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NASA NTRS1998

Operational processing and cloud boundary detection from micro pulse lidar data

Micro Pulse Lidar (MPL) was developed at NASA Goddard Space Flight Center (GSFC) as the result of research on space-borne lidar techniques. It was designed to provide continuous, unattended observations of all significant atmospheric cloud and aerosol structure with a rugged, compact system design and the benefit of eye safety (Spinhirne 1993). The significant eye safety feature is achieved by using low pulse energies and high pulse repetition rates compared to standard lidar systems. MPL systems use a diode pumped 10 microj, 2500 Hz doubled Nd:YLF laser. In addition, a solid state Geiger mode avalanche photo diode (GAPD) photon counting detector is used allowing for quantum efficiencies approaching 70%. Other design features have previously been noted by Spinhirne (1995). Though a commercially available instrument, with nearly 20 systems operating around the world, the most extensive MPL work has come from those operated by the Atmospheric Radiation Measurement (ARM) (Stokes and Schwartz 1994) program. The diverse ability of the instrument relating to the measurement of basic cloud macrophysical structure and both cloud and aerosol radiative properties well suits the ARM research philosophy. MPL data can be used to yield many parameters including cloud boundary heights to the limit of signal attenuation, cloud scattering cross sections and optical thicknesses, planetary boundary layer heights and aerosol scattering profiles, including those into the stratosphere in nighttime cases (Hlavka et al 1996). System vertical resolution ranges from 30 m to 300 m (i.e. high and low resolution respectively) depending on system design. The lidar research group at GSFC plays an advisory role in the operation, calibration and maintenance of NASA and ARM owned MPL systems. Over the past three years, processing software and system correction techniques have been developed in anticipation of the increasing population of systems amongst the community. Datasets produced by three ARM-owned systems have served as the basis for this development. With two operating at the southern Great Plains Cloud and Radiation Testbed Site (SGP CART) since December 1993 and another at the Manus Island Atmospheric Radiation and Cloud Station (TWP ARCS) location in the tropical western Pacific since February 1997, the ARM archive contains over 4 years of observations. In addition, high resolution systems planning to come on-line at the North Slope, AK CART shortly with another scheduled to follow at the TWP ARCS-II will diversify this archive with more extensive observations.

Campbell, James R.↗

NASA NTRS2017

The Integrated Medical Model: Outcomes from Independent Review

In 2016, the Integrated Medical Model (IMM) v4.0 underwent an extensive external review in preparation for transition to an operational status. In order to insure impartiality of the review process, the Exploration Medical Capabilities Element of NASA's Human Research Program convened the review through the Systems Review Office at NASA Goddard Space Flight Center (GSFC). The review board convened by GSFC consisted of persons from both NASA and academia with expertise in the fields of statistics, epidemiology, modeling, software development, aerospace medicine, and project management (see Figure 1). The board reviewed software and code standards, as well as evidence pedigree associated with both the input and outcomes information. The board also assesses the models verification, validation, sensitivity to parameters and ability to answer operational questions. This talk will discuss the processes for designing the review, how the review progressed and the findings from the board, as well as summarize the IMM project responses to those findings. Overall, the board found that the IMM is scientifically sound, represents a necessary, comprehensive approach to identifying medical and environmental risks facing astronauts in long duration missions and is an excellent tool for communication between engineers and physicians. The board also found IMM and its customer(s) should convene an additional review of the IMM data sources and to develop a sustainable approach to augment, peer review, and maintain the information utilized in the IMM. The board found this is critically important because medical knowledge continues to evolve. Delivery of IMM v4.0 to the Crew Health and Safety (CHS) Program will occur in the 2017. Once delivered for operational decision support, IMM v4.0 will provide CHS with additional quantitative capability in to assess astronaut medical risks and required medical capabilities to help drive down overall mission risks.

medical equipment↗

NASA NTRSDate not supplied

Isolator Shock Train Response to Dual-Mode Scramjet Throttling-Like Forcing

Typical dual-mode scramjet designs employ an isolator component situated between the inlet and combustor. The isolator fluidically buffers the inlet from the combustor via the formation of a shock train, an interconnected series of shock-boundary layer interactions. The isolator is required during ramjet-engine operation and at the lower edges of the scramjet operational envelope. In an operational engine, transient behavior imposes fluidic dynamics due to attitude adjustments, combustor throttling, aberrant resonances within the combustor, etc. To accommodate these transients, the shock train adjusts on both temporal and spatial scales. Sufficiently large downstream pressure increases in the combustor play a critical role during unstart, whereby strong compression waves propagate upstream in the isolator to ultimately force the shock train out the inlet, if it cannot stabilize to the pressure increase. This critically reduces the engine mass flow rate, potentially causing a catastrophic reduction in thrust and loss of the vehicle. A nearly instantaneous unstart can, in principle, be predicted by a ratio of the normal shock pressure jump in an inviscid isolator. However, lower-magnitude perturbations, below this bound, also induce unstarts. Understanding the transient behavior inherent in unsteady isolator operation is critical to reducing safety margin in isolator length and refining engine control algorithms to improve engine performance. The Isolator Dynamics Research Lab (IDRL) is a cold-flow, direct-connect isolator facility ideal for assessing the fundamental fluid dynamics of shock trains. The internal flow of a scramjet isolator is modeled using a 2D, symmetric Mach 2.5 nozzle followed by a 695 mm long, 25.4x50.8 mm rectangular isolator test section. A blockage cone on a high-speed ball screw actuator downstream of the isolator is pulsed to simulate the throttling behavior in an operational combustor. Wall flush high-speed static pressure transducers detect the streamwise shock train within the isolator. Time delays between the onset of the throttling and the response motion of the shock train are analyzed to assess the controlling scales and parameters. From this analysis, the response time of the isolator shock train system is found to depend primarily on the isolator back pressure time-rate-of-change but quickly saturates to a maximum propagation speed along the isolator. These results inform dual-mode scramjet engine designers of the critical parameters and limits of an unstart resistant engine.

scramjet↗

NASA NTRS2005

The Apollo Capsule Optimization for Improved Stability and Computational/Experimental Data Comparisons

Numerical optimization was employed on the Apollo Command Module to modify its external shape. The Apollo Command Module (CM) that was used on all NASA human space flights during the Apollo Space Program is stable and trimmed in an apex forward (alpha of approximately 40 to 80 degrees) position. This poses a safety risk if the CM separates from the launch tower during abort. Optimization was employed on the Apollo CM to remedy the undesirable stability characteristics of the configuration. Geometric shape changes were limited to axisymmetric modifications that altered the radius of the apex (R(sub A)), base radius (R(sub O)), corner radius (R(sub C)), and the cone half angle (theta), while the maximum diameter of the CM was held constant. The results of multipoint optimization on the CM indicated that the cross-range performance can be improved while maintaining robust apex-aft stability with a single trim point. Navier-Stokes computations were performed on the baseline and optimized configurations and confirmed the Euler-based optimization results. Euler Analysis of ten alternative CM vehicles with different values of the above four parameters are compared with the published experimental results of numerous wind tunnel tests during the late 1960's. These comparisons cover a wide Mach number range and a full 180-degree pitch range and show that the Euler methods are capable of fairly accurate force and moment computations and can separate the vehicle characteristics of these ten alternative configurations.

Cliff, Susan E.↗

NASA NTRS2006

Comanche Helmet-Mounted Display Symbology Simulation

The Aeroflightdynamics Directorate (AMRDEC) conducted a simulation to examine the performance of the Comanche Contact Analog world-referenced symbology displayed on the Comanche HIDSS when compared with a compressed symbology design similar to that specified by the former MIL-STD 1295. Six experimental test pilots flew one modified ADS-33 maneuver (hover turn, bob-up), an unusual attitude recovery, and two terrain flight tactical tasks in the NASA Vertical Motion Simulator (VMS). Analysis of the pilot objective performance data and subjective data showed the following results. Objective test results showed that 1295 symbology yielded more rapid maneuvering in the hover turn bob-up than Contact Analog symbology. The average margin of difference in the time to complete the maneuver was approximately two seconds, which was statistically significant. There were no significant differences measured between symbology sets with respect to altitude or position performance measures for all other maneuvers. The NOE target ID task data showed improved accuracy in determining heading to target when using Contact Analog over MIL-STD-1295. Subjective test results, including handling qualities ratings (HQRs) and NASA-TLX workload ratings, showed small but consistent advantages of 1295 symbology over Contact Analog for most parameters. For the bob-up maneuver, 1295 symbology handling qualities were rated Desired for lateral position error and time to complete whereas Contact Analog was rated adequate. The average HQRs for all other maneuvers were rated the same for both symbology sets. Pilot comments and the results of an online questionnaire more strongly favored 1295 over Contact Analog. Repeated comments from all six pilots led to a focus on design issues with six Contact Analog symbols. Those symbols were the heading tape, horizon line, radar altitude six-second predictor, the position of the torque symbol, the absence of a hover position cue, and the widespread positioning of the symbology to the outer edges of the display field-of-view. Pilots rated the present design of three of the six symbols as having safety-of-flight implications. Those symbols were the heading tape, horizon line, and six-second predictor. To summarize, test results showed no objective data that would warrant restricting experimental test pilots from flying constrained tasks. Small but consistent advantages were recorded for the MILSTD-1295 symbology design when executing ADS-33 constrained tasks. No general performance differences were recorded for the operational maneuvers (NOE, contour flight modes) except for the azimuth-to-target task, which favored the Contact Analog heading tape design. There were consistent and strong pilot comments supporting the MIL-STD-1295 design over Contact Analog in both this current simulation and in Comanche Sim I. Three Contact Analog symbols warrant modification and further evaluation to mitigate safety-of-flight implications noted by participating test pilots. A plan for symbology redesign and testing was developed. However, the Comanche program was cancelled by Department of Army in February 2004.

Turpin, Terry S.↗

NASA NTRS2015

PLSS 2.5 Fan Design and Development

NASA is building a high fidelity prototype of an advanced portable life support system (PLSS) as part of the Advanced Exploration Systems Program. This new PLSS, designated as PLSS 2.5, will advance component technologies and systems knowledge in order to inform a future flight program. The oxygen ventilation loop of its predecessor, PLSS 2.0, is driven by a centrifugal fan developed using specifications from over five years ago. PLSS technology and system parameters have matured to the point where the existing fan will not perform adequately for the new prototype. In addition, areas of potential improvement have been identified with the existing fan that could be addressed in a new design. As a result, a new fan was designed and tested for the PLSS 2.5. The PLSS 2.5 fan is a derivative of the one used in PLSS 2.0. It uses the same basic non-metallic can around the motor, but with a larger volute and impeller to meet the higher pressure drop requirements of the PLSS 2.5 loop. This allows it to operate at rotational speeds that are matched to rolling element bearings, and which create reasonably low impeller tip speeds. Development of the fan also considered a shrouded impeller design that allows larger clearances for greater oxygen safety and better performance.

Converse, David↗

NASA NTRS2007

Understanding and Counteracting Fatigue in Flight Crews

The materials included in the collection of documents describe the research of the NASA Ames Fatigue Countermeasures Group (FCG), which examines the extent to which fatigue, sleep loss, and circadian disruption affect flight-crew performance. The group was formed in 1980 in response to a Congressional request to examine a possible safety problem of uncertain magnitude due to transmeridian flying and a potential problem due to fatigue in association with various factors found in air-transport operations and was originally called the Fatigue/Jet Lag Program. The goals of the FCG are: (1) the development and evaluation of strategies for mitigating the effects of sleepiness and circadian disruption on pilot performance levels; (2) the identification and evaluation of objective approaches for the prediction of alertness changes in flight crews; and (3) the transfer and application of research results to the operational field via classes, workshops, and safety briefings. Some of the countermeasure approaches that have been identified to be scientifically valid and operationally relevant are brief naps (less than 40 min) in the cockpit seat and 7-min activity breaks, which include postural changes and ambulation. Although a video-based alertness monitor based on slow eyelid closure shows promise in other operational environments, research by the FCG has demonstrated that in its current form at the time of this reporting, it is not feasible to implement it in the cockpit. Efforts also focus on documenting the impact of untreated fatigue on various types of flight operations. For example, the FCG recently completed a major investigation into the effects of ultra-long-range flights (20 continuous hours in duration) on the alertness and performance of pilots in order to establish a baseline set of parameters against which the effectiveness of new ultra-long-range fatigue remedies can be judged.

Mallis, Melissa↗

NASA NTRS2010

Projection Operator: A Step Towards Certification of Adaptive Controllers

One of the major barriers to wider use of adaptive controllers in commercial aviation is the lack of appropriate certification procedures. In order to be certified by the Federal Aviation Administration (FAA), an aircraft controller is expected to meet a set of guidelines on functionality and reliability while not negatively impacting other systems or safety of aircraft operations. Due to their inherent time-variant and non-linear behavior, adaptive controllers cannot be certified via the metrics used for linear conventional controllers, such as gain and phase margin. Projection Operator is a robustness augmentation technique that bounds the output of a non-linear adaptive controller while conforming to the Lyapunov stability rules. It can also be used to limit the control authority of the adaptive component so that the said control authority can be arbitrarily close to that of a linear controller. In this paper we will present the results of applying the Projection Operator to a Model-Reference Adaptive Controller (MRAC), varying the amount of control authority, and comparing controller s performance and stability characteristics with those of a linear controller. We will also show how adjusting Projection Operator parameters can make it easier for the controller to satisfy the certification guidelines by enabling a tradeoff between controller s performance and robustness.

Larchev, Gregory V.↗

NASA NTRS2019

Integrated Refrigeration and Storage of LNG for Compositional Stability

Growing interest in liquefied natural gas (LNG) as a rocket fuel necessitates a greater technical understanding of the compositional changes due to preferential boil-off (or weathering) that occurs during long duration storage. The purity of methane in LNG can range from 90 to 98%, and is subject to preferential boil-off due to its low boiling point compared to other constituents despite the use of high-performance thermal insulation systems. Active heat extraction (i.e. refrigeration) is required to completely eliminate weathering. For future operational safety and reliability, and to better understand the quality and efficiency of the LNG as a cryofuel, a 400-liter Cryostat vessel was designed and constructed to measure the composition and temperatures of the LNG at a number of different liquid levels over long durations. The vessel is the centerpiece of a custom-designed lab-scale integrated refrigeration and storage (IRaS) system employing a pulse tube cryocooler capable of roughly 50 W of lift at 100 K. Instrumentation includes ten temperature sensors mounted on a vertical rake and five liquid sample tubes corresponding to five liquid levels. Two modes of operation are studied. The first is without refrigeration in order to determine a baseline in the change in composition, and to study stratification of the LNG. The second is performed with the cryocooler active to determine the operational parameters of the IRaS system for eliminating the weathering as well as stratification effects in the bulk liquid. The apparatus design and test method, as well as preliminary test results are presented in this paper. As a bonus in cost-saving and operational efficiency, the capability of the IRaS system to provide zero-loss capabilities such as zero boil-off (ZBO) keeping of the LNG and zero-loss filling/transfer operations are also discussed.

Rose, Laura P.↗

NASA NTRS2019

Integrated Refrigeration and Storage of LNG for Compositional Stability

Rose, L.↗

NASA NTRSDate not supplied

Data-Centric Operational Design Domain Characterization for Machine Learning-Based Aeronautical Products

We give for Machine Learning (ML)-based aeronautical products, a first rigorous characterization of Operational Design Domains (ODDs). Unlike in other application sectors (such as self-driving road vehicles) where ODD development is scenario-based, our approach is data-centric: we propose the dimensions along which the parameters that define an ODD can be explicitly captured, using a top-down approach starting from system specifications, and a bottom-up approach starting from detailed ML Model (MLM) designs. Then we give a categorization of the data that ML-based applications can encounter in operation, identifying their system-level relevance and impact. Specifically, we discuss how those data categories are useful to determine: (1) the requirements necessary to drive the design of MLMs; (2) the potential effects on the MLM and higher levels of the system hierarchy; (3) the learning assurance processes that may be needed, and (4) system architectural considerations. We illustrate the underlying concepts with an example of an aircraft flight envelope. The approach in this paper is one of the cornerstones of a future process guidance for development and certification/approval of safety-related aeronautical products implementing Artificial Intelligence (AI), currently being developed through aviation industry-based consensus, jointly by the SAE G-34 Committee for AI in aviation, and EUROCAE WG-114 for AI.

Aeronautical products↗

NASA NTRS2007

Silicon Carbide Gas Sensors for Propulsion Emissions and Safety Applications

Silicon carbide (SiC) based gas sensors have the ability to meet the needs of a range of aerospace propulsion applications including emissions monitoring, leak detection, and hydrazine monitoring. These applications often require sensitive gas detection in a range of environments. An effective sensing approach to meet the needs of these applications is a Schottky diode based on a SiC semiconductor. The primary advantage of using SiC as a semiconductor is its inherent stability and capability to operate at a wide range of temperatures. The complete SiC Schottky diode gas sensing structure includes both the SiC semiconductor and gas sensitive thin film metal layers; reliable operation of the SiC-based gas sensing structure requires good control of the interface between these gas sensitive layers and the SiC. This paper reports on the development of SiC gas sensors. The focus is on two efforts to better control the SiC gas sensitive Schottky diode interface. First, the use of palladium oxide (PdOx) as a barrier layer between the metal and SiC is discussed. Second, the use of atomically flat SiC to provide an improved SiC semiconductor surface for gas sensor element deposition is explored. The use of SiC gas sensors in a multi-parameter detection system is briefly discussed. It is concluded that SiC gas sensors have potential in a range of propulsion system applications, but tailoring of the sensor for each application is necessary.

Hunter, G. W.↗

NASA NTRS1999

Control Design Strategies to Enhance Long-Term Aircraft Structural Integrity

Over the operational lifetime of both military and civil aircraft, structural components are exposed to hundreds of thousands of low-stress repetitive load cycles and less frequent but higher-stress transient loads originating from maneuvering flight and atmospheric gusts. Micro-material imperfections in the structure, such as cracks and debonded laminates, expand and grow in this environment, reducing the structural integrity and shortening the life of the airframe. Extreme costs associated with refurbishment of critical load-bearing structural components in a large fleet, or altogether reinventoring the fleet with newer models, indicate alternative solutions for life extension of the airframe structure are highly desirable. Increased levels of operational safety and reliability are also important factors influencing the desirability of such solutions. One area having significant potential for impacting crack growth/fatigue damage reduction and structural life extension is flight control. To modify the airframe response dynamics arising from command inputs and gust disturbances, feedback loops are routinely applied to vehicles. A dexterous flight control system architecture senses key vehicle motions and generates critical forces/moments at multiple points distributed throughout the airframe to elicit the desired motion characteristics. In principle, these same control loops can be utilized to influence the level of exposure to harmful loads during flight on structural components. Project objectives are to investigate and/or assess the leverage control has on reducing fatigue damage and enhancing long-term structural integrity, without degrading attitude control and trajectory guidance performance levels. In particular, efforts have focused on the effects inner loop control parameters and architectures have on fatigue damage rate. To complete this research, an actively controlled flexible aircraft model and a new state space modeling procedure for crack growth have been utilized. Analysis of the analytical state space model for crack growth revealed the critical mathematical factors, and hence the physical mechanism they represent, that influenced high rates of airframe crack growth. The crack model was then exercised with simple load inputs to uncover and expose key crack growth behavior. To characterize crack growth behavior, both "short-term" laboratory specimen test type inputs and "long-term" operational flight type inputs were considered. Harmonic loading with a single overload revealed typical exponential crack growth behavior until the overload application, after which time the crack growth was retarded for a period of time depending on the overload strength. An optimum overload strength was identified which leads to maximum retardation of crack growth. Harmonic loading with a repeated overload of varying strength and frequency again revealed an optimum overload trait for maximizing growth retardation. The optimum overload strength ratio lies near the range of 2 to 3 with dependency on frequency. Experimental data was found to correlate well with the analytical predictions.

Newman, Brett A.↗

NASA NTRS2008

X-ray Fluorescence Spectroscopy Study of Coating Thickness and Base Metal Composition

For electrical, electronic, and electromechanical (EEE) parts to be approved for space use, they must be able to meet safety standards approved by NASA. A fast, reliable, and precise method is needed to make sure these standards are met. Many EEE parts are coated in gold (Au) and nickel (Ni), and the thickness coating is crucial to a part s performance. A nondestructive method that is efficient in measuring coating thickness is x-ray fluorescence (XRF) spectroscopy. The XRF spectrometer is a machine designed to measure layer thickness and composition of single or multilayered samples. By understanding the limitations in the collection of the data by this method, accurate composition and thickness measurements can be obtained for samples with Au and Ni coatings. To understand the limitations of data found, measurements were taken with the XRF spectrometer and compared to true values of standard reference materials (SRM) that were National Institute of Standards and Technology (NIST) traceable. For every sample, six different parameters were varied to understand measurement error: coating/substrate combination, number of layers, counting interval, collimator size, coating thickness, and test area location. Each measurement was taken in accordance with standards set by the American Society for Testing and Materials (ASTM) International Standard B 568.

Rolin, T. D.↗

NASA NTRS2006

Embedded Sensors for Measuring Surface Regression

The development and evaluation of new hybrid and solid rocket motors requires accurate characterization of the propellant surface regression as a function of key operational parameters. These characteristics establish the propellant flow rate and are prime design drivers affecting the propulsion system geometry, size, and overall performance. There is a similar need for the development of advanced ablative materials, and the use of conventional ablatives exposed to new operational environments. The Miniature Surface Regression Sensor (MSRS) was developed to serve these applications. It is designed to be cast or embedded in the material of interest and regresses along with it. During this process, the resistance of the sensor is related to its instantaneous length, allowing the real-time thickness of the host material to be established. The time derivative of this data reveals the instantaneous surface regression rate. The MSRS could also be adapted to perform similar measurements for a variety of other host materials when it is desired to monitor thicknesses and/or regression rate for purposes of safety, operational control, or research. For example, the sensor could be used to monitor the thicknesses of brake linings or racecar tires and indicate when they need to be replaced. At the time of this reporting, over 200 of these sensors have been installed into a variety of host materials. An MSRS can be made in either of two configurations, denoted ladder and continuous (see Figure 1). A ladder MSRS includes two highly electrically conductive legs, across which narrow strips of electrically resistive material are placed at small increments of length. These strips resemble the rungs of a ladder and are electrically equivalent to many tiny resistors connected in parallel. A substrate material provides structural support for the legs and rungs. The instantaneous sensor resistance is read by an external signal conditioner via wires attached to the conductive legs on the non-eroding end of the sensor. The sensor signal can be transmitted from inside a high-pressure chamber to the ambient environment, using commercially available feedthrough connectors. Miniaturized internal recorders or wireless data transmission could also potentially be employed to eliminate the need for producing penetrations in the chamber case. The rungs are designed so that as each successive rung is eroded away, the resistance changes by an amount that yields a readily measurable signal larger than the background noise. (In addition, signal-conditioning techniques are used in processing the resistance readings to mitigate the effect of noise.) Hence, each discrete change of resistance serves to indicate the arrival of the regressing host material front at the known depth of the affected resistor rung. The average rate of regression between two adjacent resistors can be calculated simply as the distance between the resistors divided by the time interval between their resistance jumps. Advanced data reduction techniques have also been developed to establish the instantaneous surface position and regression rate when the regressing front is between rungs.

Gramer, Daniel J.↗

NASA NTRS2016

Portable Life Support System 2.5 Fan Design and Development

NASA is building a high-fidelity prototype of an advanced Portable Life Support System (PLSS) as part of the Advanced Exploration Systems Program. This new PLSS, designated as PLSS 2.5, will advance component technologies and systems knowledge to inform a future flight program. The oxygen ventilation loop of its predecessor, PLSS 2.0, was driven by a centrifugal fan developed using specifications from the Constellation Program. PLSS technology and system parameters have matured to the point where the existing fan will not perform adequately for the new prototype. In addition, areas of potential improvement were identified with the PLSS 2.0 fan that could be addressed in a new design. As a result, a new fan was designed and tested for the PLSS 2.5. The PLSS 2.5 fan is a derivative of the one used in PLSS 2.0, and it uses the same nonmetallic, canned motor, with a larger volute and impeller to meet the higher pressure drop requirements of the PLSS 2.5 ventilation loop. The larger impeller allows it to operate at rotational speeds that are matched to rolling element bearings, and which create reasonably low impeller tip speeds consistent with prior, oxygen-rated fans. Development of the fan also considered a shrouded impeller design that could allow larger clearances for greater oxygen safety, assembly tolerances and particle ingestion. This paper discusses the design, manufacturing and performance testing of the new fans.

Quinn, Gregory↗

NASA NTRS2011

Use of Smoothed Measured Winds to Predict and Assess Launch Environments

Since many of the larger launch vehicles are operated near their design limits during the ascent phase of flight to optimize payload to orbit, it often becomes necessary to verify that the vehicle will remain within certification limits during the ascent phase as part of the go/no-go review made prior to launch. This paper describes the approach used to predict Ares I-X launch vehicle structural air loads and controllability prior to launch which represents a distinct departure from the methodology of the Space Shuttle and Evolved Expendable Launch Vehicle (EELV) programs. Protection for uncertainty of key environment and trajectory parameters is added to the nominal assessment of launch capability to ensure that critical launch trajectory variables would be within the integrated vehicle certification envelopes. This process was applied by the launch team as a key element of the launch day go/no-go recommendation. Pre-launch assessments of vehicle launch capability for NASA's Space Shuttle and the EELV heavy lift versions require the use of a high-resolution wind profile measurements, which have relatively small sample size compared with low-resolution profile databases (which include low-resolution balloons and radar wind profilers). The approach described in this paper has the potential to allow the pre-launch assessment team to use larger samples of wind measurements from low-resolution wind profile databases that will improve the accuracy of pre-launch assessments of launch availability with no degradation of mission assurance or launch safety.

Cordova, Henry S.↗

NASA NTRSDate not supplied

Probabalistic Risk Analysis and Thermal Margin Process for an Inflatable Aeroshell

Uncertainties always exist in atmospheric entry aeroheating environments and the thermal response of thermal protection system (TPS) material. These uncertainties are mitigated in the design by ap-plying margin and factors of safety to the TPS. Entry vehicle TPS is often conservatively over-sized for the heat loads that are experienced along the entry trajectory by stacking worst-case scenarios together. Additionally, the current TPS design and margin process used by NASA offers very little insight into the risk of over-temperature during flight and the reliability of the heat shield performance [1,3]. A probabilistic margin process can be used to calculate the amount of TPS margin necessary to survive a given entry heat load at a specified level of risk [2,3,4]. The vehicle’s initial entry state (entry velocity, flight path angle, and entry mass) determines the expected atmospheric entry environmental conditions and resulting heat load that the entry vehicle will experience. If there is flexibility in the entry state, then this process can be used to select an appropriate combination of entry state parameters and TPS size to target a desired reentry reliability. This probabilistic margin process allows engineers to make informed aeroshell design, entry-trajectory design, and TPS performance risk trades while preventing excessive TPS margin from being applied. The probabilistic TPS margin process has been performed to determine TPS thickness and entry heating constraints given an acceptable risk level for the Low Earth Orbit Flight Experiment of an Inflatable Decelerator (LOFTID) flight project. The process is used in a manner to size the entry heat load for a given flexible TPS (FTPS) thickness so that it meets project reliability standards while allowing the FTPS and the underlying inflatable structure (IS) to be pushed to adequately high temperatures. Since the LOFTID project is an experimental flight demonstration, it is de-sired to drive the FTPS and IS to temperatures that cover a large range of their thermal response models’ applicability. This will allow the thermal response models to be better improved and validated post-flight using LOFTID’s extensive instrumentation embedded within the aeroshell. The presentation demonstrates how uncertainty analysis is carried out using an end-to-end Monte Carlo process where three separate Monte Carlo simulations are run in sequence. The first Monte Carlo simulation operates on the entry trajectory model to generate trajectory parameter dispersions that are fed into the second Monte Carlo simulation. The second Monte Carlo simulation operates on the aerothermodynamics model to generate aeroheating parameter dispersions that are fed into the third Monte Carlo simulation. The third Monte Carlo simulation operates on the FTPS material thermal response model to generate the final FTPS/IS thermal response dispersions. The end-to-end Monte Carlo simulation propagates the uncertainties of each model into the next to quantify the resulting uncertainty of the FTPS/IS thermal response. The fractional contributions of the uncertain parameters in the trajectory, aerothermal, and thermal response models to the variance in the FTPS/IS thermal response is determined as a byproduct of the Monte Carlo analysis. The structural uncertainty of the FTPS thermal response model is evaluated by flight relevant ground testing and model error analysis using test measurements. This probabilistic TPS margin process had never been applied to an entry vehicle and it is one of the LOFTID project’s goals to demonstrate its merits.

Steven A. Tobin↗