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At least 127 records · Page 7

LWS/SET Technology Experiment Carrier

This paper examines the approach taken to building a low-cost, modular spacecraft bus that can be used to support a variety of technology experiments in different space environments. It describes the techniques used and design drivers considered to ensure experiment independence from as yet selected host spacecraft. It describes the technology experiment carriers that will support NASA's Living With a Star Space Environment Testbed space missions. NASA has initiated the Living With a Star (LWS) Program to develop a better scientific understanding to address the aspects of the connected Sun-Earth system that affect life and society. A principal goal of the program is to bridge the gap between science, engineering, and user application communities. The Space Environment Testbed (SET) Project is one element of LWS. The Project will enable future science, operational, and commercial objectives in space and atmospheric environments by improving engineering approaches to the accommodation and/or mitigation of the effects of solar variability on technological systems. The SET Project is highly budget constrained and must seek to take advantage of as yet undetermined partnering opportunities for access to space. SET will conduct technology validation experiments hosted on available flight opportunities. The SET Testbeds will be developed in a manner that minimizes the requirements for accommodation, and will be flown as flight opportunities become available. To access the widest range of flight opportunities, two key development requirements are to maintain flexibility with respect to accommodation constraints and to have the capability to respond quickly to flight opportunities. Experiments, already developed to the technology readiness level of needing flight validation in the variable Sun-Earth environment, will be selected on the basis of the need for the subject technology, readiness for flight, need for flight resources and particular orbit. Experiments will be accumulated by the Project and manifested for specific flight opportunities as they become available. The SET Carrier is designed to present a standard set of interfaces to SET technology experiments and to be modular and flexible enough to interface to a variety of possible host spacecraft. The Carrier will have core components and mission unique components. Once the core carrier elements have been developed, only the mission unique components need to be defined and developed for any particular mission. This approach will minimize the mission specific cost and development schedule for a given flight opportunity. The standard set of interfaces provided by SET to experiments allows them to be developed independent of the particulars of a host spacecraft. The Carrier will provide the power, communication, and the necessary monitoring features to operate experiments. The Carrier will also provide all of the mechanical assemblies and harnesses required to adapt experiments to a particular host. Experiments may be hosted locally with the Carrier or remotely on the host spacecraft. The Carrier design will allow a single Carrier to support a variable number of experiments and will include features that support the ability to incrementally add experiments without disturbing the core architecture.

Sherman, Barry↗

Decision Gate Process for Assessment of a NASA Technology Development Portfolio

The NASA Dust Management Project (DMP) was established to provide technologies (to Technology Readiness Level (TRL) 6) required to address adverse effects of lunar dust to humans and to exploration systems and equipment, to reduce life cycle cost and risk, and to increase the probability of sustainable and successful lunar missions. The technology portfolio of DMP consisted of different categories of technologies whose final product was either a technology solution in itself, or one that contributes toward a dust mitigation strategy for a particular application. A Decision Gate Process (DGP) was developed to assess and validate the achievement and priority of the dust mitigation technologies as the technologies progress through the development cycle. The DGP was part of continuous technology assessment and was a critical element of DMP risk management. At the core of the process were technology-specific criteria developed to measure the success of each DMP technology in attaining the technology readiness levels assigned to each decision gate. The DGP accounts for both categories of technologies and qualifies the technology progression from technology development tasks to application areas. The process provided opportunities to validate performance, as well as to identify non-performance in time to adjust resources and direction. This paper describes the overall philosophy of the DGP and the methodology for implementation for DMP, and describes the method for defining the technology evaluation criteria. The process is illustrated by example of an application to a specific DMP technology.

Kohli, Rajiv↗

Lithium-Ion Batteries for Aerospace Applications

This presentation reviews: (1) the goals and objectives, (2) the NASA and Airforce requirements, (3) the potential near term missions, (4) management approach, (5) the technical approach and (6) the program road map. The objectives of the program include: (1) develop high specific energy and long life lithium ion cells and smart batteries for aerospace and defense applications, (2) establish domestic production sources, and to demonstrate technological readiness for various missions. The management approach is to encourage the teaming of universities, R&D organizations, and battery manufacturing companies, to build on existing commercial and government technology, and to develop two sources for manufacturing cells and batteries. The technological approach includes: (1) develop advanced electrode materials and electrolytes to achieve improved low temperature performance and long cycle life, (2) optimize cell design to improve specific energy, cycle life and safety, (3) establish manufacturing processes to ensure predictable performance, (4) establish manufacturing processes to ensure predictable performance, (5) develop aerospace lithium ion cells in various AH sizes and voltages, (6) develop electronics for smart battery management, (7) develop a performance database required for various applications, and (8) demonstrate technology readiness for the various missions. Charts which review the requirements for the Li-ion battery development program are presented.

Surampudi, S.↗

Benefits of Using a Mars Forward Strategy for Lunar Surface Systems

This paper identifies potential risk reduction, cost savings and programmatic procurement benefits of a Mars Forward Lunar Surface System architecture that provides commonality or evolutionary development paths for lunar surface system elements applicable to Mars surface systems. The objective of this paper is to identify the potential benefits for incorporating a Mars Forward development strategy into the planned Project Constellation Lunar Surface System Architecture. The benefits include cost savings, technology readiness, and design validation of systems that would be applicable to lunar and Mars surface systems. The paper presents a survey of previous lunar and Mars surface systems design concepts and provides an assessment of previous conclusions concerning those systems in light of the current Project Constellation Exploration Architectures. The operational requirements for current Project Constellation lunar and Mars surface system elements are compared and evaluated to identify the potential risk reduction strategies that build on lunar surface systems to reduce the technical and programmatic risks for Mars exploration. Risk reduction for rapidly evolving technologies is achieved through systematic evolution of technologies and components based on Moore's Law superimposed on the typical NASA systems engineering project development "V-cycle" described in NASA NPR 7120.5. Risk reduction for established or slowly evolving technologies is achieved through a process called the Mars-Ready Platform strategy in which incremental improvements lead from the initial lunar surface system components to Mars-Ready technologies. The potential programmatic benefits of the Mars Forward strategy are provided in terms of the transition from the lunar exploration campaign to the Mars exploration campaign. By utilizing a sequential combined procurement strategy for lunar and Mars exploration surface systems, the overall budget wedges for exploration systems are reduced and the costly technological development gap between the lunar and Mars programs can be eliminated. This provides a sustained level of technological competitiveness as well as maintaining a stable engineering and manufacturing capability throughout the entire duration of Project Constellation.

Mulqueen, Jack↗

NLF technology is ready to go

Natural laminar flow (NLF) can reduce drag on aircraft developed using modern structural design methods. Modern metal and composite construction methods can meet NLF requirements for subsonic commuter and business airframes. NLF research at NASA concentrates on expanding the practical application of NLF drag reduction technology; payoffs include progress with liquid-crystal flow visualization, NLF on three-dimensional bodies, and the effects of acoustics on laminar stability. Fuel savings from 2 to 4 percent are expected if laminar flow could be achieved over the forward 50 percent of engine nacelles on large transports depending on the configuration. It is concluded that the skill required to use NLF for drag reduction depends on understanding the conservative design corridors within which laminar flow is durable and reliable.

Holmes, Bruce J.↗

SEP full-scale wing technology development

A technology development program has generated a detail design of a lightweight 25 kW solar array for Solar Electric Propulsion (SEP). The fabrication and test of a full-scale array wing, 32.0 m x 4.06 m, is in progress to demonstrate technology readiness for fabrication, testing and flight of the large area lightweight solar array system. This paper presents the requirements for the 66 W/kg array and the component testing that has been performed to demonstrate technology readiness in the areas of SEP mission environmental survival, zero-gravity flat-fold array retraction, and NDT development testing. A zero-gravity test program was performed in the NASA KC-135 aircraft using a three-panel, full-width segment of the flat-fold array blanket with three degrees of panel stiffening. The full-scale solar array wing being fabricated is composed of three electrical modules, 76 x 200 cm, and mass simulator panels each 76 x 400 cm employing 2 x 4 cm glass slides (4.5 panels) and aluminum mass simulators (35 panels).

Elms, R. V., Jr.↗

Review of NASA In-Space Propulsion Technology Program Inflatable Decelerator Investments

The NASA In-Space Propulsion Technology (ISPT) Program is managed by the NASA Headquarters Science Mission Directorate and is implemented by the Marshall Space Flight Center in Huntsville, Alabama. The ISPT objective is to fund development of promising in- space propulsion technologies that can decrease flight times, decrease cost, or increase delivered payload mass for future science missions. Before ISPT will invest in a technology, the Technology Readiness Level (TRL) of the concept must be estimated to be at TRL 3. A TRL 3 signifies that the technical community agrees that the feasibility of the concept has been proven through experiment or analysis. One of the highest priority technology investments for ISPT is Aerocapture. The aerocapture maneuver uses a planetary atmosphere to reduce or alter the speed of a vehicle allowing for quick, propellantless (or using very little propellant) orbit capture. The atmosphere is used as a brake, transferring the energy associated with the vehicle s high speed into thermal energy. The ISPT Aerocapture Technology Area (ATA) is currently investing in the development of advanced lightweight ablative thermal protection systems, high temperature composite structures, and heat-flux sensors for rigid aeroshells. The heritage of rigid aeroshells extends back to the Apollo era and this technology will most likely be used by the first generation aerocapture vehicle. As a second generation aerocapture technology, ISPT is investing in three inflatable aerodynamic decelerator concepts for planetary aerocapture. They are: trailing ballute (balloon-parachute), attached afterbody ballute, and an inflatable aeroshell. ISPT also leverages the NASA Small Business Innovative Research Program for additional inflatable decelerator technology development. In mid-2004 ISPT requested an independent review of the three inflatable decelerator technologies funded directly by ISPT to validate the TRL and to identify technology maturation concerns. An independent panel with expertise in advanced thin film materials, aerothermodynamics, trajectory design, and inflatable structures was convened to assess the ISPT investments. The panel considered all major technical subsystems including materials, aerothermodynamics, structural dynamics, packaging, and inflation systems. The panel assessed the overall technology readiness of inflatable decelerators to be a 3 and identified fluid-structure interaction, aeroheating, and structural adhesives to be of highest technical concern.

Richardson, Erin H.↗

Review of NASA In-Space Propulsion Technology Program Inflatable Decelerator Investments

The NASA In-Space Propulsion Technology (ISPT) Program is managed by the NASA Headquarters Science Mission Directorate and is implemented by the Marshall Space Flight Center in Huntsville, Alabama. The ISPT objective is to fund development of promising in-space propulsion technologies that can decrease flight times, decrease cost, or increase delivered payload mass for future science missions. Before ISPT will invest in a technology, the Technology Readiness Level (TRL) of the concept must be estimated to be at TRL 3. A TRL 3 signifies that the technical community agrees that the feasibility of the concept has been proven through experiment or analysis. One of the highest priority technology investments for ISPT is Aerocapture. The aerocapture maneuver uses a planetary atmosphere to reduce or alter the speed of a vehicle allowing for quick, propellantless (or using very little propellant) orbit capture. The atmosphere is used as a brake, transferring the energy associated with the vehicle's high speed into thermal energy. The ISPT Aerocapture Technology Area (ATA) is currently investing in the development of advanced lightweight ablative thermal protection systems, high temperature composite structures, and heat-flux sensors for rigid aeroshells. The heritage of rigid aeroshells extends back to the Apollo era and this technology will most likely be used by the first generation aerocapture vehicle. As a second generation aerocapture technology, ISPT is investing in three inflatable aerodynamic decelerator concepts for planetary aerocapture. They are: trailing ballute (balloon-parachute), attached afterbody ballute, and an inflatable aeroshell. ISPT also leverages the NASA Small Business Innovative Research Program for additional inflatable decelerator technology development. In mid-2004 ISPT requested an independent review of the three inflatable decelerator technologies funded directly by ISPT to validate the TRL and to identify technology maturation concerns. An independent panel with expertise in advanced thin film materials, aerothermodynamics, trajectory design, and inflatable structures was convened to assess the ISPT investments. The panel considered all major technical subsystems including materials, aerothermodynamics, structural dynamics, packaging, and inflation systems. The panel assessed the overall technology readiness of inflatable decelerators to be a 3 and identified fluid- structure interaction, aeroheating, and structural adhesives to be of highest technical concern.

Richardson, E. H.↗

Evaluating the Acoustic Benefits of Over-the-Rotor Acoustic Treatments Installed on the Advanced Noise Control Fan

Over the last 15 years, over-the-rotor acoustic treatments have been evaluated by NASA with varying success. Recently, NASA has been developing the next generation of over-the-rotor acoustic treatments for fan noise reduction. The NASA Glenn Research Centers Advanced Noise Control Fan was used as a Low Technology Readiness Level test bed. A rapid prototyped in-duct array consisting of 50 microphones was employed, and used to correlate the in-duct analysis to the far-field acoustic levels and to validate an existing beam-former method. The goal of this testing was to improve the Technology Readiness Level of various over-the-rotor acoustic treatments by advancing the understanding of the physical mechanisms and projecting the far-field acoustic benefit.

Fan noise↗

Sensor technology for advanced space missions

The capability and applications of two sensors, Spatial, High-Accuracy, Position-Encoding Sensor (SHAPES) and Fiber Optics Rotation Sensor (FORS), for advanced missions are discussed. The multiple target, 3-D position sensing capability of SHAPES meets a critical technology need for many developing applications. A major milestone of the SHAPES task was completed on schedule on May 30, 1986, by demonstrating simultaneous ranging to eight moving targets at a rate of 10 measurements per second. The range resolution to static target was shown to be 25 microns. SHAPES scheduled technology readiness will support the sensor needs of a number of early users. The next phase in the development of SHAPES is to incorporate an angular measurement CCD to provide the full 3-dimensional sensing. A flight unit design and fabrication can be complete by FY89. FORS, with its significant improvement over present technology in lifetime, performance, weight, power, and recurrent cost, will be an important technology for future space systems. Technology readiness will be demonstrated with a FORS brassboard with fully integrated IO chips by FY88. The unique capability of miniature remote sensing heads, connected to a central system, will open up new areas in control and stability of large space structures. This application requires additional study.

Nerheim, N. M.↗

Fiber-Based, Trace-Gas, Laser Transmitter Technology Development for Space

NASA’s Goddard Space Flight Center (GSFC) is working on maturing the technology readiness of a laser transmitter designed for use in atmospheric CO2 remote-sensing. GSFC has been developing an airplane-based CO2 lidar instrument over several years to demonstrate the efficacy of the instrumentation and measurement technique and to link the science models to the instrument performance. The ultimate goal is to make space-based satellite measurements with global coverage. In order to accomplish this, we must demonstrate the technology readiness and performance of the components as well as demonstrate the required power-scaling to make the link with the required signal-to-noise-ratio (SNR). To date, all the instrument components have been shown to have the required performance with the exception of the laser transmitter.In this program we are working on a fiber-based master oscillator power amplifier (MOPA) laser transmitter architecture where we will develop a ruggedized package and perform the relevant environmental tests to demonstrate TRL-6. In this paper we will review our transmitter architecture and progress on the performance and packaging of the laser transmitter.

Laser↗

Recent Electric Propulsion Development Activities for NASA Science Missions

(The primary source of electric propulsion development throughout NASA is managed by the In-Space Propulsion Technology Project at the NASA Glenn Research Center for the Science Mission Directorate. The objective of the Electric Propulsion project area is to develop near-term electric propulsion technology to enhance or enable science missions while minimizing risk and cost to the end user. Major hardware tasks include developing NASA s Evolutionary Xenon Thruster (NEXT), developing a long-life High Voltage Hall Accelerator (HIVHAC), developing an advanced feed system, and developing cross-platform components. The objective of the NEXT task is to advance next generation ion propulsion technology readiness. The baseline NEXT system consists of a high-performance, 7-kW ion thruster; a high-efficiency, 7-kW power processor unit (PPU); a highly flexible advanced xenon propellant management system (PMS); a lightweight engine gimbal; and key elements of a digital control interface unit (DCIU) including software algorithms. This design approach was selected to provide future NASA science missions with the greatest value in mission performance benefit at a low total development cost. The objective of the HIVHAC task is to advance the Hall thruster technology readiness for science mission applications. The task seeks to increase specific impulse, throttle-ability and lifetime to make Hall propulsion systems applicable to deep space science missions. The primary application focus for the resulting Hall propulsion system would be cost-capped missions, such as competitively selected, Discovery-class missions. The objective of the advanced xenon feed system task is to demonstrate novel manufacturing techniques that will significantly reduce mass, volume, and footprint size of xenon feed systems over conventional feed systems. This task has focused on the development of a flow control module, which consists of a three-channel flow system based on a piezo-electrically actuated valve concept, as well as a pressure control module, which will regulate pressure from the propellant tank. Cross-platform component standardization and simplification are being investigated through the Standard Architecture task to reduce first user costs for implementing electric propulsion systems. Progress on current hardware development, recent test activities and future plans are discussed.

Pencil, Eric J.↗

Oxide_Oxide Ceramic Matrix Composite (CMC) Exhaust Mixer Development in the NASA Environmentally Responsible Aviation (ERA) Project

LibertyWorks®, a subsidiary of Rolls-Royce Corporation, first studied CMC (ceramic matrix composite) exhaust mixers for potential weight benefits in 2008. Oxide CMC potentially offered weight reduction, higher temperature capability, and the ability to fabricate complex-shapes for increased mixing and noise suppression. In 2010, NASA was pursuing the reduction of NOx emissions, fuel burn, and noise from turbine engines in Phase I of the Environmentally Responsible Aviation (ERA) Project (within the Integrated Systems Research Program). ERA subtasks, including those focused on CMC components, were being formulated with the goal of maturing technology from Proof of Concept Validation (Technology Readiness Level 3 (TRL 3)) to System/Subsystem or Prototype Demonstration in a Relevant Environment (TRL 6). LibertyWorks®, a subsidiary of Rolls-Royce Corporation, first studied CMC (ceramic matrix composite) exhaust mixers for potential weight benefits in 2008. Oxide CMC potentially offered weight reduction, higher temperature capability, and the ability to fabricate complex-shapes for increased mixing and noise suppression. In 2010, NASA was pursuing the reduction of NOx emissions, fuel burn, and noise from turbine engines in Phase I of the Environmentally Responsible Aviation (ERA) Project (within the Integrated Systems Research Program). ERA subtasks, including those focused on CMC components, were being formulated with the goal of maturing technology from Proof of Concept Validation (Technology Readiness Level 3 (TRL 3)) to System/Subsystem or Prototype Demonstration in a Relevant Environment (TRL 6). Oxide CMC component at both room and elevated temperatures. A TRL≈5 (Component Validation in a Relevant Environment) was attained and the CMC mixer was cleared for ground testing on a Rolls-Royce AE3007 engine for performance evaluation to achieve TRL 6.

propulsion↗

Introduction: Prediction of F-16XL Flight Flow Physics

This special section is the result of fruitful endeavors by an international group of researchers in industry, government laboratories and university-led efforts to improve the technology readiness level of their CFD solvers through comparisons with flight data collected on the F-16XL-1 aircraft at a variety of test conditions. These 1996 flight data were documented and detailed the flight-flow physics of this aircraft through surface tufts and pressures, boundary-layer rakes and skin-friction measurements. The flight project was called the Cranked Wing Aerodynamics Project (CAWAP), due to its leading-edge sweep crank (70 degrees inboard, 50 degrees outboard), and served as a basis for the International comparisons to be made, called CAWAPI. This highly focused effort was one of two vortical flow studies facilitated by the NATO Research and Technology Organization through its Applied Vehicle Panel with a title of Understanding and Modeling Vortical Flows to Improve the Technology Readiness Level for Military Aircraft. It was given a task group number of AVT-113 and had an official start date of Spring 2003. The companion part of this task group dealt with fundamentals of vortical flow from both an experimental and numerical perspective on an analytically describable 65 degree delta-wing model for which much surface pressure data had already been measured at NASA Langley Research Center at a variety of Mach and Reynolds numbers and is called the Vortex Flow Experiment - 2 (VFE-2). These two parts or facets helped one another in understanding the predictions and data that had been or were being collected.

Lamar, John E.↗

Solar Thermal Upper Stage Liquid Hydrogen Pressure Control Testing

High-energy cryogenic propellant is an essential element in future space exploration programs. Therefore, NASA and its industrial partners are committed to an advanced development/technology program that will broaden the experience base for the entire cryogenic fluid management community. Furthermore, the high cost of microgravity experiments has motivated NASA to establish government/aerospace industry teams to aggressively explore combinations of ground testing and analytical modeling to the greatest extent possible, thereby benefitting both industry and government entities. One such team consisting of ManTech SRS, Inc., Edwards Air Force Base, and Marshall Space Flight Center (MSFC) was formed to pursue a technology project designed to demonstrate technology readiness for an SRS liquid hydrogen (LH2) in-space propellant management concept. The subject testing was cooperatively performed June 21-30, 2000, through a partially reimbursable Space Act Agreement between SRS, MSFC, and the Air Force Research Laboratory. The joint statement of work used to guide the technical activity is presented in appendix A. The key elements of the SRS concept consisted of an LH2 storage and supply system that used all of the vented H2 for solar engine thrusting, accommodated pressure control without a thermodynamic vent system (TVS), and minimized or eliminated the need for a capillary liquid acquisition device (LAD). The strategy was to balance the LH2 storage tank pressure control requirements with the engine thrusting requirements to selectively provide either liquid or vapor H2 at a controlled rate to a solar thermal engine in the low-gravity environment of space operations. The overall test objective was to verify that the proposed concept could enable simultaneous control of LH2 tank pressure and feed system flow to the thruster without necessitating a TVS and a capillary LAD. The primary program objectives were designed to demonstrate technology readiness of the SRS concept at a system level as a first step toward actual flight vehicle demonstrations. More specific objectives included testing the pressure and feed control system concept hardware for functionality, operability, and performance. Valuable LH2 thermodynamic and fluid dynamics data were obtained for application to both the SRS concept and to future missions requiring space-based cryogen propellant management.

Moore, J. D.↗

Multi-Wavelength, Multi-Beam, and Polarization-Sensitive Laser Transmitter for Surface Mapping

A multi-beam, multi-color, polarized laser transmitter has been developed for mapping applications. It uses commercial off-the-shelf components for a lowcost approach for a ruggedized laser suitable for field deployment. The laser transmitter design is capable of delivering dual wavelengths, multiple beams on each wavelength with equal (or variable) intensities per beam, and a welldefined state of polarization. This laser transmitter has been flown on several airborne campaigns for the Slope Imaging Multi-Polarization Photon Counting Lidar (SIMPL) instrument, and at the time of this reporting is at a technology readiness level of between 5 and 6. The laser is a 1,064-nm microchip high-repetition-rate laser emitting energy of about 8 microjoules per pulse. The beam was frequency-doubled to 532 nm using a KTP (KTiOPO4) nonlinear crystal [other nonlinear crystals such as LBO (LiB3O5) or periodically poled lithium niobiate can be used as well, depending on the conversion efficiency requirements], and the conversion efficiency was approximately 30 percent. The KTP was under temperature control using a thermoelectric cooler and a feedback monitoring thermistor. The dual-wavelength beams were then spectrally separated and each color went through its own optical path, which consisted of a beam-shaping lens, quarterwave plate (QWP), and a birefringent crystal (in this case, a calcite crystal, but others such as vanadate can be used). The QWP and calcite crystal set was used to convert the laser beams from a linearly polarized state to circularly polarized light, which when injected into a calcite crystal, will spatially separate the circularly polarized light into the two linear polarized components. The spatial separation of the two linearly polarized components is determined by the length of the crystal. A second set of QWP and calcite then further separated the two beams into four. Additional sets of QWP and calcite can be used to further split the beams into multiple orders of two. The spatially separated beams had alternating linearly polarization states; a half-wave plate (HWP) array was then made to rotate the alternating states of A multi-beam, multi-color, polarized laser transmitter has been developed for mapping applications. It uses commercial off-the-shelf components for a lowcost approach for a ruggedized laser suitable for field deployment. The laser transmitter design is capable of delivering dual wavelengths, multiple beams on each wavelength with equal (or variable) intensities per beam, and a welldefined state of polarization. This laser transmitter has been flown on several airborne campaigns for the Slope Imaging Multi-Polarization Photon Counting Lidar (SIMPL) instrument, and at the time of this reporting is at a technology readiness level of between 5 and 6. The laser is a 1,064-nm microchip high-repetition-rate laser emitting energy of about 8 microjoules per pulse. The beam was frequency-doubled to 532 nm using a KTP (KTiOPO4) nonlinear crystal [other nonlinear crystals such as LBO (LiB3O5) or periodically poled lithium niobiate can be used as well, depending on the conversion efficiency requirements], and the conversion efficiency was approximately 30 percent. The KTP was under temperature control using a thermoelectric cooler and a feedback monitoring thermistor. The dual-wavelength beams were then spectrally separated and each color went through its own optical path, which consisted of a beam-shaping lens, quarterwave plate (QWP), and a birefringent crystal (in this case, a calcite crystal, but others such as vanadate can be used). The QWP and calcite crystal set was used to convert the laser beams from a linearly polarized state to circularly polarized light, which when injected into a calcite crystal, will spatially separate the circularly polarized light into the two linear polarized components. The spatial separation of the two linearly polarized components is determined by the length of the crystal. A cond set of QWP and calcite then further separated the two beams into four. Additional sets of QWP and calcite can be used to further split the beams into multiple orders of two. The spatially separated beams had alternating linearly polarization states; a half-wave plate (HWP) array was then made to rotate the alternating states of

Yu, Anthony W.↗

Instrumentation Working Group Summary

The Instrumentation Working Group compiled a summary of measurement techniques applicable to gas turbine engine aerosol precursors and particulates. An assessment was made of the limits, accuracy, applicability, and technology readiness of the various techniques. Despite advances made in emissions characterization of aircraft engines, uncertainties still exist in the mechanisms by which aerosols and particulates are produced in the near-field engine exhaust. To adequately assess current understanding of the formation of sulfuric acid aerosols in the exhaust plumes of gas turbine engines, measurements are required to determine the degree and importance of sulfur oxidation in the turbine and at the engine exit. Ideally, concentrations of all sulfur species would be acquired, with emphasis on SO2 and SO3. Numerous options exist for extractive and non-extractive measurement of SO2 at the engine exit, most of which are well developed. SO2 measurements should be performed first to place an upper bound on the percentage of SO2 oxidation. If extractive and non-extractive techniques indicate that a large amount of the fuel sulfur is not detected as SO2, then efforts are needed to improve techniques for SO3 measurements. Additional work will be required to account for the fuel sulfur in the engine exhaust. Chemical Ionization Mass Spectrometry (CI-MS) measurements need to be pursued, although a careful assessment needs to be made of the sampling line impact on the extracted sample composition. Efforts should also be placed on implementing non-intrusive techniques and extending their capabilities by maximizing exhaust coverage for line-of-sight measurements, as well as development of 2-D techniques, where feasible. Recommendations were made to continue engine exit and combustor measurements of particulates. Particulate measurements should include particle size distribution, mass fraction, hydration properties, and volatile fraction. However, methods to ensure that unaltered samples are obtained need to be developed. Particulate speciation was also assigned a high priority for quantifying the fractions of carbon soot, PAH, refractory materials, metals, sulfates, and nitrates. High priority was also placed on performing a comparison of particle sizing instruments. Concern was expressed by the workshop attendees who routinely make particulate measurements about the variation in number density measured during in-flight tests by different instruments. In some cases, measurements performed by different groups of researchers during the same flight tests showed an order of magnitude variation. Second priority was assigned to measuring concentrations of odd hydrogen and oxidizing species. Since OH, HO2, H2O2, and O are extremely reactive, non-extractive measurements are recommended. A combination of absorption and fluorescence is anticipated to be effective for OH measurements in the combustor and at the engine exit. Extractive measurements of HO2 have been made in the stratosphere, where the ambient level of OH is relatively low. Use of techniques that convert HO2 to OH for combustor and engine exit measurements needs to be evaluated, since the ratio of HO2/OH may be 1% or less at both the combustor and engine exit. CI-MS might be a viable option for H2O2, subject to sampling line conversion issues. However, H2O2 is a low priority oxidizing species in the combustor and at the engine exit. Two candidates for atomic oxygen measurements are Resonance Enhanced Multi-Photon Ionization (REMPI) and Laser-Induced Fluorescence (LIF). Particulate measurement by simultaneous extractive and non-extractive techniques was given equal priority to the oxidizer measurements. Concern was expressed over the ability of typical ground test sampling lines to deliver an unaltered sample to a remotely located instrument. It was suggested that the sampling probe and line losses be checked out by attempting measurements using an optical or non-extractive technique immediately upstream of the sampling probe. This is a possible application for Laser Induced Incandescence (LII) as a check on the volume fraction of soot. Optical measurements of size distribution are not well developed for ultrafine particles less than about 20 nm in diameter, so a non-extractive technique for particulate size distribution cannot be recommended without further development. Carbon dioxide measurements need to be made to complement other extractive measurement techniques. CO2 measurements enable conversion of other species concentrations to emission indices. Carbon monoxide, which acts as a sink for oxidizing species, should be measured using non-extractive techniques. CO can be rapidly converted to CO2 in extractive probes, and a comparison between extractive and non-extractive measurements should be performed. Development of non-extractive techniques would help to assess the degree of CO conversion, and might be needed to improve the concentration measurement accuracy. Measurements of NO(x) will continue to be critical due to the role of NO and NO2 in atmospheric chemistry, and their influence on atmospheric ozone. Time-resolved measurements of temperature, velocity, and species concentrations were included on the list of desired measurement. Thermocouples are typically adequate for engine exit measurements. PIV and LDV are well established for obtaining velocity profiles. The techniques are listed in the accompanying table; are divided into extractive and non-extractive techniques. Efforts were made to include a measurement uncertainty for each technique. An assessment of the technology readiness was included.

Zaller, Michelle↗

Technology requirements and readiness for very large vehicles

Common concerns of very large vehicles in the areas of economics, transportation system interfaces and operational problems were reviewed regarding their influence on vehicle configurations and technology. Fifty-four technology requirements were identified which are judged to be unique, or particularly critical, to very large vehicles. The requirements were about equally divided among the four general areas of aero/hydrodynamics, propulsion and acoustics, structures, and vehicle systems and operations. The state of technology readiness was judged to be poor to fair for slightly more than one half of the requirements. In the classic disciplinary areas, the state of technology readiness appears to be more advanced than for vehicle systems and operations.

Conner, D. W.↗