stability has several benefits: enabling higher loading and fewer blades, increasing safety throughout a mission, increasing tolerance to stage mismatch during part-speed operation and speed transients, and providing an opportunity to match stages at the compressor maximum efficiency point, thus reducing fuel burn. Mass injection upstream of the tip of a high-speed axial compressor rotor is a stability enhancement approach known to be effective in suppressing stall in tip-critical rotors if the injection is activated before stall occurs. This approach to stall suppression requires that a reliable stall warning system be available. Tests have recently been performed to assess whether steady injection can also be used to recover from fully developed stall. If mass injection is effective in recovering from stall quickly enough to avoid structural damage or loss of engine power, then a stall warning system may not be required. The stall recovery tests were performed on a transonic compressor rotor at its design tip speed of 1475 ft/sec using four injectors evenly spaced around the compressor case upstream of the rotor. The injectors were connected to an external air source. In an actual engine application, the injected air would be supplied with compressor bleed air. The injectors were isolated from the air source by a fast-acting butterfly valve. With the injectors turned off, the compressor was throttled into stall. Air injection was then activated with no change in throttle setting by opening the butterfly valve. The compressor recovered from stall at a fixed throttle setting with the aid of tip injection. The unsteady operating characteristic of the rotor was measured during these tests using high-response pressure sensors located upstream and downstream of the rotor. The figure shows the results, where the unsteady pressure and mass flow are superimposed on the steady operating characteristic. The total injected mass flow was equal to 1.3 percent of the compressor flow. The solid line with no solid squares on it denotes the operating point during the beginning of throttle closure and the initial drop into stall. The gray traces denote the operating point during an additional throttle closure that occurred over the next 1200 rotor revolutions (4 sec). The dashed line denotes the recovery from stall that occurred during 90 rotor revolutions (0.3 sec) after the injectors were activated with no change in throttle setting. Tip injection not only recovers the compressor from stall, but also restores the compressor to its pre-stall level of pressure rise. In contrast, standard stall recovery schemes such as compressor bleed, stator vane actuation, or engine throttle modulation result in a loss of pressure rise across the compressor, which results in a loss of engine power.
Suder, Ken L.↗
New Compressor Added to Glenn's 450- psig Combustion Air System
In September 1999, the Central Process Systems Engineering Branch and the Maintenance and the Central Process Systems Operations Branch, released for service a new high pressure compressor to supplement the 450-psig Combustion Air System at the NASA Glenn Research Center at Lewis Field. The new compressor, designated C-18, is located in Glenn s Central Air Equipment Building and is remotely operated from the Central Control Building. C-18 can provide 40 pounds per second (pps) of airflow at pressure to our research customers. This capability augments our existing system capacity (compressors C 4 at 38 pps and C-5 at 32 pps), which is generated from Glenn's Engine Research Building. The C-18 compressor was originally part of Glenn's 21-Inch Hypersonic Tunnel, which was transferred from the Jet Propulsion Laboratory to Glenn in the mid-1980's. With the investment of construction of facilities funding, the compressor was modified, new mechanical and electrical support equipment were purchased, and the unit was installed in the basement of the Central Air Equipment Building. After several weeks of checkout and troubleshooting, the new compressor was ready for long-term, reliable operations. With a total of 110 pps in airflow now available, Glenn is well positioned to support the high-pressure air test requirements of our research customers.
Swan, Jeffrey A.↗
Long-Duration Testing of a Temperature-Swing Adsorption Compressor for Carbon Dioxide for Closed-Loop Air Revitalization Systems
This paper describes the results of long-duration testing of a temperature-swing adsorption compressor that has application in the International Space Station (ISS) and future spacecraft for closing the air revitalization loop. The air revitalization system of the ISS operates in an open loop mode and relies on the resupply of oxygen and other consumables from Earth for the life support of astronauts. A compressor is required for delivering the carbon dioxide from a removal assembly to a reduction unit to recover oxygen and thereby closing the air-loop. The TSAC is a solid-state compressor that has the capability to remove CO2 from a low-pressure source, and subsequently store, compress, and deliver at a higher pressure as required by a processor. The TSAC is an ideal interface device for CO2 removal and reduction units in the air revitalization loop of a spacecraft for oxygen recovery. The TSAC was developed and its operation was successfully verified in integration tests with the flight-like Carbon Dioxide Removal Assembly (CDRA) at Marshall Space Flight Center prior to the long-duration tests. Long-duration tests reveal the impacts of repeated thermal cycling on the compressor components and the adsorbent material.
Rosen, Micha↗
Development of compressor end seals, stator interstage seals, and stator pivot seals in advanced air breathing propulsion systems Semiannual report, 1 Jan. - 30 Jun. 1966
Design of compressor end, stator interface, and stator pivot seals for advanced air breathing engines
SEALING↗
Static sea-level performance of an axial-flow-compressor turbojet engine with an air-cooled turbine
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Cochran, Reeves P↗
The impact of circulation control on rotary aircraft controls systems
Application of circulation to rotary wing systems is a new development. Efforts to determine the near and far field flow patterns and to analytically predict those flow patterns have been underway for some years. Rotary wing applications present a new set of challenges in circulation control technology. Rotary wing sections must accommodate substantial Mach number, free stream dynamic pressure and section angle of attack variation at each flight condition within the design envelope. They must also be capable of short term circulation blowing modulation to produce control moments and vibration alleviation in addition to a lift augmentation function. Control system design must provide this primary control moment, vibration alleviation and lift augmentation function. To accomplish this, one must simultaneously control the compressed air source and its distribution. The control law algorithm must therefore address the compressor as the air source, the plenum as the air pressure storage and the pneumatic flow gates or valves that distribute and meter the stored pressure to the rotating blades. Also, mechanical collective blade pitch, rotor shaft angle of attack and engine power control must be maintained.
Kingloff, R. F.↗
Effect of humidity on jet engine axial-flow compressor performance
Two problems related to the ingestion of humid air into jet engine axial compressors have been studied: (1) the changes in the performance of the compressor in the absence of condensation and (2) the changes in the entry conditions to the compressor when condensation occurs. Regarding the first, the extent of changes are predicted and also measured in the case of a six-stage compressor operated with air-methane gas mixture utilizing the similarities in the thermodynamic properties between water vapor and methane. For the condensation process in an inlet, a model is described that takes into account the presence of micro particulates and the flow field changes in the vicinity of the inlet wall. Finally, the effects of humidity on engine performance are discussed in relation to engine trim and control schemes.
Ehresman, C. M.↗
Investigation of an Experimental Supersonic Axial-Flow Compressor
An investigation is in progress at the Langley Laboratory of the NACA to explore the possibilities of axial-flow compressors operating with supersonic velocities relative to the blade rows. The first phase of this investigation, a study of supersonic diffusers, has been reported. The second phase, an analysis of supersonic compressors, has also been reported. Preliminary calculations have shown that very high pressure ratios across a stage, together with somewhat increased mass flows, are possible with compressors which decelerate air through the speed of sound in their rotor blading. These performance characteristics are desirable in compressors for aircraft jet propulsion units, gas turbines, or superchargers. The third phase, presented here, is a preliminary experimental investigation of a supersonic compressor designed to produce a high pressure ratio in a single stage.
Erwin, John R.↗
Numerical Simulation of Non-Rotating and Rotating Coolant Channel Flow Fields
Future generations of ultra high bypass-ratio jet engines will require far higher pressure ratios and operating temperatures than those of current engines. For the foreseeable future, engine materials will not be able to withstand the high temperatures without some form of cooling. In particular the turbine blades, which are under high thermal as well as mechanical loads, must be cooled. Cooling of turbine blades is achieved by bleeding air from the compressor stage of the engine through complicated internal passages in the turbine blades (internal cooling, including jet-impingement cooling) and by bleeding small amounts of air into the boundary layer of the external flow through small discrete holes on the surface of the blade (film cooling and transpiration cooling). The cooling must be done using a minimum amount of air or any increases in efficiency gained through higher operating temperature will be lost due to added load on the compressor stage. Turbine cooling schemes have traditionally been based on extensive empirical data bases, quasi-one-dimensional computational fluid dynamics (CFD) analysis, and trial and error. With improved capabilities of CFD, these traditional methods can be augmented by full three-dimensional simulations of the coolant flow to predict in detail the heat transfer and metal temperatures. Several aspects of turbine coolant flows make such application of CFD difficult, thus a highly effective CFD methodology must be used. First, high resolution of the flow field is required to attain the needed accuracy for heat transfer predictions, making highly efficient flow solvers essential for such computations. Second, the geometries of the flow passages are complicated but must be modeled accurately in order to capture all important details of the flow. This makes grid generation and grid quality important issues. Finally, since coolant flows are turbulent and separated the effects of turbulence must be modeled with a low Reynolds number turbulence model to accurately predict details of heat transfer.
Rigby, David L.↗