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The Photonics group top level processes regarding development and qualification of space flight hardware for optoelectronic components and optical systems along with our contributions and successes with LIDAR systems, space flight reliability heritage, and our testing capabilities are conveyed with this power point presentation.
Dual-function integrated circuits use one shift register array instead of two to feed video data to seven tape channels thereby making system more reliable and less expensive than predecessor. Basic principle of reformatter is also applicable to multiplexing and frequency conversion circuits.
The Xpress Transfer Protocol (XTP) is a reliable, real-time, light weight transfer layer protocol. Current transport layer protocols such as DoD's Transmission Control Protocol (TCP) and ISO's Transport Protocol (TP) were not designed for the next generation of high speed, interconnected reliable networks such as fiber distributed data interface (FDDI) and the gigabit/second wide area networks. Unlike all previous transport layer protocols, XTP is being designed to be implemented in hardware as a VLSI chip set. By streamlining the protocol, combining the transport and network layers and utilizing the increased speed and parallelization possible with a VLSI implementation, XTP will be able to provide the end-to-end data transmission rates demanded in high speed networks without compromising reliability and functionality. This paper describes the operation of the XTP protocol and in particular, its error, flow and rate control; inter-networking addressing mechanisms; and multicast support features, as defined in the XTP Protocol Definition Revision 3.4.
A recent study performed for the Aerocapture Technology Area in the In-Space Propulsion Technology Projects Office at the Marshall Space Flight Center investigated the relative risk of various capture techniques for Mars missions. Aerocapture has been proposed as a possible capture technique for future Mars missions but has been perceived by many in the community as a higher risk option as compared to aerobraking and propulsive capture. By performing a probabilistic risk assessment on aerocapture, aerobraking and propulsive capture, a comparison was made to uncover the projected relative risks of these three maneuvers. For mission planners, this knowledge will allow them to decide if the mass savings provided by aerocapture warrant any incremental risk exposure. The study focuses on a Mars Sample Return mission currently under investigation at the Jet Propulsion Laboratory (JPL). In each case (propulsive, aerobraking and aerocapture), the Earth return vehicle is inserted into Martian orbit by one of the three techniques being investigated. A baseline spacecraft was established through initial sizing exercises performed by JPL's Team X. While Team X design results provided the baseline and common thread between the spacecraft, in each case the Team X results were supplemented by historical data as needed. Propulsion, thermal protection, guidance, navigation and control, software, solar arrays, navigation and targeting and atmospheric prediction were investigated. A qualitative assessment of human reliability was also included. Results show that different risk drivers contribute significantly to each capture technique. For aerocapture, the significant drivers include propulsion system failures and atmospheric prediction errors. Software and guidance hardware contribute the most to aerobraking risk. Propulsive capture risk is mainly driven by anomalous solar array degradation and propulsion system failures. While each subsystem contributes differently to the risk of each technique, results show that there exists little relative difference in the reliability of these capture techniques although uncertainty for the aerocapture estimates remains high given the lack of in-space demonstration.
The Xpress Transfer Protocol (XTP) is a reliable, light weight transfer layer protocol. Current transport layer protocols such as DoD's Transmission Control Protocol (TCP) and ISO's Transport Protocol (TP) were not designed for the next generation of high speed, interconnected reliable networks such as fiber distributed data interface (FDDI) and the gigabit/second wide area networks. Unlike all previous transport layer protocols, XTP is being designed to be implemented in hardware as a VLSI chip set. By streamlining the protocol, combining the transport and network layers, and utilizing the increased speed and parallelization possible with a VLSI implementation, XTP will be able to provide the end-to-end data transmission rates demanded in the high speed networks without compromising reliability and functionality. This tutorial briefly describes the operation of the XTP protocol and in particular, its error, flow and rate control; inter-networking addressing mechanisms; and multicast support features, as defined in the XTP Protocol Definition Revision 3.4.
The requirement for ultradependability of computer systems in future avionics and space applications necessitates a top-down, integrated systems engineering approach for design, implementation, testing, and operation. The functional analyses of hardware and software systems must be combined by models that are flexible enough to represent their interactions and behavior. The information contained in these models must be accessible throughout all phases of the system life cycle in order to maintain consistency and accuracy in design and operational decisions. One approach being taken by researchers at Ames Research Center is the creation of an object-oriented environment that integrates information about system components required in the reliability evaluation with behavioral information useful for diagnostic algorithms. Procedures have been developed at Ames that perform reliability evaluations during design and failure diagnoses during system operation. These procedures utilize information from a central source, structured as object-oriented fault trees. Fault trees were selected because they are a flexible model widely used in aerospace applications and because they give a concise, structured representation of system behavior. The utility of this integrated environment for aerospace applications in light of our experiences during its development and use is described. The techniques for reliability evaluation and failure diagnosis are discussed, and current extensions of the environment and areas requiring further development are summarized.
The requirement for ultradependability of computer systems in future avionics and space applications necessitates a top-down, integrated systems ingeneering approach for design, implementation, testing, and operation. The functional analyses of hardware and software systems must be combined by models that are flexible enough to represent their interactions and behavior. The information contained in these models must be accessible throughout all phases of the system life cycle in order to maintain consistency and accuracy in design and operational decisions. One approach being taken by researchers at Ames Research Center is the creation of an object-oriented environment that integrates information about system components required in the reliability evaluation with behavioral information useful for diagnostic algorithms.
The problem in assessment of electroencephalograms has been the subject element in visual evaluations of such records as well as the need for extensive training (2 or 3 years) before suitable candidate is capable to perform this task. There have been variety of computer type analyses of EEG during the past ten years or more, and we are still searching for a reliable, inexpensive, online data processing procedure for this task. We are aware our Soviet colleagues are routinely using computer assessment of EEGs, using relatively inexpensive hardware. In our report we endeavor to show how four different assessment procedures (visual assessment, zero crossing technique, smoothing and peak counting technique, and Weibull Statistic) may be utilized, and indicate the advantages and disadvantages peculiar to each. We wish to point out that in the Gemini 7 EEG analog records supplied to the Henry Ford Hospital, Channel II (mid-line central to mid-line occipital electrodes) could not be assessed reliably after approximately 26 hours of flight. Channel I (left central to left occipital electrodes) provided the more assessable EEG up to approximately 54 hours of flight. These channels were compared during the first 26 hours of flight, and no significant difference was found by our various assessment procedures. We, therefore, have limited this report to the findings on the analyses of Channel I during its approximate 54 hours of flight.
The memory requirements for ultra-reliable computers are expected to increase due to future increases in mission functionality and operating-system requirements. This increase will have a negative effect on the reliability and cost of the system. Increased memory size will also reduce the ability to reintegrate a channel after a transient fault, since the time required to reintegrate a channel in a conventional fault-tolerant processor is dominated by memory realignment time. A Byzantine Resilient Fault-Tolerant Processor with Fault-Tolerant Shared Memory (FTP/FTSM) is presented as a solution to these problems. The FTSM uses an encoded memory system, which reduces the memory requirement by one-half compared to a conventional quad-FTP design. This increases the reliability and decreases the cost of the system. The realignment problem is also addressed by the FTSM. Because any single error is corrected upon a read from the FTSM, a faulty channel's corrupted memory does not need realignment before reintegration of the faulty channel. A combination of correct-on-access and background scrubbing is proposed to prevent the accumulation of transient errors in the memory. With a hardware-implemented scrubber, the scrubbing cycle time, and therefore the memory fault latency, can be upper-bounded at a small value. This technique increases the reliability of the memory system and facilitates validation of its reliability model.
Hardware and system testing mathematical model assuming stochastic design and manufacturing defects detection and nondetection probability equations analogous to reliability function
The paper outlines the approach used in developing DMS (Data Management Subsystem) alternatives for the LST (Large Space Telescope) and in selecting the concept considered to be the most cost effective means of implementing the LST DMS requirements. Two candidate DMS concepts are discussed: a functionally integrated and a functionally separated one. For the single vehicle LST program, separation of the DMS functions best provides high reliability, operations flexibility, minimal interface complexity, and the least complex software development and verification task. The use of available hardware and NASA standard components is stressed.
Demonstration advanced anionics system (DAAS) function description, hardware description, operational evaluation, and failure mode and effects analysis (FMEA) are provided. Projected advanced avionics system (PAAS) description, reliability analysis, cost analysis, maintainability analysis, and modularity analysis are discussed.
Two programs are described which involve the application of advanced control techniques to the design of engine control algorithms. Multivariable control theory is used in the F100 MVCS (multivariable control synthesis) program to design controls which coordinate the control inputs for improved engine performance. A systematic method for handling a complex control design task is given. Methods of analytical redundancy are aimed at increasing the control system reliability. The F100 DIA (detection, isolation, and accommodation) program, which investigates the uses of software to replace or augment hardware redundancy for certain critical engine sensor, is described.
The Diagnostic Emulation Technique was developed within the System Validation Methods Branch as a part of the development of methods for the analysis of the reliability of highly reliable, fault tolerant digital avionics systems. This is a general technique which allows for the emulation of a digital hardware system. The technique is general in the sense that it is completely independent of the particular target hardware which is being emulated. Parts of the system are described and emulated at the logic or gate level, while other parts of the system are described and emulated at the functional level. This algorithm allows for the insertion of faults into the system, and for the observation of the response of the system to these faults. This allows for controlled and accelerated testing of system reaction to hardware failures in the target machine. This document describes in detail how the algorithm was implemented at NASA Langley Research Center and gives instructions for using the system.
Flight systems must be validated to show that they are consistent with the requirements of their intended applications. While high reliability is difficult to validate, the additional complexity of fault tolerance further compounds the validation problem. The objective of NASA’s research is to develop a methodology for designing validatable fault-tolerant systems. Under the design-for-validation philosophy, emphasis is placed on developing validation methods that can be incorporated into the design process right from the start and design methods and guidance which, while incorporating fault tolerance, can assure validatability. This paper examines the statistical issues of validating highly reliable, fault tolerant system. There are many problems associated with traditional methods of designing and validating these potentially complex hardware and software systems. Useful design-for-validation methods, which include structured specification and design methodologies, mathematical proof techniques, analytical modeling, simulation and emulation, and physical testing, are discussed. Important design issues associated with fault tolerance are presented along with the related validation concerns which must be addressed. Experience has shown that synchronization and Byzantine resilience must accompany fault tolerance. Other design attributes associated with fault tolerance may be used by a designer on the basis of cost, weight, performance, and validation considerations.
This paper describes a design for fiber-optic sensing and control in advanced aircraft Electronic Engine Control (EEC). The recommended architecture is an on-engine EEC which contains electro-optic interface circuits for fiber-optic sensors. Size and weight are reduced by multiplexing arrays of functionally similar sensors on a pairs of optical fibers to common electro-optical interfaces. The architecture contains interfaces to seven sensor groups. Nine distinct fiber-optic sensor types were found to provide the sensing functions. Analysis revealed no strong discriminator (except reliability of laser diodes and remote electronics) on which to base a selection of preferred common interface type. A hardware test program is recommended to assess the relative maturity of the technologies and to determine real performance in the engine environment.
Advantage of non-intrusiveness, capability for field measurement, and increased availability of IR imaging systems have resulted in their wider use for aerodynamic research. However, certain difficulties persist while using currently available systems for such applications. A critical evaluation of the IR imaging systems is presented on the basis of the state-of-the-art of IR imaging technology and experiences in wind tunnel and flight testing at NASA's Langley Research Center. The requirements for using IR thermography as a measurement tool in aerodynamic research are examined in terms of range, sensitivity, and accuracy of temperature measurement, temporal and spatial resolution, and features of target. Deficiencies of present IR imaging systems are identified, and user precautions to avoid such problems by proper selection and operation of these units are suggested. Different aspects of imager performance such as imager optics, video capabilities, and environmental tolerance are discussed. Electronic data recording and image processing hardware and software requirements are evaluated. Slit response tests and spatial resolution are discussed with the objective of obtaining reliable, accurate, and meaningful information from IR thermography measurements for aerodynamic studies.
Doppler and ranging measurements between spacecraft can be obtained only when the ratio of the total received signal power to noise power density (P(sub t/N(sub 0)) at the receiving spacecraft is sufficiently large that reliable signal detection can be achieved within a reasonable time period. In this paper, the requirements on P(sub t)/N(sub 0) for reliable carrier signal detection is calculated as a function of various system parameters, including characteristics of the spacecraft computing hardware and a priori uncertainty in spacecraft-spacecraft relative velocity and acceleration. Also calculated is the P(sub t)/N(sub 0) requirement for relaible detection of a ranging signal, consistting of a carrier with pseudo-noise phase modulation. Once the P(sub t)/N(sub 0) requirement is determined, then for a given set of assumed spacecraft telecommunication characteristics (transmitted signal power, antenna gains, receiver noise temperatures) it is possible to calculate the maximum range at which a carrier signal or ranging signal may be acquired. A brief error covariance analysis has been conducted to illustrate the utility of in situ Doppler and ranging measurements for Mars approach navigation. The results indicate that navigation accuracies of a few kilometers can be achieved with either data type. The analysis also illustrates dependency of the achievable accuracy on the approach trajectory velocity.