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A methodology for specifying the radiometric requirements for ocean color remote sensing systems is described. Consideration is given to the noise equivalent radiance, the saturation radiance, the polarization sensitivity, and the calibration and stability. The degree of polarization across a scan line for the CZCS and MODIS orbits is presented.
This slide presentation presents an overview of work system requirements, extravehicular activity system evolution, key issues, future needs, and a summary. Key issues include pressure suits, life support systems, system integration, biomedical requirements, and work and mobility aids.
Recent advanced in model-checking have made it practical to formally verify the correctness of many complex synchronous systems (i.e., systems driven by a single clock). However, many computer systems are implemented by asynchronously composing several synchronous components, where each component has its own clock and these clocks are not synchronized. Formal verification of such Globally Asynchronous/Locally Synchronous (GA/LS) architectures is a much more difficult task. In this report, we describe a methodology for developing and reasoning about such systems. This approach allows a developer to start from an ideal system specification and refine it along two axes. Along one axis, the system can be refined one component at a time towards an implementation. Along the other axis, the behavior of the system can be relaxed to produce a more cost effective but still acceptable solution. We illustrate this process by applying it to the synchronization logic of a Dual Fight Guidance System, evolving the system from an ideal case in which the components do not fail and communicate synchronously to one in which the components can fail and communicate asynchronously. For each step, we show how the system requirements have to change if the system is to be implemented and prove that each implementation meets the revised system requirements through modelchecking.
Spacecraft and launch vehicle ground processing activities require a variety of unique human activities. These activities are being documented in a Primitive motion capture library. The Library will be used by the human factors engineering in the future to infuse real to life human activities into the CAD models to verify ground systems human factors requirements. As the Primitive models are being developed for the library the project has selected several current human factors issues to be addressed for the SLS and Orion launch systems. This paper explains how the Motion Capture of unique ground systems activities are being used to verify the human factors analysis requirements for ground system used to process the STS and Orion vehicles, and how the primitive models will be applied to future spacecraft and launch vehicle processing.
During Phase II, the Engineering Development Phase, the SP-100 Project has defined and is pursuing a new approach to developing reliable power systems. The approach to developing such a system during the early technology phase is described along with some preliminary examples to help explain the approach. Developing reliable components to meet space reactor power system requirements is based on a top-down systems approach which includes a point design based on a detailed technical specification of a 100-kW power system. The SP-100 system requirements implicitly recognize the challenge of achieving a high system reliability for a ten-year lifetime, while at the same time using technologies that require very significant development efforts. A low-cost method for assessing reliability, based on an understanding of fundamental failure mechanisms and design margins for specific failure mechanisms, is being developed as part of the SP-100 Program.
The success of any effort depends upon the effective initial definition of its purpose, in terms of the needs to be satisfied and the goals to be fulfilled. If the desired product is "A System" that is well-characterized, these high-level need and goal statements can be transformed into system requirements by traditional systems engineering techniques. The satisfaction of well-designed requirements can be tracked by fairly straightforward cost, schedule, and technical performance metrics. Unfortunately, some types of efforts, including those that NASA terms "Programs," tend to resist application of traditional systems engineering practices. In the NASA hierarchy of efforts, a "Program" is often an ongoing effort with broad, high-level goals and objectives. A NASA "project" is a finite effort, in terms of budget and schedule, that usually produces or involves one System. Programs usually contain more than one project and thus more than one System. Special care must be taken in the formulation of NASA Programs and their projects, to ensure that lower-level project requirements are traceable to top-level Program goals, feasible with the given cost and schedule constraints, and measurable against top-level goals. NASA Programs and projects are tasked to identify the advancement of technology as an explicit goal, which introduces more complicating factors. The justification for funding of technology development may be based on the technology's applicability to more than one System, Systems outside that Program or even external to NASA. Application of systems engineering to broad-based technology development, leading to effective measurement of the benefits, can be valid, but it requires that potential beneficiary Systems be organized into a hierarchical structure, creating a "system of Systems." In addition, these Systems evolve with the successful application of the technology, which creates the necessity for evolution of the benefit metrics to reflect the changing baseline. Still, economic metrics for technology development in these Programs and projects remain fairly straightforward, being based on reductions in acquisition and operating costs of the Systems. One of the most challenging requirements that NASA levies on its Programs is to plan for the commercialization of the developed technology. Some NASA Programs are created for the express purpose of developing technology for a particular industrial sector, such as aviation or space transportation, in financial partnership with that sector. With industrial investment, another set of goals, constraints and expectations are levied on the technology program. Economic benefit metrics then expand beyond cost and cost savings to include the marketability, profit, and investment return requirements of the private sector. Commercial investment criteria include low risk, potential for high return, and strategic alignment with existing product lines. These corporate criteria derive from top-level strategic plans and investment goals, which rank high among the most proprietary types of information in any business. As a result, top-level economic goals and objectives that industry partners bring to cooperative programs cannot usually be brought into technical processes, such as systems engineering, that are worked collaboratively between Industry and Government. In spite of these handicaps, the top-level economic goals and objectives of a joint technology program can be crafted in such a way that they accurately reflect the fiscal benefits from both Industry and Government perspectives. Valid economic metrics can then be designed that can track progress toward these goals and objectives, while maintaining the confidentiality necessary for the competitive process.
The key systems requirements for ceilometer systems are described. The following items are included: range must be 10,000 ft.; laser emission must conform to the bureau of radiological health class I performance; system must detect two lowest cloud layers; display must be in either English or metric units; and system must be capable of self monitoring and testing performance. Based upon the requirements competitive prototype cloud height indicator systems are to be built.
Performance deficiencies in aerial liquid and dry dispersal systems are identified. Five control system concepts are explored: (1) end of field on/off control; (2) manual control of particle size and application rate from the aircraft; (3) manual control of deposit rate on the field; (4) automatic alarm and shut-off control; and (5) fully automatic control. Operational aspects of the concepts and specifications for improved control configurations are discussed in detail. A research plan to provide the technology needed to develop the proposed improvements is presented along with a flight program to verify the benefits achieved.
Today's complex systems require understanding beyond one person s capability to comprehend. Each system requires a team to divide the system into understandable subsystems which can then be analyzed with an Integrated Hazard Analysis. The team must have both specific experiences and diversity of experience. Safety experience and system understanding are not always manifested in one individual. Group dynamics make the difference between success and failure as well as the difference between a difficult task and a rewarding experience. There are examples in the news which demonstrate the need to connect the pieces of a system into a complete picture. The Columbia disaster is now a standard example of a low consequence hazard in one part of the system; the External Tank is a catastrophic hazard cause for a companion subsystem, the Space Shuttle Orbiter. The interaction between the hardware, the manufacturing process, the handling, and the operations contributed to the problem. Each of these had analysis performed, but who constituted the team which integrated this analysis together? This paper will explore some of the methods used for dividing up a complex system; and how one integration team has analyzed the parts. How this analysis has been documented in one particular launch space vehicle case will also be discussed.
The Advanced Electric Propulsion System (AEPS) program will develop a flight 13kW Hall thruster propulsion system based on NASA's HERMeS thruster. The AEPS system includes the Hall Thruster, the Power Processing Unit (PPU) and the Xenon Flow Controller (XFC). These three primary components must operate together to ensure that the system generates the required combinations of thrust and specific impulse at the required system efficiencies for the desired system lifetime. At the highest level, the AEPS system will be integrated into the spacecraft and will receive power, propellant, and commands from the spacecraft. Power and propellant flow rates will be determined by the throttle set points commanded by the spacecraft. Within the system, the major control loop is between the mass flow rate and thruster current, with time-dependencies required to handle all expected transients, and additional, much slower interactions between the thruster and cathode temperatures, flow controller and PPU. The internal system interactions generally occur on shorter timescales than the spacecraft interactions, though certain failure modes may require rapid responses from the spacecraft. The AEPS system performance model is designed to account for all these interactions in a way that allows evaluation of the sensitivity of the system to expected changes over the planned mission as well as to assess the impacts of normal component and assembly variability during the production phase of the program. This effort describes the plan for the system performance model development, correlation to NASA test data, and how the model will be used to evaluate the critical internal and external interactions. The results will ensure the component requirements do not unnecessarily drive the system cost or overly constrain the development program. Finally, the model will be available to quickly troubleshoot any future unforeseen development challenges.
The engineering process of Design for Reliability (DfR) is well established in the automotive and aerospace industries. DfR should be useful in the future development of space life support systems. DfR is a sequence of tasks that develop system requirements and plan reliability analysis and testing. First and fundamentally, the reliability requirement is defined. Next the system reliability model is developed, often using a reliability block diagram. The overall system reliability requirement is allocated to the subsystems and an estimate of the attainable reliability is made. This expected reliability can be improved by simplifying the design by removing components or by replacing less reliable components. Improving reliability can require difficult compromises, such as reducing performance requirements, increasing budget, or extending testing. The actual system reliability can be determined only by testing, which should continue long enough to provide the required confidence in the measured value. New systems often have unexpected design errors that cause failures in early testing. The usual reliability improvement process of testing, finding the failure modes, and redesigning to remove them reduces the failure rate and is referred to as “reliability growth.” After redesign has been completed, the system should be further tested to determine the actual achieved reliability more accurately. If the final system failure rate is too high, redundant systems can be used to improve overall operational reliability. Adding redundancy simply to increase the one- or two-fault tolerance metric may sometimes reduce reliability. Reliability can be improved in three ways: redesigning the system to include more reliable subsystems and components, reliability growth testing and failure mode removal, and by using parallel redundant systems. DfR should combine these approaches to achieve the required reliability while managing performance, cost, and schedule.
The engineering process of Design for Reliability (DfR) is well established in the automotive and aerospace industries. DfR should be useful in the future development of space life support systems. DfR is a sequence of tasks that develop system requirements and plan reliability analysis and testing. First and fundamentally, the reliability requirement is defined. Next the system reliability model is developed, often using a reliability block diagram. The overall system reliability requirement is allocated to the subsystems and an estimate of the attainable reliability is made. This expected reliability can be improved by simplifying the design by removing components or by replacing less reliable components. Improving reliability can require difficult compromises, such as reducing performance requirements, increasing budget, or extending testing. The actual system reliability can be determined only by testing, which should continue long enough to provide the required confidence in the measured value. New systems often have unexpected design errors that cause failures in early testing. The usual reliability improvement process of testing, finding the failure modes, and redesigning to remove them reduces the failure rate and is referred to as “reliability growth.” After redesign has been completed, the system should be further tested to determine the actual achieved reliability more accurately. If the final system failure rate is too high, redundant systems can be used to improve overall operational reliability. Adding redundancy simply to increase the one- or two-fault tolerance metric may sometimes reduce reliability. Reliability can be improved in three ways: redesigning the system to include more reliable subsystems and components, reliability growth testing and failure mode removal, and by using parallel redundant systems. DfR should combine these approaches to achieve the required reliability while managing performance, cost, and schedule.
Space Station FDMA communications system requirements, design, and analysis are addressed. The analysis is primarily based on numerical results generated by a computer simulation system called SCSS. The time-line communications performance during real time mission operation is also discussed. The purpose of this paper is three-fold: introduction to Space Station multiple access communications system requirements, demonstration of system analysis by a computer tool, and design of an FDMA communications system for the Space Station.
The SLS Program chose to implement a Model-based Design and Model-based Requirements approach for managing component design information and system requirements. This approach differs from previous large-scale design efforts at Marshall Space Flight Center where design documentation alone conveyed information required for vehicle design and analysis and where extensive requirements sets were used to scope and constrain the design. The SLS Navigation Team has been responsible for the Program-controlled Design Math Models (DMMs) which describe and represent the performance of the Inertial Navigation System (INS) and the Rate Gyro Assemblies (RGAs) used by Guidance, Navigation, and Controls (GN&C). The SLS Navigation Team is also responsible for the navigation algorithms. The navigation algorithms are delivered for implementation on the flight hardware as a DMM. For the SLS Block 1-B design, the additional GPS Receiver hardware is managed as a DMM at the vehicle design level. This paper provides a discussion of the processes and methods used to engineer, design, and coordinate engineering trades and performance assessments using SLS practices as applied to the GN&C system, with a particular focus on the Navigation components. These include composing system requirements, requirements verification, model development, model verification and validation, and modeling and analysis approaches. The Model-based Design and Requirements approach does not reduce the effort associated with the design process versus previous processes used at Marshall Space Flight Center. Instead, the approach takes advantage of overlap between the requirements development and management process, and the design and analysis process by efficiently combining the control (i.e. the requirement) and the design mechanisms. The design mechanism is the representation of the component behavior and performance in design and analysis tools. The focus in the early design process shifts from the development and management of design requirements to the development of usable models, model requirements, and model verification and validation efforts. The models themselves are represented in C/C++ code and accompanying data files. Under the idealized process, potential ambiguity in specification is reduced because the model must be implementable versus a requirement which is not necessarily subject to this constraint. Further, the models are shown to emulate the hardware during validation. For models developed by the Navigation Team, a common interface/standalone environment was developed. The common environment allows for easy implementation in design and analysis tools. Mechanisms such as unit test cases ensure implementation as the developer intended. The model verification and validation process provides a very high level of component design insight. The origin and implementation of the SLS variant of Model-based Design is described from the perspective of the SLS Navigation Team. The format of the models and the requirements are described. The Model-based Design approach has many benefits but is not without potential complications. Key lessons learned associated with the implementation of the Model Based Design approach and process from infancy to verification and certification are discussed
Helicopter aerial dispersal systems were studied to ascertain constraints to the system, the effects of removal of limitations (technical and FAA regulations), and subsystem improvements. Productivity indices for the aircraft and swath effects were examined. Typical missions were formulated through conversations with operators, and differing gross weight aircraft were synthesized to perform these missions. Economic analysis of missions and aircraft indicated a general correlation of small aircraft (3000 lb gross weight) suitability for small fields (25 acres), and low dispersion rates (less than 32 lb/acre), with larger aircraft (12,000 lb gross weight) being more favorable for bigger fields (200 acres) and heavier dispersal rates (100 lb/acre). Operator problems, possible aircraft and system improvements, and selected removal of operating limitations were reviewed into recommendations for future NASA research items.
EVA missions, beginning with the Space Shuttle and continuing through the 80's and 90's are evaluated in an evolutionary perspective. The projected large space structures, to be assembled in orbit, will require EVA for their construction. Various parameters of EVA are analyzed, including: EVA tasks and system requirements, system pressurization level, and radiation protection. Existing technology and already proven EVA techniques are considered adequate for future EVA missions, with some modifications. Space suit pressure, for example, can be increased from 4 to 8 psi, depending on ambient pressure in the spacecraft. Radiation shielding might have to be improved for higher altitude LEO's. Also, new restraint systems will probably be required for the more complicated EVA tasks.
This report's objectives were to review regulatory and guidance documents to identify requirements and considerations for the design and operation of automated systems that perform the functions of automated collision avoidance (ACA) and/or automated return to course (ARTC). The importance of this work is twofold: 1) to help focus efforts addressing new automated collision avoidance and return to course systems requirements and considerations, and 2) in pulling together available requirements from multiple sources, generate a master resource for these automated system requirements and considerations. Reviewed documents include regulations, industry standards and research papers that specifically address ACA and ARTC systems, other automatedl control and guidance systems, ICAO (International Civil Aviation Organization) and Eurocontrol documents.
Civil and military applications of space transportation have been pursued for just over 50 years and there has been, and still is, a need for safe, dependable, affordable, and sustainable space transportation systems. Fully expendable and partially reusable space transportation systems have been developed and put in operation that have not adequately achieved this need. Access to space is technically achievable, but presently very expensive and will remain so until there is a breakthrough in the way we do business. Since 1991 the national Space Propulsion Synergy Team (SPST) has reviewed and assessed the lessons learned from the major U.S. space programs of the past decades focusing on what has been learned from the assessment and control of Life Cycle Cost (LCC) from these systems. This paper presents the results of a selected number of studies and analyses that have been conducted by the SPST addressing the need, as well as the solutions, for improvement in LCC. The major emphasis of the SPST processes is on developing the space transportation system requirements first (up front). These requirements must include both the usual system flight performance requirements and also the system functional requirements, including the infrastructure on Earth's surface, in-space and on the Moon and Mars surfaces to determine LCC. This paper describes the development of specific innovative engineering and management approaches and processes. This includes a focus on flight hardware maturity for reliability, ground operations approaches, and business processes between contractor and government organizations. A major change in program/project cost control is being proposed by the SPST to achieve a sustainable space transportation system LCC - controlling cost as a program metric in addition to the existing practice of controlling performance and weight. Without a firm requirement and methodically structured cost control, it is unlikely that an affordable and sustainable space transportation system LCC will ever be achieved. '