Development of the risk oriented program optimization tool (ROPOT) for orbital space plane design concepts

Indian Space Program (ISP) is one of the largest civilian space programs in the world. ISP is rapidly evolving from a primarily Earth Observation (EO) orientation to a full-fledged space exploitation that includes deep space exploration missions and observatories, satellite navigational (PNT) system, oceanic and climate studies mission and a human spaceflight program. From the beginning, ISP has been envisaged as a self-reliant and largely indigenous space program catering to the socio-economic benefits of the citizens. The space transportation systems (STS) development, starting with the first successful launcher, SLV-3 that placed a 35 kg Rohini satellite on July 18, 1980, is a critical component of the ISP accounting for about 40% of the entire space budget. At present, India has two operational small/ medium satellite launch systems, PSLV and GSLV for polar and geostationary orbits respectively. An up-rated PSLV-XL successfully launched India‟s first deep space mission, Chandrayaan-1 Orbiter, to the lunar polar orbit on October 22, 2008. Two heavy lift launchers, GSLV MK II and GSLV MK III are in development for achieving complete self-reliance in launch vehicles with indigenous cryogenic, liquid and solid motor engines. An ambitious Reusable Launch Vehicle (RLV) program utilizing scramjet and advanced materials is also under development. The recently initiated Indian Human Spaceflight (HSF) program will rely on indigenous STS with a first HSF mission to low Earth orbit planned for 2016.

Our aerospace customers are demanding that we drastically reduce the cost of operating and supporting our products. Our space customer in particular is looking for the next generation of reusable launch vehicle systems to support more aircraft like operation. To achieve this goal requires more than an evolution in materials, processes and systems, what is required is a paradigm shift in the design of the launch vehicles and the processing systems that support the launch vehicles. This paper describes the Automated Informed Maintenance System (AIM) we are developing for NASA's Space Launch Initiative (SLI) Second Generation Reusable Launch Vehicle (RLV). Our system includes an Integrated Health Management (IHM) system for the launch vehicles and ground support systems, which features model based diagnostics and prognostics. Health Management data is used by our AIM decision support and process aids to automatically plan maintenance, generate work orders and schedule maintenance activities along with the resources required to execute these processes. Our system will automate the ground processing for a spaceport handling multiple RLVs executing multiple missions. To accomplish this task we are applying the latest web based distributed computing technologies and application development techniques.

The objective of the study described in this paper was to develop an efficient uncertainty quantification framework capable of analyzing uncertainty in integrated spacecraft system models. Specifically, this paper discusses the capabilities of the developed framework and the res ults when applied to the multidisciplinary analysis of a reusable launch veh icle (RLV). This particular framework is capable of efficiently propagating mixed (inherent and epistemic) uncertainties through complex simulation codes. The Second-Order Probability Theory utilizing a stochastic response surface obta ined with Point-Collocation NonIntrusive Polynomial Chaos was used for the propagation of the mixed uncertainties. This particular methodology was applied to the RLV analysis, and the uncertainty in the output parameters of interested was obtained in terms of intervals at various probability levels. This study has also demonstrated the feasibility of the developed uncertainty quantifica tion framework for efficient propagation of mixed uncertainties in the analysis of complex aerospace systems. I. Introduction ncertainties are generally ubiquitous in the analys is and design of highly complex engineering systems. Uncertainties can arise from the lack of k nowledge in physical modeling (epistemic uncertainty), inherent variations in the systems (a leatory uncertainty), and numerical errors in the computational procedures used for analysis. It is i mportant to account for all of these uncertainties in applications such as robust and reliable design of multi-disciplinary aerospace systems. A reusable la unch vehicle (RLV) is a highly complex aerospace system, which represents a cost viable option for access t o space missions due to its reusability aspect. Since an RLV system is composed of various subsystems which must work together to serve an over-arching p urpose or mission objective/constraint, a multidisciplinary approach for the analysis and des ign of RLV systems is required. Uncertainties are generally present in the models used in each discipline of an RLV multidisciplinary analysis framework. It is important to account for all of these uncertainties for accurate and reliabl e estimates of the RLV system performance. The primary purpose of this paper will be to demonstrate an efficient approach for uncertainty quantification in the multidiscipl inary analysis of a RLV, which has a mixture of aleatory (inherent) and epistemic input uncertainties. Uncertainty quantification will be performed on various output variables of interest f or an RLV system such as the weight and vehicle crossrange. For the propagation of mixed (aleatory-epist emic) uncertainty, Second-Order Probability Theory utilizing Point-Collocation Non-Intrusive Polynomia l Chaos (NIPC) 1 will be used. 2-3 In general, the NIPC methods, which are based on the spectral representa tion of uncertainty, are computationally more effic ient than traditional Monte Carlo methods for moderate number of uncertain variables and can give highly 1 Aerospace Engineer, 4020 Long Beach Blvd, Long Beach, CA, 90807, Member AIAA

It has long been proposed that a reliable space transportation system (STS) should be developed in Japan as in many other countries. Recently this becomes inevitable, due to successive launch failures of rockets which occurred in the last three years in Japan. Consequently an ad hoc Committee on Future Space Transportation System has oeen established by the Science and Technology Agency of Japan. Extensive investigations were conducted about the essential features of a future reusable launch vehicle (RLV) and the technological maturity for vehicle core systems. In conclusion, a fully reusable two stage to orbit (TSTO) spaceplane system equipped with air-breathing engines for its first stage was selected as a target RLV system to be developed in Japan around the mid 2010’s. A road map for research and development towards a future RLV was shown where a well-mixed space infrastructure of RLV and expendable launch vehicle (ELV) was proposed. This paper overviews the technological background to develop a future RLV in Japan and outlines the Committee’s final report issued in June 2000.

NASA's Integrated Space Transportation Plan — 3 rd generation reusable launch vehicle technology update

A major objective of next generation reusable launch vehicle (RLV) programs includes significant improvements in vehicle safety, reliability, and operational costs. In this paper, trajectory generation on approach/ landing (A&L) for RLVs using motion primitives (MPs) and neighboring optimal control (NOC) is discussed. The proposed trajectory generation approach at A&L phase is based on an MP scheme which consists of trims and maneuvers. From an initial point to a given touchdown point, all feasible trajectories that satisfy certain constraints are generated and saved into a trajectory database. An optimal trajectory can then be found off-line by using Dijkstra's algorithm. After perturbations are imposed on the initial states of the off-line optimal trajectory, it is reshaped into a neighboring feasible trajectory on-line by using NOC approach. At this point, a neighboring feasible trajectory existence theorem (NFTET) is investigated and its proof is provided as well. The results show that the vehicle with stuck effectors can be recovered from failures in real time. Finally, robustness issues on NOC approach are briefly discussed.

<div class="htmlview paragraph">Current space transportation systems such as expendable launch vehicles (ELVs) and the reusable Space Transportation System (space shuttle) are very expensive. In some instances they are based on 10 to 20 year old technologies. Newer, lower cost technologies must be applied to designing and manufacturing the next generation of vehicles. Competing effectively in the launch vehicle industry requires an order-of-magnitude reduction in the cost per pound of payloads delivered to orbit.</div> <div class="htmlview paragraph">Payload fairings are critical structural elements in all ELV systems. In addition to carrying and transmitting vehicle loads, they also protect payloads from severe launch and in-flight environments. Developing improved fairing hardware and the associated structural, thermal, acoustic, and separation subsystems is a major undertaking. It requires extensive design, analysis, testing, and systems integration activity.</div> <div class="htmlview paragraph">To maintain a competitive posture in the launch service industry, three years ago McDonnell Douglas Space System Company (MDSSC) initiated a research and development program on composite fairings. The program investigated and refined design and analysis, process techniques, and fabrication methods for large diameter composite fairings (<span class="xref">1</span>,<span class="xref">2</span>).<span class="xref">*</span></div> <div class="htmlview paragraph">Substantial progress was achieved in composite fairing material selection, design and analysis, and manufacturing methods characterization. Eight composite fairing configurations with different constructions, and five aerodynamic forebody shapes with improved performance characteristics, were evaluated and studied. A final fairing configuration was selected for further study and the fabrication of a demonstration/test article is underway. This paper presents the composite fairing structural alternatives investigated and summarizes the results of the major trade studies undertaken.</div>

The development and application of the differential evolution (DE) optimisation algorithm to the problem of worst-case analysis of nonlinear control laws for hypersonic re-entry vehicles is described. The algorithm is applied to the problem of evaluating a proposed nonlinear handling qualities clearance criterion for a detailed simulation model of a hypersonic re-entry vehicle (also known as a reusable launch vehicle (RLV)) having a full-authority nonlinear dynamic inversion (NDI) flight control law. A hybrid version of the differential evolution algorithm, incorporating local gradient-based optimisation, is also developed and evaluated. Comparisons of computational complexity and global convergence properties reveal the significant benefits which may be obtained through hybridisation of the standard differential evolution algorithm. The proposed optimisation-based approach to worst-case analysis is shown to have significant potential for improving both the reliability and efficiency of the flight clearance process for next generation RLV’s.

NASA has a technology program in place to enable the development of a next generation Reusable Launch Vehicle that will carry our future payloads into orbit at a much-reduced cost. The VentureStar, Lifting Body (LB) flight vehicle, is one of the potential reusable launch vehicle configurations being studied. A LB vehicle has no wings and derives its lift solely from the shape of its body, and has the unique advantages of superior volumetric efficiency, better aerodynamic efficiency at high angles-of-attack and hypersonic speeds, and reduced thermal protection system weight. Classically, in a ballistic vehicle, drag has been employed to control the level of deceleration in reentry. In the LB, lift enables the vehicle to decelerate at higher altitudes for the same velocity and defines the reentry corridor which includes a greater cross range. This paper outlines the flight stability and control aspects of our LB heritage which was utilized in the design of the VentureStar LB and its test version, the X-33. NASA and the U.S. Air Force have a rich heritage of LB vehicle design and flight experience. In the initial LB Program, eight LB's were built and over 225 LB test flights were conducted through 1975. Three LB series were most significant in the advancement of today's LB technolocy: the M2-F; the HL-10; and the X-24 series. The M2-F series was designed by NASA Ames Research Center, the HL-10 series by NASA Langley Research Center, and the X-24 series by the U. S. Air Force. LB vehicles are alive again today with the X- 33, X-38, and VentureStar.

Federal tax incentives and financing for the reusable launch vehicle

An analysis of cost savings that could be realized on active debris removal (ADR) missions through the use of reusable launch vehicles (RLVs) has been performed. Launch vehicle price estimates were established for three levels of RLV development, based on varying levels of technological development and market competition. An expendable launch vehicle (ELV) price estimate was also established as a point of comparison. These price estimates were used to form two separate debris removal mission cost estimates, based on previously proposed debris removal mission concepts. The results of this analysis indicate that RLVs could reduce launch prices to levels between 19.6% and 92.8% cheaper than ELVs, depending on the level of RLV maturity. It was also determined that a RLV could be used to realize total ADR mission cost savings of between 2.8% (for a partially RLV in an uncompetitive market) and 21.7% (for a fully RLV in a competitive market).

In the preliminary design phase of aircraft design, estimating the production cost accurately is a challenging task. At this stage, many design parameters that affect the overall cost are still undefined. This paper establishes cost-estimation models for civil, commercial aircraft using a parametric cost analysis (PCA) approach. Aircraft are characterised based on their size, ranging from a wide body to executive jets, into four categories. Key design parameters, such as maximum take-off weight, number of passengers, range, wing area, span, fuselage length, to name a few, are likely to be available in the preliminary design stage and significantly impact the aircraft design. These variables either directly or indirectly affect the overall production cost or performance. The PCA approach includes both correlation and multiple linear regression techniques. The empirical models thus developed were able to predict the aircraft cost with an error of less than ±4% for all aircraft categories considered. Two aircraft in each defined category were not part of the PCA models and were used to verify the models. The proposed models provide the ability to estimate the aircraft cost quickly in the early stages of the preliminary design phase and provide the possibility of performing parametric studies involving the key variables to determine the cost sensitivity to the main design parameters.

Main Focus is the design of an air breathing TSTO reusable launch vehicle using in -flight LO X-collection. Integration of such a device in a winged space transport system can have potentials to reduce weight and hence costs for access to space. Plant architecture, influencing parameters, mode of operation and optimum working conditions of the coll ection device will be discussed. This will lead to a suitable design point, field of operation and requirements for application in the first stage of a reusable space transport. Preliminary investigations have shown that a turbojet propelled Mach -4 vehicle is most suitable for application of the extraction system. Based on this a complete design of the TSTO for ACES application is carried out in terms of performance, feasibility and safety reasons as well as reductions of complexity and development effort. The evaluation of potentials and risks of a technology is only possible by its examination as integral part of the entire system.

To enhance the flexibility and responsiveness of commercial space operations, improved surveillance techniques are required to track the ascent and descent of a space vehicle through the national airspace system. In this paper, a prototype UAT Automatic Dependent Surveillance - Broadcast (or ADS-B) transmitter developed by Embry-Riddle Aeronautical University for research purposes is presented. The past flight history and design of this prototype is briefly discussed. Next, the experimental results of a flight onboard a commercial reusable launch vehicle (RLV) that functions similar to a sounding rocket is presented and analyzed. The paper concludes with a discussion of possible future work in adapting ADS-B to support commercial space operations for not just reusable launch vehicles, but other commercial space transportation vehicles as well.

A High-Speed Flight Demonstration (HSFD) project is the latest in a series of flight experiments in a joint research program into reusable space transportation systems being conducted by the National Aerospace Laboratory of Japan (NAL) and the National Space Development Agency of Japan (NASDA) (NAL and NASDA were merged to form Japan Aerospace Exploration Agency-JAXA on October 1, 2003). The project consists of two phases: Phases I and II. The objective of Phase I is to verify approach and landing systems for the final part of the return phase of a winged re-entry vehicle. The Phase I flight experiment was conducted in Christmas Island in Republic of Kiribati, from October to December 2002. Through three flights, the approach and landing system was evaluated during steep gliding, which is characteristic of return flights of winged reentry vehicles, and autonomous flight technologies essential for future space transportation systems were accumulated. The purpose of the Phase II is to clarify the transonic aerodynamic characteristics of the demonstrator, and data obtained from wind tunnel tests and Computational Fluid Dynamics (CFD) are evaluated against the flight test data. A Flight experiment of Phase II was conducted at Esrange test site in Sweden from May to July 2003 in collaboration with Centre National d'Etudes Spatiales of France (CNES) which developed and operated the balloon system. Although the flight experiment was interrupted due to the anomaly of the recovery system, various data at Mach 0.8 were obtained through the flight, and fully autonomous flight control technologies for transonic speed region were demonstrated. The data obtained from the HSFD project are expected to provide essential information for development of future reusable space transportation systems.