This dissertation is focused on the quantitative metrics of performance, cost, and reliability for future launch vehicles. Methods are developed that hold performance constant for a required mission and payload so that cost and reliability can be traded. Reliability strategies such as reducing the number of engines, increasing the thrust-to-weight ratio, and adding redundant subsystems all increase launch vehicle reliability. However, there are few references that illustrate the cost of increasing launch vehicle reliability in a disciplined, integrated approach. For launch vehicle design, integrated performance, cost, and reliability disciplines are required to show the sensitivity of cost to different reliability strategies. A methodology is presented that demonstrates how to create the necessary launch vehicle reliability models and integrate them with the performance and cost disciplines. An integrated environment is developed for conceptual design that can rapidly assess thousands of launch vehicle configurations. The design process begins with a feasible launch vehicle configuration and its mission objectives. The performance disciplines, such as trajectory analysis, propulsion, and mass estimation are modeled to include the effects of using different reliability strategies. Reliability models are created based upon the launch vehicle configuration. Engine reliability receives additional attention because engines are historically one of the leading causes of launch vehicle failure. Additionally, the reliability of the propulsion subsystem changes dynamically when a launch vehicle design includes engine out capability. Cost estimating techniques which use parametric models are employed to capture the dependencies on system cost of increasing launch vehicle reliability. Uncertainty analysis is included within the cost and reliability disciplines because of the limited historical database for launch vehicles. Optimization is applied within the integrated design environment to find the best launch vehicle configuration based upon a particular weighting of cost and reliability. The results show that both the Saturn V and future launch vehicles could be optimized to be significantly cheaper, be more reliable, or have a compromise solution by illustrating how cost and reliability are coupled with vehicle configuration changes.

ITC/USA 2006 Conference Proceedings / The Forty-Second Annual International Telemetering Conference and Technical Exhibition / October 23-26, 2006 / Town and Country Resort & Convention Center, San Diego, California / XX launch vehicle has been flying onboard video system which includes video cameras, data compression devices and channel switch device for the second Chinese spaceflight. The camera is a PAL analog camera that been sampled and compressed by compression device. The compressed digital video data is combined with telemetry data into the telemetry radio channel. Lighting is provided by sunlight, or a light has been equipped when sunlight is unavailable. IRIG-B timing is used to correlate the video with other vehicle telemetry. The video system’s influences to the vehicle flight have been decreased to minimum.

ITC/USA 2006 Conference Proceedings / The Forty-Second Annual International Telemetering Conference and Technical Exhibition / October 23-26, 2006 / Town and Country Resort & Convention Center, San Diego, California / The National Aeronautics and Space Administration is developing a new launch vehicle to replace the Space Shuttle. The Crew Launch Vehicle (CLV) will be a combination of new design hardware and heritage Apollo and Space Shuttle hardware. The current CLV configuration is a 5 segment solid rocket booster First Stage and a new Upper Stage design with a modified Apollo era J-2 engine. The current schedule has an Ascent Development Test Flight (ADFT-0) with a First Stage and a dummy structurally identical, but without engine, Upper Stage. The ADFT-0 test results will determine if there will be multiple ADFT flights. There will be a minimum of two test flights with a full complement of flight hardware. After the completion of the test flights, the first manned flight to the International Space Station is scheduled for late 2014. To verify the CLV’s design margins a developmental flight instrumentation (DFI) system is needed. The DFI system will collect environmental and health data from the various CLV subsystems’ and either transmit it to the ground or store it onboard for later evaluation on the ground. The CLV consists of 4 major elements: the First Stage, the Upper Stage, the Upper Stage Engine and the integration of these elements together. It is anticipated that each of CLV’s elements will have some version of DFI. This paper will discuss a conceptual DFI design for each element and also of an integrated CLV DFI system.

International Telemetering Conference Proceedings / October 25-28, 1999 / Riviera Hotel and Convention Center, Las Vegas, Nevada / Collecting and analyzing vehicle performance data is an essential part of the launch process. Performance data is used to determine mission success. Performance data also provides essential feedback to the launch vehicle design engineers. This feedback can be used to improve the overall vehicle design and thereby improve the probability of a successful launch. Various Telemetry products are used to gather and process critical information on board launch vehicles. Data is transmitted by RF links to fixed or mobile receiving stations. These Telemetry products are ruggedized for the extreme launch environments. This paper discusses the use of video telemetry as a means of providing launch vehicle performance data.

International Telemetering Conference Proceedings / October 25-28, 1999 / Riviera Hotel and Convention Center, Las Vegas, Nevada / Launch Vehicle tracking is indispensable due to the fact that wayward vehicles must be destroyed lest they cause loss of life and/or damage to property. Launch Vehicle tracking data is also useful in assessing vehicle performance and mission success. Cincinnati Electronics (CE) has developed a Global Positioning Satellite (GPS) Receiver/Transmitter Unit (RTU), specifically for use with launch vehicles. The CE GPS RTU was flown as an experiment on the Missile Technology Demonstration (MTD) flight at White Sands Missile Range (WSMR). This paper provides an overview of CE’s GPS RTU and provides the results of CE’s GPS RTU MTD-3 flight performance.

ITC/USA 2007 Conference Proceedings / The Forty-Third Annual International Telemetering Conference and Technical Exhibition / October 22-25, 2007 / Riviera Hotel & Convention Center, Las Vegas, Nevada / Telemetry Prognostics is Failure Prediction using telemetry for launch vehicle and satellite space flight equipment to stop launch failures, launch pad delays, satellite infant mortalities and satellite on orbit failures. This technology characterizes telemetry behaviors that are latent, transient, and go undetected by the most experienced engineering personnel and software diagnostic tools during integration and test, launch operations and on orbit activities stopping launch pad delays, launch failures, infant mortalities and on orbit failures. Telemetry prognostics yield a technology with state-of-the-art innovative techniques for determining critical on-board equipment remaining useful life taking into account system states, attitude reorientations, equipment usage patterns, failure modes and piece part failure characteristics to increase the reliability, usability, serviceability, availability and safety of our nation’s space systems.

Launch vehicle development programs have experienced significant difficulties in achieving first flight. Optimism during the initiation of these complex programs, coupled with the innovative nature of the technologies they employ, has resulted in a long list of programs unable to remain within the national means. A recent example of this challenge is the Constellation program which was canceled in 2011 due to excessive cost overruns and schedule slippage. The budgetary constraints currently placed on NASA's Space Launch System (SLS) highlights the need for a greater emphasis on affordability. Where affordability is defined in this research as the ability to remain under the mandated funding curve for all points in a system's life cycle while simultaneously meeting schedule goals given that performance requirements are met. The proposed research aims to address the gap between current practices and an affordability-centric design approach by capturing manufacturing technology effects on the affordability of the baseline vehicle concept. Historically, cost overruns and schedule slippages escalate once production begins and are only truly realized at the first launch of a system. These trends, based upon systems which leveraged traditional materials and processes, suggest a shortcoming in the ability of current practices to assess manufacturing implications during the early design phases. The advent of advanced materials and the new process required to fabricate parts from them, further challenges these practices, and threaten to exacerbate the already excessive overruns experience once production begins. Manufacturing technologies, such as composite materials, automated fabrication processes, and the use of stiffener concepts, can no longer be considered independently. This observation leads to the conclusion that improvements in vehicle affordability can only be realized by bringing manufacturing information forward into the Conceptual Design phase. The goal of this research is to support the development of affordable launch vehicles by quantitatively capturing the effects of manufacturing technology selection during Conceptual Design. A manufacturing influenced design methodology is combined with established techniques of time-phasing and risk propagation to evaluate the expected affordability of a launch vehicle baseline concept. The method is benchmarked against expected performance and affordability trends established in literature. The experiments used to build this methodology provide interesting insight into the excess risk typically carried into Preliminary Design due to a lack of the temporal nature of cost. Fundamental implications include the notion that the most expensive candidate (i.e. the highest total cost) does not correspond to the candidate with the highest annual cost insurance. Furthermore, the assessment of risk — within the traditional total cost domain — by overlaying vertical constraints onto uncertainty distributions results in the inclusion of many unaffordable candidates. The final chapter of this thesis applies the method to a relevant launch vehicle, the Exploration Upper Stage (EUS) of the SLS Block IB, which is currently in its Conceptual Design phase. This chapter compares two viable candidate manufacturing technologies based on affordability criteria established herein. The application of this methodology provides the decision maker with a significant amount of information previously unavailable and affords her additional degrees of freedom regarding appropriate Design, Development, Testing, Evaluation, and Production (DDTE&P) planning. This will ultimately enable the selection of an affordable vehicle baseline which will be robust to uncertainty in congress-appropriated funding and thus circumvent risks associated with government program cancellation.

International Telemetering Conference Proceedings / October 28-31, 1996 / Town and Country Hotel and Convention Center, San Diego, California / An S-band telemetry transmitter has been developed for Expendable Launch Vehicles (ELV's) that can downlink data through NASA's Tracking and Data Relay Satellite System (TDRSS). The transmitter operates in the 2200 to 2300 MHz range and provides a number of unique features to achieve optimum performance in the launch vehicle environment: · Commandable QPSK or BPSK modulation format. · Data rates up to 10 Mbps. · Commandable concatenated coding provides superior link performance. · Premodulation filtering produces excellent spectral containment characteristics. · Phase noise of less than 3 degrees rms is maintained through launch and ascent vibration profiles. · A 30 watt nominal RF output power provides a robust RF link. · Two RF antenna output ports with commandable selection of all power out to either port or power split evenly between ports. · Operating modes and conditions of the unit can be monitored through a number of bilevel and analog outputs. · A ruggedized mechanical design provides a reliable communications link for launch vehicle environments.

ITC/USA 2011 Conference Proceedings / The Forty-Seventh Annual International Telemetering Conference and Technical Exhibition / October 24-27, 2011 / Bally's Las Vegas, Las Vegas, Nevada / When embarking into the design of a new launch vehicle, engineering models of expected vehicle performance are always generated. While many models are well established and understood, some models contain design features that are only marginally known. Unfortunately, these analytical models produce uncertainties in design margins. The best way to answer these analytical issues is with vehicle level testing. The National Aeronautics and Space Administration respond to these uncertainties by using a vehicle level system called the Development Flight Instrumentation, or DFI. This DFI system can be simple to implement, with only a few measurements, or it may be a sophisticated system with hundreds of measurement and video, without a recording capability. From experience with DFI systems, DFI never goes away. The system is renamed and allowed to continue, in most cases. Proper system design can aid the transition to future data requirements. This paper will discuss design features that need to be considered when developing a DFI system for a launch vehicle. It will briefly review the data acquisition units, sensors, multiplexers and recorders, telemetry components and harnessing. It will present a reasonable set of requirements which should be implemented in the beginning of the program in order to start the design. It will discuss a simplistic DFI architecture that could be the basis for the next NASA launch vehicle. This will be followed by a discussion of the "experiences gained" from a past DFI system implementation, such as the very successful Ares I-X test flight. Application of these design considerations may not work for every situation, but they may direct a path toward success or at least make one pause and ask the right questions.

In this work the conditions for which a (0,90,90,0,0,90)s BMI-carbon fiber laminate will initiate transverse microcracking are determined for the fabrication of a cryogenic fuel tank for use in a Reusable Launch Vehicle (RLV). This is accomplished using a quadratic interaction criterion failure analysis on the total stress state at possible launch conditions. There are three major sources of stress, that is, thermal residual stress, internal pressure stress, and applied load stress, that are evaluated at the launch stage to determine the total stress state. To assess the accuracy of the analysis the well known X-33 cryogenic fuel tank failure was analyzed as an example. The results of the X-33 example show that the analysis accurately portrays the failure of the X-33 and provides evidence that the analysis can be used to provide reliable conditions for the initiation of microcracking. The final result of this study is a range of launch conditions that can be used without the initiation of microcracking and a limiting range of conditions that cause complete microcracking throughout the laminate.