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22 N HPGP Thruster Life Testing

In the ever-changing paradigm of efficient and capable spacecraft design, scientific missions continue pushing the envelope enabling spacecraft subsystems to deliver effective solutions to meet challenging new mission/spacecraft needs. From an in-space storable liquid chemical propulsion perspective, monopropellant hydrazine has been, and continues to be, a dependable propellant with considerable flight heritage, a variety of engine thrust classes available from multiple vendors, with repeatable and reliable performance. Additionally, the space propulsion industry has learned to successfully handle hydrazine, its regulations, the safety protocols, the personnel protective equipment, and the unique training standards–all requisite for loading spacecraft propulsion systems with toxic hypergolic hydrazine. The question now arises as to “what is next for in-space chemical propulsion?” Further, with the evolution and concrete advancements in innovative in-space green propellant technologies, capable of providing realizable benefits to scientific missions, concern over the reliability and availability of this higher performing and safer to handle class of propellants is waning. As science missions move forward with the potential flight in fusion of High Performance Green Propulsion (HPGP), NASA and its industry partners are working to address any gaps in system reliability, performance, or unique operational considerations. Propellant technology that offers both higher performance and significant reduction in personnel hazards compared to hydrazine presents an attractive propulsion subsystem design opportunity. Increased propulsion subsystem performance can result in lower spacecraft launch mass, larger scientific payloads, or extended on-orbit lifetimes. Mission trades using green propulsion technologies have been documented on multiple NASA Goddard Space Flight Center (GSFC) mission classes, examining various parameters and requirements to support mission architectures in Low Earth Orbit (LEO), High Earth Orbit (HEO), geostationary, lunar, planetary, and Quasi-halo orbit around Sun-Earth Lagrange point (L2). The results of these trade studies show promising, attainable benefits. The perceived programmatic risk of flying a newer propulsion technology has, unfortunately, not outweighed the benefits to date. To take advantage of the improved performance and mitigate programmatic risk, HPGP engines must demonstrate life testing at higher propellant throughputs than have currently been demonstrated. In an effort to proactively address the challenges with technology infusion into a risk-averse community, NASA and the Swedish National Space Agency (SNSA) outlined a collaborative Implementing Arrangement (IA) for the respective agencies to pursue increased HPGP technology maturation. This initial IA effort began in 2013, fresh off the heels of the successful PRISMA HPGP technology demonstration mission. The IA targeted objective is to reduce risk to potential future HPGP missions and fully characterize the LMP-103S propellant and associated engine performance. Over the past eight years, HPGP has flown in propulsion systems on twenty-five(25) spacecraft from seven(7) different Launch Ranges around the globe and on seven (7) different Launch Vehicles. Six(6) of these launches involved multiple loading operations for multiple spacecraft. For U.S. Range operations, nine (9) HPGP systems have been processed at Vandenberg Space Force Base(VSFB):six(6) in 2017, and three (3) in 2018. Six (6) more have been processed at Cape Canaveral Air Force Station (CCAFS)in May 2020, with three (3) systems launched in June 2020 and the remaining three (3) system were left loaded and ready until their launch in August of 2020. Three (3) more systems have been processed at Wallops Flight Facility(WFF)and launched in June 2021. In addition, these propulsion subsystems employed heritage propulsion subsystem component such as valves, filters, and pressure transducers, and have further demonstrated nominal functionality in both diaphragm and Propellant Management Device (PDM) propellant tanks. Based on these successes, HPGP technology continues to be considered for NASA Science Mission Directorate missions at GSFC. The work presented herein represents many years of development and collaborative efforts to successfully align higher performance, low toxicity hydrazine alternatives into scientific missions. NASA GSFC Propulsion Engineering, in collaboration with Bradford ECAPS, has developed mission specific thruster design and testing requirements to establish GSFC’s desired test conditions and firing sequences.In2017, the first flight-like 22N HPGP thruster Engineering Qualification Model (EQM-1)was designed and built by Bradford ECAPS to prove out the thruster design, materials, build process, and test campaign with respect to NASA GSFC critical component and mission requirements. This test program was developed to comprehensively test the thruster, the technology, and ultimately increase the 22N HPGP Technology Readiness Level(TRL). EQM-1was tested to environmental qualification levels prior to hot fire performance testing to represent the relevant end-to-end environment (launch to on-orbit operation)with required margin. This thruster demonstrated steady-state and pulse mode operational capability with propellant thruster throughput up to~53kg.At this throughput level, the EQM-1 engine began to present off-nominal performance and the test campaign was halted to allow for non-destructive testing and identify the root cause for the an omalous performance. Capitalizing on the successful elements of the EQM-1 campaign, an upgraded 22N HPGP EQM-2 has been manufactured by Bradford ECAPS to meet the complete GSFC requirements. The EQM-2 thruster’s test campaign has further demonstrated the robustness of the HPGP propulsion technology and increased the Technology Readiness Level (TRL) by undergoing a full acceptance test program, then proceeding into qualification, including environmental testing (vibration and shock to qualification levels),as well as hot-fire life testing, operating at steady-state and pulse modes with increased propellant thruster throughput to~150kg. The HPGP thruster performance testing enables effective HPGP thruster readiness evaluation to meet NASA candidate mission requirements in the future.

High↗

DOE OSTI2021

Testing of an Ultrasonic Thermometer Developed at the Idaho National Laboratory

Ultrasonic thermometry has great potential to improve temperature measurements in irradiation experiments performed in nuclear test reactors. The flexibility of the measurement makes it applicable to a wide range of potential test conditions, in terms of temperature, radiation, and chemical environment. However, the range of several typical ultrasonic thermometer configurations are not well defined. The research described in this document attempts to define an operational envelop for multiple configurations and sensor materials.

46 - INSTRUMENTATION RELATED TO NUCLEAR SCIENCE AN↗

DOE OSTI2025

Updating the Building Science Advisor (BSA): A Tool to Assist in the Design of Durable Building Envelopes

Predicting the moisture durability of building envelope components remains challenging due to multiple influencing factors, including material selection, assembly positioning, local climate conditions, air tightness, interior environment, and construction quality. Building codes increasingly emphasize energy efficiency through enhanced insulation and tighter envelopes but offer limited guidance on moisture durability considerations. Consequently, builders face uncertainty, particularly as new materials and assemblies enter the market.The Building Science Advisor (BSA) is a free, web-based expert system developed to address these challenges by providing actionable insights into the moisture durability and energy efficiency of both new and retrofit wall designs. Recently updated, we are now providing version 3.0 of the tool. BSA features significant user interface improvements, enhancing navigation and user interaction through a refreshed, intuitive design. Additionally, the tool incorporates a newly developed database containing pre-simulated wall assembly cases, significantly reducing response times and improving the accuracy of moisture durability assessments. Furthermore, the updated BSA includes moisture content as a performance criterion, providing users with a more comprehensive understanding of moisture-related durability risks. These enhancements enable rapid, reliable assessments tailored to specific climate zones and local building practices. BSA continues to offer targeted guidance on wall retrofit scenarios and delivers access to an expanded library of location-specific building science resources.This paper describes these key updates, highlighting the enhanced features, expanded capabilities, and overall improvements to user experience and educational content. The paper includes a demonstration that illustrates how the revised BSA effectively supports practitioners in designing durable, energy-efficient building envelope assemblies.

Salonvaara, Mikael [ORNL] (ORCID:0000000318991554)↗

NASA NTRS2010

Orion Entry Performance-Based Center-of-Gravity Box

The Orion capsule is designed both for Low Earth Orbit missions to the ISS and for missions to the moon. For ISS class missions, the capsule will use an Apollo-style direct entry. For lunar return missions, depending on the timing of the mission, the capsule could perform a direct entry or a skip entry of up to 4800 n.mi. in order to land in the coastal waters of California. The physics of atmospheric re-entry determine the capability of the Orion vehicle. For a given vehicle mass and shape, physics tells us that the driving parameters for an entry vehicle are the hypersonic lift-to-drag ratio (L/D) and the flight path angle at entry interface (gamma(sub EI)). The design of the Orion atmospheric re-entry must meet constraints during both nominal and dispersed flight conditions on landing accuracy, heating rate, total heat load, sensed acceleration, and proper disposal of the Service Module. These constraints define an entry corridor in the space of L/D-gamma(sub EI); if the vehicle falls within this corridor, then all constraints are met. The gamma(sub EI) dimension of the corridor can be further constrained by the gloads experienced during emergency entries. Thus, the entry performance for the Orion vehicle can be described completely by the L/D. Bounds on the hypersonic L/D necessary to achieve all the mission requirements can be defined for the given entry corridor. Landing accuracy performance drives the lower limit on L/D. In order to achieve the desired landing accuracy, a minimum L/D must be ensured. The design of the Thermal Protection System (TPS) drives the upper limit on L/D. A higher L/D can drive mass into the design of the TPS. Conversely, once the TPS is designed, the L/D must be ensured to stay below a certain limit in order for the TPS to stay within its design envelop. The L/D must stay within its upper and lower bounds during dispersed flight conditions. L/D is a function of both the aerodynamics and the center-of-gravity (CG) of the vehicle. The aerodynamics of the vehicle are determined by Computational Fluid Mechanics (CFD) and wind tunnel tests. However, the aerodynamics are not known precisely. Instead, an aerodynamic database has been developed where the aerodynamic coefficients are known to fall within a probabilistic band defined by upper and lower bounds. It is expected that the probabilistic band will shrink after the first missions are flown and real-world data is collected. Until that time, the Orion must be designed to the current aerodynamic database. Thus, for a given aerodynamic database with given uncertainties, the allowable range in L/D can be mapped to an allowable box for the CG location. The CG box is used to set requirements on the dispersions allowed for vehicle packaging and cargo storage. As the aerodynamic uncertainties decrease, the size of the CG box can increase. This paper discusses the technique used to map the minimum and maximum L/D bounds set by the entry performance requirements to the allowable dispersions in CG while accounting for aerodynamic uncertainties. The L/D is defined as the ratio of the lift force to the drag force. It is equivalent to the ratio of lift coefficient (C(sub L)) over drag coefficient (C(sub D)). C(sub L) and C(sub D) are functions of Mach number (M) and angle of attack (alpha). A Mach number of 25 is used as a measuring point of the hypersonic L/D. Variations in C(sub L), C(sub D) and alpha cause variations in L/D. Equation (1) shows the three contributions to the variation in L/D.

Rea, Jeremy R.↗

DOE OSTI2014

R&D for Safety Codes and Standards: Materials and Components Compatibility

A principal challenge to the widespread adoption of hydrogen infrastructure is the lack of quantifiable data on its safety envelope and concerns about additional risk from hydrogen. To convince regulatory officials, local fire marshals, fuel suppliers, and the public at large that hydrogen refueling is safe for consumer use, the risk to personnel and bystanders must be quantified and minimized to an acceptable level. Such a task requires strong confidence in the safety performance of high pressure hydrogen systems. Developing meaningful materials characterization and qualification methodologies in addition to enhancing understanding of performance of materials is critical to eliminating barriers to the development of safe, low-cost, high-performance high-pressure hydrogen systems for the consumer environment.

36 MATERIALS SCIENCE↗

DOE OSTI2016

PSR B0329+54: STATISTICS OF SUBSTRUCTURE DISCOVERED WITHIN THE SCATTERING DISK ON RADIOASTRON BASELINES OF UP TO 235,000 km

We discovered fine-scale structure within the scattering disk of PSR B0329+54 in observations with the RadioAstron ground–space radio interferometer. Here we describe this phenomenon, characterize it with averages and correlation functions, and interpret it as the result of decorrelation of the impulse-response function of interstellar scattering between the widely separated antennas. This instrument included the 10 m Space Radio Telescope, the 110 m Green Bank Telescope, the 14 × 25 m Westerbork Synthesis Radio Telescope, and the 64 m Kalyazin Radio Telescope. The observations were performed at 324 MHz on baselines of up to 235,000 km in 2012 November and 2014 January. In the delay domain, on long baselines the interferometric visibility consists of many discrete spikes within a limited range of delays. On short baselines it consists of a sharp spike surrounded by lower spikes. The average envelope of correlations of the visibility function shows two exponential scales, with characteristic delays of τ{sub 1}=4.1±0.3 μs and τ{sub 2}=23±3 μs, indicating the presence of two scales of scattering in the interstellar medium. These two scales are present in the pulse-broadening function. The longer scale contains 0.38 times the scattered power of the shorter one. We suggest that the longer tail arises from highly scattered paths, possibly from anisotropic scattering or from substructure at large angles.

79 ASTRONOMY AND ASTROPHYSICS↗

NASA NTRS1986

Photovoltaic array loss mechanisms

Loss mechanisms which come into play when solar cell modules are mounted in arrays are identified. Losses can occur either from a reduction in the array electrical performance or with nonoptimal extraction of power from the array. Electrical performance degradation is caused by electrical mismatch, transmission losses from cell surface soiling and steep angle of reflectance, and electrical losses from field wiring resistance and the voltage drop across blocking diodes. The second type of loss, concerned with the operating points of the array, can involve nonoptimal load impedance and limiting the operating envelope of the array to specific ranges of voltage and current. Each of the loss mechanisms are discussed and average energy losses expected from soiling, steep reflectance angles and circuit losses are calculated.

Gonzalez, Charles↗

NASA NTRS2005

Reconfiguration of Analog Electronics for Extreme Environments

This paper argues in favor of adaptive reconfiguration as a technique to expand the operational envelope of analog electronics for extreme environments (EE). On a reconfigurable device, although component parameters change in EE, as long as devices still operate, albeit degraded, a new circuit design, suitable for new parameter values, may be mapped into the reconfigurable structure to recover the initial circuit function. Laboratory demonstrations of this technique were performed by JPL in several independent experiments in which bulk CMOS reconfgurable devices were exposed to, and degraded by, high temperatures (approx.300 C) or radiation (300kRad TID), and then recovered by adaptive reconfiguration using evolutionary search algorithms.

evolutionary search algorithms↗

NASA NTRS2011

Relaxing USOS Solar Array Constraints for Russian Vehicle Undocking

With the retirement of Space Shuttle cargo delivery capability and the ten year life extension of the International Space Station (ISS) more emphasis is being put on preservation of the service life of ISS critical components. Current restrictions on the United States Orbital Segment (USOS) Solar Array (SA) positioning during Russian Vehicle (RV) departure from ISS nadir and zenith ports cause SA to be positioned in the plume field of Service Module thrusters and lead to degradation of SAs as well as potential damage to Sun tracking Beta Gimbal Assemblies (BGA). These restrictions are imposed because of the single fault tolerant RV Motion Control System (MCS), which does not meet ISS Safety requirements for catastrophic hazards and dictates 16 degree Solar Array Rotary Joint position, which ensures that ISS and RV relative motion post separation, does lead to collision. The purpose of this paper is to describe a methodology and the analysis that was performed to determine relative motion trajectories of the ISS and separating RV for nominal and contingency cases. Analysis was performed in three phases that included ISS free drift prior to Visiting Vehicle separation, ISS and Visiting Vehicle relative motion analysis and clearance analysis. First, the ISS free drift analysis determined the worst case attitude and attitude rate excursions prior to RV separation based on a series of different configurations and mass properties. Next, the relative motion analysis calculated the separation trajectories while varying the initial conditions, such as docking mechanism performance, Visiting Vehicle MCS failure, departure port location, ISS attitude and attitude rates at the time of separation, etc. The analysis employed both orbital mechanics and rigid body rotation calculations while accounting for various atmospheric conditions and gravity gradient effects. The resulting relative motion trajectories were then used to determine the worst case separation envelopes during the clearance analysis. Analytical models were developed individually for each stage and the results were used to build initial conditions for the following stages. In addition to the analysis approach, this paper also discusses the analysis results, showing worst case relative motion envelopes, the recommendations for ISS appendage positioning and the suggested approach for future analyses.

Menkin, Evgeny↗

NASA NTRS2015

NASA Glenn Icing Research Tunnel: 2014 and 2015 Cloud Calibration Procedures and Results

This report summarizes the current status of the NASA Glenn Research Center (GRC) Icing Research Tunnel cloud calibration: specifically, the cloud uniformity, liquid water content, and drop-size calibration results from both the January-February 2014 full cloud calibration and the January 2015 interim cloud calibration. Some aspects of the cloud have remained the same as what was reported for the 2014 full calibration, including the cloud uniformity from the Standard nozzles, the drop-size equations for Standard and Mod1 nozzles, and the liquid water content for large-drop conditions. Overall, the tests performed in January 2015 showed good repeatability to 2014, but there is new information to report as well. There have been minor updates to the Mod1 cloud uniformity on the north side of the test section. Also, successful testing with the OAP-230Y has allowed the IRT to re-expand its operating envelopes for large-drop conditions to a maximum median volumetric diameter of 270 microns. Lastly, improvements to the collection-efficiency correction for the SEA multi-wire have resulted in new calibration equations for Standard- and Mod1-nozzle liquid water content.

aircraft icing↗

NASA NTRS2016

System Design and Implementation of the Detector Assembly of the Astro-H Soft X-Ray Spectrometer

The soft x-ray spectrometer (SXS) onboard Astro-H presents to the science community unprecedented capability (less than 7 eV at 6 keV) for high-resolution spectral measurements in the range of 0.5-12 keV to study extended celestial sources. At the heart of this SXS is the x-ray calorimeter spectrometer (XCS) where detectors (calorimeter array and anticoincidence detector) operate at 50 mK, the bias circuit operates at nominal 1.3 K, and the first stage amplifiers operateat 130 K, all within a nominal 20 cm envelope. The design of the detector assembly in this XCS originates from the Astro-E x-ray spectrometer (XRS) and lessons learned from Astro-E and Suzaku. After the production of our engineering model, additional changes were made in order to improve our flight assembly process for better reliability and overall performance. In this poster, we present the final design and implementation of the flight detector assembly, show comparison of parameters and performance to Suzakus XRS, and list susceptibilities to other subsystems as well as our lessons learned.

soft x-ray spectrometer↗

DOE OSTI2022

Analysis of sparse recovery for Legendre expansions using envelope bound

We provide novel sufficient conditions for the uniform recovery of sparse Legendre expansions using ℓ 1 minimization, where the sampling points are drawn according to orthogonalization (uniform) measure. So far, conditions of the form m ≳ Θ 2 s x log factors have been relied on to determine the minimum number of samples m that guarantees successful reconstruction of s-sparse vectors when the measurement matrix is associated to an orthonormal system. However, in case of sparse Legendre expansions, the uniform bound Θ of Legendre systems is so high that these conditions are unable to provide meaningful guarantees. Here, in this paper, we present an analysis which employs the envelop bound of all Legendre polynomials instead, and prove a new recovery guarantee for s-sparse Legendre expansions, m ≳ Θs 2 x log factors, which is independent of Θ. Arguably, this is the first recovery condition established for orthonormal systems without assuming the uniform boundedness of the sampling matrix. The key ingredient of our analysis is an extension of chaining arguments, recently developed in Bourgain and Chkifa et al., to handle the envelope bound. Furthermore, our recovery condition is proved via restricted eigenvalue property, a less demanding replacement of restricted isometry property which is perfectly suited to the considered scenario. Along the way, we derive simple criteria to detect good sample sets. Our numerical tests show that sets of uniformly sampled points that meet these criteria will perform better recovery on average.

97 MATHEMATICS AND COMPUTING↗

DOE OSTI2018

Solar Hot Air Balloons for Terrestrial and Planetary Atmospheres

The first solar hot air balloon was constructed in the early 1970s (Besset, 2016). Over the following decades the Centre National d'Etudes Spatiales (CNES) developed the Montgolfiere Infrarouge (MIR) balloon, which flew on solar power during the day and infrared radiation from the Earth's surface at night (Fommerau and Rougeron, 2011). The balloons were capable of flying for over 60 days and apparently reached altitudes of 30 km at least once (Malaterre, 1993). Solar balloons were the subject of a Jet Propulsion Laboratory study that performed test flights on Earth (Jones and Wu 1999) and discussed their mission potential for Mars, Jupiter, and Venus (Jones and Heun, 1997). The solar balloons were deployed from the ground and dropped from hot air balloons; some were altitude controlled by means of a remotely-commanded air valve at the top of the envelope. More recently, solar balloons have been employed for infrasound studies in the lower stratosphere (see Table 1). The program began in 2015, when a prototype balloon reached an altitude of 22 kilometers before terminating just prior to float (Bowman et al., 2015). An infrasound sensor was successfully deployed on a solar balloon during the 2016 SISE/USIE experiment, in which an acoustic signal from a ground explosion was captured at a range of 330 km (Anderson et al. 2018; Young et al. 2018). This led to the launch of a 5-balloon infrasound network during the Heliotrope experiment (Bowman and Albert, 2018). The balloons were constructed by the researchers themselves at a materials of less than $50 per envelope.

14 SOLAR ENERGY↗

DOE OSTI2022

Impact of Mass Wood Walls on Building Energy Use, Peak Demand, and Thermal Comfort

Thermal mass moderates indoor temperatures, allowing the heating, ventilation, and air conditioning (HVAC) system to operate more efficiently during peak hours. Cross-Laminated Timber (CLT) and other mass wood products provide thermal mass to the building envelope in a lighter and more environmentally friendly form than concrete. However, the influence of CLTs on heating and cooling energy, peak energy demand, and the indoor climate is not well-known. This project was formed to find the energy efficiency benefits of mass timber structures. The objectives were to (1) test the thermal performance of insulated mass wall structures in controlled laboratory conditions, (2) validate simulation models to expand performance to natural climatic conditions, (3) isolate the impact of the thermal properties of wood (thermal mass, thermal conductivity, moisture storage) on heating and cooling energy use and peak energy demand, and to optimize the performance. Currently, the designs used for CLT buildings have focused on structural and fire issues, and they do not consider the thermal mass benefits. Therefore, the thermal performance is not necessarily optimized.

32 ENERGY CONSERVATION, CONSUMPTION, AND UTILIZATI↗

NASA NTRS1981

Report on research and technology-FY 1981

More than 65 technical reports, papers, and articles published by personnel and contractors at the Dryden Flight Research Center are listed. Activities performed for the Offices of Aeronautics and Space Technology, Space and Terrestrial Applications, Space Transportation Systems, and Space Tracking and Data Systems are summarized. Preliminary stability and control derivatives were determined for the shuttle orbiter at hypersonic speeds from the data obtained at reentry. The shuttle tile tests, spin research vehicle nose shapes flight investigations, envelope expansion flights for the Ames tilt rotor research aircraft, and the AD-1 oblique wing programs were completed as well as the KC-135 winglet program.

Source record↗

NASA NTRS1982

Rate 3/4 16-PSK phase codes

An investigation is carried out of convolutional-coding coupled with 16-PSK modulation for bandwidth efficient transmission. Small memory codes (rate three-fourths) are found that are optimum in the free-distance sense on the Gaussian channel. These codes furnish up to 4.8 dB of coding gain with 32 states over uncoded 8-PSK, a scheme having the same spectral efficiency as the codes described. The performance is compared with the earlier findings of Ungerboeck (1976, 1982) and certain recent results on R = 2/3 coded 8-PSK. It is shown that with 4 decoder states the coding gain is 3.5 dB and with 32 states it is 4.8 dB. What is more, the codes compare favorably with coded 16-QASK when envelope variations and peak energy constraints are considered.

Wilson, S. G.↗

NASA NTRS1996

Development of High Speed Imaging and Analysis Techniques Compressible Dynamics Stall

Dynamic stall has limited the flight envelope of helicopters for many years. The problem has been studied in the laboratory as well as in flight, but most research, even in the laboratory, has been restricted to surface measurement techniques such as pressure transducers or skin friction gauges, except at low speed. From this research, it became apparent that flow visualization tests performed at Mach numbers representing actual flight conditions were needed if the complex physics associated with dynamic stall was to be properly understood. However, visualization of the flow field during compressible conditions required carefully aligned and meticulously reconstructed holographic interferometry. As part of a long-range effort focused on exposing of the physics of compressible dynamic stall, a research wind tunnel was developed at NASA Ames Research Center which permits visual access to the full flow field surrounding an oscillating airfoil during compressible dynamic stall. Initially, a stroboscopic schlieren technique was used for visualization of the stall process, but the primary research tool has been point diffraction interferometry(PDI), a technique carefully optimized for use in th is project. A review of the process of development of PDI will be presented in the full paper. One of the most valuable aspects of PDI is the fact that interferograms are produced in real time on a continuous basis. The use of a rapidly-pulsed laser makes this practical; a discussion of this approach will be presented in the full paper. This rapid pulsing(up to 40,000 pulses/sec) produces interferograms of the rapidly developing dynamic stall field in sufficient resolution(both in space and time) that the fluid physics of the compressible dynamic stall flowfield can be quantitatively determined, including the gradients of pressure in space and time. This permits analysis of the influence of the effect of pitch rate, Mach number, Reynolds number, amplitude of oscillation, and other parameters on the dynamic stall process. When interferograms can be captured in real time, the potential for real-time mapping of a developing unsteady flow such as dynamic stall becomes a possibility. This has been achieved in the present case through the use of a high-speed drum camera combined with electronic circuitry which has resulted in a series of interferograms obtained during a single cycle of dynamic stall; images obtained at the rate of 20 KHz will be presented as a part of the formal presentation. Interferometry has been available for a long time; however, most of its use has been limited to visualization. The present research has focused on use of interferograms for quantitative mapping of the flow over oscillating airfoils. Instantaneous pressure distributions can now be obtained semi-automatically, making practical the analysis of the thousands of interferograms that are produced in this research. A review of the techniques that have been developed as part of this research effort will be presented in the final paper.

Chandrasekhara, M. S.↗

NASA NTRS2011

Piloted Simulation to Evaluate the Utility of a Real Time Envelope Protection System for Mitigating In-Flight Icing Hazards

The utility of the Icing Contamination Envelope Protection (ICEPro) system for mitigating a potentially hazardous icing condition was evaluated by 29 pilots using the NASA Ice Contamination Effects Flight Training Device (ICEFTD). ICEPro provides real time envelope protection cues and alerting messages on pilot displays. The pilots participating in this test were divided into two groups; a control group using baseline displays without ICEPro, and an experimental group using ICEPro driven display cueing. Each group flew identical precision approach and missed approach procedures with a simulated failure case icing condition. Pilot performance, workload, and survey questionnaires were collected for both groups of pilots. Results showed that real time assessment cues were effective in reducing the number of potentially hazardous upset events and in lessening exposure to loss of control following an incipient upset condition. Pilot workload with the added ICEPro displays was not measurably affected, but pilot opinion surveys showed that real time cueing greatly improved their situation awareness of a hazardous aircraft state.

Ranaudo, Richard J.↗