Conceptual design and integration of a space station resistojet propulsion assembly

The resistojet propulsion module is designed as a simple, long life, low risk system offering operational flexibility to the space station program. It can dispose of a wide variety of typical space station waste fluids by using them as propellants for orbital maintenance. A high temperature mode offers relatively high specific impulse with long life while a low temperature mode can propulsively dispose of mixtures that contain oxygen or hydrocarbons without reducing thruster life or generating particulates in the plume. A low duty cycle and a plume that is confined to a small aft region minimizes the impacts on the users. Simple interfaces with other space station systems facilitate integration. It is concluded that there are no major obstacles and many advantages to developing, installing, and operating a resistojet propulsion module aboard the Initial Operational Capability (IOC) space station

Two-phase flow

An experimental program to characterize the spray from candidate nozzles for icing-cloud simulation is discussed. One canidate nozzle, which is currently used for icing research, has been characterized for flow and drop size. The median-volume diameter (MVD) from this air-assist nozzle is compared with correlations in the literature. The new experimental spray facility is discussed, and the drop-size instruments are discussed in detail. Since there is no absolute standard for drop-size measurements and there are other limitations, such as drop -size range and velocity range, several instruments are used and results are compared. A two-phase model was developed at Pennsylvania State University. The model uses the k-epsilon model of turbulence in the continous phase. Three methods for treating the discrete phase are used: (1) a locally homogeneous flow (LHF) model, (2) a deterministic separated flow (DSF) model, and (3) a stochastic separated flow (SSF) model. In the LHF model both phases have the same velocity and temperature at each point. The DSF model provides interphase transport but ignores the effects of turbulent fluctuations. In the SSF model the drops interact with turbulent eddies whose properties are determined by the k-epsilon turbulence model. The two-phase flow model has been extended to include the effects of evaporation and combustion

Combustor technology for future aircraft

The continuing improvement of aircraft gas turbine engine operating efficiencies involves increases in overall engine pressure ratio increases that will result in combustor inlet pressure and temperature increases, greater combustion temperature rises, and higher combustor exit temperatures. These conditions entail the development of fuel injectors generating uniform circumferential and radial temperature patterns, as well as combustor liner configurations and materials capable of withstanding increased thermal radiation even as the amount of cooling air is reduced. Low NO(x)-emitting combustor concepts are required which will employ staged combustion. The development status of component technologies answering these requirements are presently evaluated

Analysis of Total-Pressure Loss and Airflow Distribution for Annular Gas Turbine Combustors

Computer analysis of total pressure loss and air flow distribution in annular flow gas turbine combustion chamber

Low NO(x) potential of gas turbine engines

The purpose is to correlate emission levels of gas turbine engines. The predictions of NO(x) emissions are based on a review of the literature of previous low NO(x) combustor programs and analytical chemical kinetic calculations. Concepts included in the literature review consisted of lean-premixed-prevaporized (LPP), rich burn/quick quench/lean burn (RQL), and direct injection. The NO(x) emissions were found to be an exponential function of adiabatic combustion temperature over a wide range of inlet temperatures, pressures and (lean) fuel-air ratios. A simple correlation of NO(x) formation with time was not found. The LPP and direct injection (using gaseous fuels) concepts have the lowest NO(x) emissions of the three concepts. The RQL data has higher values of NO(x) than the LPP concept, probably due to the stoichiometric temperatures and NO(x) production that occur during the quench step. Improvements in the quick quench step could reduce the NO(x) emissions to the LPP levels. The low NO(x) potential of LPP is offset by the operational disadvantages of its narrow stability limits and its susceptibility to autoignition/flashback. The Rich-Burn/Quick-Quench/Lean-Burn (RQL) and the direct injection concepts have the advantage of wider stability limits comparable to conventional combustors

Effects of Spent Cooling and Swirler Angle on a 9-Point Swirl-Venturi Injector

This paper presents multipoint Lean-Direct-Injection (LDI) emissions results for flame tube combustion tests at an inlet pressure of 1034 kPa and inlet temperatures between 835 and 865 K; these are the combustor inlet conditions that the High Speed Research (HSR) program used for supersonic cruise. It focuses on one class of LDI geometry, 9-point swirl-venturi LDI (SV-LDI). Two parameters are compared in this paper: the use of dome cooling air and the swirler blade angle. Dome cooling air is called "spent cooling" and is at combustor inlet conditions. Three cooling variations are studied: cooling at the venturi throat, cooling at the dome face, and no cooling at all. Two swirler blade angles are studied: 45deg and 60deg. The HSR 9-point SV-LDI emissions are also compared to a similar 9-point SV-LDI design which was used in the later ultra-efficient engine technology (UEET) program. The HSR and UEET designs cannot be compared directly due to different UEET combustor conditions. Therefore, this paper uses previously published UEET correlation equations to make comparisons. Results show that using a 45deg swirler produces lower NOx emissions than using a 60deg swirler. This is consistent with the later UEET results. The effects of spent cooling depend on swirler angle, spent cooling location, and the test conditions. For the configuration with 45deg swirlers, spent cooling delivers lower NOx emissions when it is injected at the throat. For the 60deg swirler, spent cooling does not have much effect on NOx emissions. These results might be caused by the location and the intensity of the flame recirculation zone

Effects of Spent Cooling and Swirler Angle on a 9-point Swirl-Venturi Injector

This paper presents multipoint lean-direct-injection (LDI) emissions results for flame tube combustion tests at an inlet pressure of 1034 kPa and inlet temperatures between 835 and 865 K; these are the combustor inlet conditions that the High Speed Research (HSR) program used for supersonic cruise. It focuses on one class of LDI geometry, 9-point swirl-venturi LDI (SV-LDI). Two parameters are compared in this paper: the use of dome cooling air and the swirler blade angle. Dome cooling air is called 'spent cooling' and is at combustor inlet conditions. Three cooling variations are studied: cooling at the venturi throat, cooling at the dome face, and no cooling at all. Two swirler blade angles are studied: 45 deg and 60 deg. The HSR 9-point SV-LDI emissions are also compared to a similar 9-point SV-LDI design which was used in the later ultra-efficient engine technology (UEET) program. The HSR and UEET designs cannot be compared directly due to different UEET combustor conditions. Therefore, this paper uses previously published UEET correlation equations to make comparisons. Results show that using a 45 deg swirler produces lower NOx emissions than using a 60 deg swirler. This is consistent with the later UEET results. The effects of spent cooling depend on swirler angle, spent cooling location, and the test conditions. For the configuration with 45 deg swirlers, spent cooling delivers lower NOx emissions when it is injected at the throat. For the 60 deg swirler, spent cooling does not have much effect on NOx emissions. These results might be caused by the location and the intensity of the flame recirculation zone

Low NOx, Lean Direct Wall Injection Combustor Concept Developed

The low-emissions combustor development at the NASA Glenn Research Center is directed toward advanced high-pressure aircraft gas turbine applications. The emphasis of this research is to reduce nitrogen oxides (NOx) at high-power conditions and to maintain carbon monoxide and unburned hydrocarbons at their current low levels at low-power conditions. Low-NOx combustors can be classified into rich burn and lean burn concepts. Lean burn combustors can be further classified into lean-premixed-prevaporized (LPP) and lean direct injection (LDI) combustors. In both concepts, all the combustor air, except for liner cooling flow, enters through the combustor dome so that the combustion occurs at the lowest possible flame temperature. The LPP concept has been shown to have the lowest NOx emissions, but for advanced high-pressure-ratio engines, the possibly of autoignition or flashback precludes its use. LDI differs from LPP in that the fuel is injected directly into the flame zone and, thus, does not have the potential for autoignition or flashback and should have greater stability. However, since it is not premixed and prevaporized, the key is good atomization and mixing of the fuel quickly and uniformly so that flame temperatures are low and NOx formation levels are comparable to those of LPP

Space station propulsion system technology

Two propulsion systems have been selected for the space station: O/H rockets for high thrust applications and the multipropellant resistojets for low thrust needs. These thruster systems integrate very well with the fluid systems on the station. Both thrusters will utilize waste fluids as their source of propellant. The O/H rocket will be fueled by electrolyzed water and the resistojets will use stored waste gases from the environmental control system and the various laboratories. This paper presents the results of experimental efforts with O/H and resistojet thrusters to determine their performance and life capability

Fundamental Study of a Single Point Lean Direct Injector. Part I: Effect of Air Swirler Angle and Injector Tip Location on Spray Characteristics

Lean direct injection (LDI) is a combustion concept to reduce oxides of nitrogen (NOx) for next generation aircraft gas turbine engines. These newer engines have cycles that increase fuel efficiency through increased operating pressures, which increase combustor inlet temperatures. NOx formation rates increase with higher temperatures; the LDI strategy avoids high temperature by staying fuel lean and away from stoichiometric burning. Thus, LDI relies on rapid and uniform fuel/air mixing. To understand this mixing process, a series of fundamental experiments are underway in the Combustion and Dynamics Facility at NASA Glenn Research Center. This first set of experiments examines cold flow (non-combusting) mixing using air and water. Using laser diagnostics, the effects of air swirler angle and injector tip location on the spray distribution, recirculation zone, and droplet size distribution are examined. Of the three swirler angles examined, 60 deg is determined to have the most even spray distribution. The injector tip location primarily shifts the flow without changing the structure, unless the flow includes a recirculation zone. When a recirculation zone is present, minimum axial velocity decreases as the injector tip moves downstream towards the venturi exit; also the droplets become more uniform in size and angular distribution