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Saturn 5 Dynamic Launch Decision Criteria model provides instantaneous criteria, derived from the parametric behavior of a complex system such as a space launch vehicle plus its payload, for the decision making of launch management personnel.
At the Kennedy Space Center (KSC) in Florida, processing facilities are being built and activated to support the processing, checkout, and launch of Space Station elements. The generic capability of these facilities will be utilized to support resupply missions for payloads, life support services, and propellants for the 30-year life of the program. Special Ground Support Equipment (GSE) is being designed for Space Station hardware special handling requirements, and a Test, Checkout, and Monitoring System (TCMS) is under development to verify that the flight elements are ready for launch. The facilities and equipment used at KSC, along with the testing required to accomplish the mission, are described in detail to provide an understanding of the complexity of operations at the launch site. Assessments of hardware processing flows through KSC are being conducted to minimize the processing flow times for each hardware element. Baseline operations plans and the changes made to improve operations and reduce costs are described, recognizing that efficient ground operations are a major key to success of the Space Station.
The successful demonstration of powered flight on Mars by Ingenuity has led to the development of next generation Martian rotorcraft. As technology advances from a technology demonstrator to possible expanded planetary science investigations, the potential to fly a high payload carrying rotorcraft relies on the capability to analyze and validate the structural integrity of the vehicle in addition to the rotor aerodynamic performance. The Mars Science Helicopter (MSH), a next generation hexacopter concept, is one example of a proposed vehicle that would perform science independently of a land-based vehicle. To ensure the feasibility and viability of MSH mission performance, it is critical to mature the structural design for these vehicle concepts to bridge the gap between the best practices of the spacecraft and aircraft communities. MSH was analyzed under operational and 26G quasi-static launch loads with different rotor arm and fuselage designs. Structural analysis results showed that rotor arms could be stiffened for significantly improved structural performance with minimal mass penalty, while the rotorcraft frame exhibited a more complex relationship between mass and launch load stresses due to the mass-dependent launch loads.
The successful demonstration of powered flight on Mars by Ingenuity has led to the development of next generation Martian rotorcraft. As technology advances from a technology demonstrator to possible expanded planetary science investigations, the potential to fly a high payload carrying rotorcraft relies on the capability to analyze and validate the structural integrity of the vehicle in addition to the rotor aerodynamic performance. The Mars Science Helicopter (MSH), a next generation hexacopter concept, is an example of a proposed vehicle that would perform science investigations independent of a land-based vehicle. To ensure the feasibility and viability of MSH mission performance, it is critical to mature the structural design of vehicle concepts to bridge the gap between the best practices of the spacecraft and aircraft communities. MSH was analyzed under operational and 26G quasi-static launch loads with different rotor arm and frame designs. Structural analysis results showed that rotor arms could be stiffened for significantly improved structural performance with minimal mass penalty, while the rotorcraft frame exhibited a more complex relationship between mass and launch load stresses due to the mass-dependent launch loads.
The use of spinning tethers to transfer payloads from low earth orbit (LEO) to geosynchronous earth orbit (GEO) has previously been considered for payload masses up to 4000 kg (4 MT). The construction of the solar power station requires a transfer of 22,568 MT per year from LEO to GEO. This is envisioned to be carried out in payload units of 20 MT or 40 MT, which implies a frequency of 1188 or 594 flights per year, respectively. We could say from the outset that the use of spinning tethers for such large payloads at such high launch frequencies does not appear promising. This is inherent in the principles of spinning tether transfer, which we will briefly sketch below. Somewhat different scenarios are possible, but the basic physics remains the same. We consider only a single stage from LEO to GTO tether system, since the complexity involved in phasing the launches, docking, and spinups for a two-stage system for so many payloads rules out a two-stage system, in our opinion. The payload must first be launched to LEO, where it docks with the tether launch platform and is connected to the tether. The tether (tens of kilometers long) is then deployed with the payload upward. In order to give the payload the velocity necessary to launch it into a geosynchronous transfer orbit (GTO), i.e., to impart the required Av, the tethered system must be spun up about the center of mass of the tether-platform-payload system. The two end masses (platform and payload) are driven to rotate about the center of mass of the tethered system. Both the final rotational velocity and the phasing of the tether spin have to be controlled so that payload is in the vertically up position at the perigee of the LEO and with the velocity required to achieve the GTO when it is released at that point. Upon release, the payload then goes into GTO, where it again requires an acceleration to reach GEO (circularization of the orbit). The platform goes into a lower orbit, from which it must be raised in order to be at the proper LEO for docking with another payload. For this scheme to make sense at all, the platform must be envisioned as having a solar powered electrical thrust system to regain LEO and to spin up the tethered system. Similarly, solar powered electrical propulsion would be used to circularize the orbit to GEO.
The complex nature of the random aerodynamic input forces associated with transonic buffet and ground winds leads to difficulty in analytical treatment of the response of launch vehicles to these important loading conditions. The application of aeroelastically scaled models in the wind tunnel as a mechanical analog of the matharnatically complex problem requiring solution is described. Scaling relationships involved in this approach are developed and some checks on their validity are presented. Model design and support techniques are discussed, and some typical results given. Some problems arising from the increased size of launch vehicles are discussed.
Spacecraft are typically designed with a primary focus on weight in order to meet launch vehicle performance parameters. However, for pressurized and/or man-rated spacecraft, it is also necessary to have an understanding of the vehicle operating environments to properly size the pressure vessel. Proper sizing of the pressure vessel requires an understanding of the space vehicle's life cycle and compares the physical design optimization (weight and launch "cost") to downstream operational complexity and total life cycle cost. This paper will provide an overview of some major environmental design drivers and provide examples for calculating the optimal design pressure versus a selected set of design parameters related to thermal and environmental perspectives. In addition, this paper will provide a generic set of cracking pressures for both positive and negative pressure relief valves that encompasses worst case environmental effects for a variety of launch / landing sites. Finally, several examples are included to highlight pressure relief set points and vehicle weight impacts for a selected set of orbital missions.
The analysis of air launched vehicles carrying payloads to low earth orbit is complex. The integration of the aircraft flight phase and the launch vehicle flight phase requires the integration of systems functioning from different energy relationships. System exergy balance provides a relationship to integrate the efficiency of the aircraft launcher and the efficiency of the launch vehicle in a single, integrated system assessment. The exergy balance of the aircraft flight phase can be calculated with the launch vehicle included as part of the vehicle mass. As exergy balance allows for separate of the masses, different velocities, and different propulsion systems to continue the mission analysis after separation through payload orbital insertion and aircraft landing. This paper presents the initial assessment of the aircraft launch phase up to separation of the launch vehicle. Showing the method of integration afforded by the system exergy balance relationship.
Titan/Centaur TC-4 was launched from the Eastern Test Range, Complex 41, at 05:22 PM EDT on Wednesday, August 20, 1975. This was the second operational flight of the newest NASA unmanned launch vehicle. The spacecraft was the Viking A, the first of two orbiting and landing missions to Mars planned for the 1975 Martian launch opportunity. The objective of the launch phase of the mission, to inject the Viking spacecraft onto the planned transfer orbit to Mars, was successfully accomplished. This report presents a review of the launch vehicle system flight data.
Titan/Centaur TC-3 was launched from the Eastern Test Range, Complex 41, at 02:34 PM EDT on Tuesday, September 9, 1975. This was the third operational flight of the newest NASA unmanned launch vehicle. The spacecraft was the Viking B, the second of two orbiting and landing missions to Mars planned for the 1975 Martian launch opportunity. The objective of the launch phase of the mission, to inject the Viking spacecraft onto the planned transfer orbit to Mars, was successfully accomplished. A review of the launch vehicle system flight data is presented.
Given the complex structural dynamics, challenging ascent performance requirements, and rigorous flight certification constraints owing to its manned capability, the NASA Space Launch System (SLS) launch vehicle requires a proven thrust vector control algorithm design with highly optimized parameters to robustly demonstrate stable and high performance flight. On its development path to preliminary design review (PDR), the stability of the SLS flight control system has been challenged by significant vehicle flexibility, aerodynamics, and sloshing propellant dynamics. While the design has been able to meet all robust stability criteria, it has done so with little excess margin. Through significant development work, an adaptive augmenting control (AAC) algorithm previously presented by Orr and VanZwieten, has been shown to extend the envelope of failures and flight anomalies for which the SLS control system can accommodate while maintaining a direct link to flight control stability criteria (e.g. gain & phase margin). In this paper, the work performed to mature the AAC algorithm as a baseline component of the SLS flight control system is presented. The progress to date has brought the algorithm design to the PDR level of maturity. The algorithm has been extended to augment the SLS digital 3-axis autopilot, including existing load-relief elements, and necessary steps for integration with the production flight software prototype have been implemented. Several updates to the adaptive algorithm to increase its performance, decrease its sensitivity to expected external commands, and safeguard against limitations in the digital implementation are discussed with illustrating results. Monte Carlo simulations and selected stressing case results are shown to demonstrate the algorithm's ability to increase the robustness of the integrated SLS flight control system.
Given the complex structural dynamics, challenging ascent performance requirements, and rigorous flight certification constraints owing to its manned capability, the NASA Space Launch System (SLS) launch vehicle requires a proven thrust vector control algorithm design with highly optimized parameters to provide stable and high-performance flight. On its development path to Preliminary Design Review (PDR), the SLS flight control system has been challenged by significant vehicle flexibility, aerodynamics, and sloshing propellant. While the design has been able to meet all robust stability criteria, it has done so with little excess margin. Through significant development work, an Adaptive Augmenting Control (AAC) algorithm has been shown to extend the envelope of failures and flight anomalies the SLS control system can accommodate while maintaining a direct link to flight control stability criteria such as classical gain and phase margin. In this paper, the work performed to mature the AAC algorithm as a baseline component of the SLS flight control system is presented. The progress to date has brought the algorithm design to the PDR level of maturity. The algorithm has been extended to augment the full SLS digital 3-axis autopilot, including existing load-relief elements, and the necessary steps for integration with the production flight software prototype have been implemented. Several updates which have been made to the adaptive algorithm to increase its performance, decrease its sensitivity to expected external commands, and safeguard against limitations in the digital implementation are discussed with illustrating results. Monte Carlo simulations and selected stressing case results are also shown to demonstrate the algorithm's ability to increase the robustness of the integrated SLS flight control system.
A series representation is presented of the dynamic behavior of incompressible non-viscous liquids contained in partially filled elastic shells of revolution. Using a series in which each term satisfies exactly Laplace's equation, the liquid portion of the problem becomes an integral equation on the wetted wall of the shell. Physical insight required in selecting the terms of the series is such that selection may be readily programed on a computer. Liquid-shell vibration interaction is investigated by modifying an existing shell of revolution vibration program with the series expansion of the liquid pressure loading on the wetted shell wall. Application of the modified program is demonstrated on a hemispherical tank and a realistic scaled model of a launch vehicle tank having a more complex geometry. Comparison of the present analysis with a NASTRAN finite element analysis for a hemispherical tank, and comparison with test results for a launch vehicle tank shows excellent agreement and rapid convergence of the series representation of the liquid to the true solution.
Historically, the federal government has been the predominant purchaser of space launches in the United States. The government met its needs through purchase of hardware and services. It also provided the infrastructure necessary to conduct launch operations through federal launch ranges, both military and NASA. Under this model, the government had the complete ownership, responsibility, liability, and expense for launch activities. As the commercial space sector grew, there emerged a corresponding growth in demand for launch range services. However, the expense and complexity of activities has thus far deterred a rapid rise in the establishment of purely commercial launch sites. In this context, purely commercial is defined as "without benefit of capabilities provided by the federal government." Consistent with the Commercial Space Launch Act, in recent years NASA and the Air Force have supported commercial launches from government launch ranges on a cost-reimbursable, non-interference basis. In this mode the commercial launch service providers contract with the government to provide services including use of facilities, tracking and data services, and range safety. As the commercial market projections began to show significant opportunities for economic development, several states established spaceports to provide the services necessary to meet these projected commercial needs. In 1997, NASA agreed to the establishment of the Virginia Space Flight Center (VSFC) at the Wallops Flight Facility. Under this arrangement, NASA agreed to allow Virginia Commercial Space Flight Authority (VCSFA) to construct facilities on NASA property and agreed to provide services in accordance with the Space Act of 1958 and the Commercial Space Launch Act of 1984 (as amended) to support VSFC launch customers. The relationship between NASA and VCSFA, however, has evolved beyond a customer supplier relationship. A partnership relationship has emerged which pairs the strengths of the established NASA test range and the state-sponsored, commercial launch facility provider, in an attempt to satisfy the needs for flexible, low-cost access to space. Furthermore, the future of the NASA/Wallops Test Range is closely linked with the success of VCSFA in promoting commercial launches from Wallops. This paper will describe the changing paradigm of the federal launch range and the unique aspects of the NASA/Wallops Facility relationship with VCSFA. Discussion will include institutional cost-sharing, business development and marketing, joint educational programs, and strategic planning.
Computerized simulation model of a launch vehicle/ground support equipment system optimizes assembly, checkout, and operation of the system. The model is used to determine performance parameters in three phases or modes - /1/ systems optimization techniques, /2/ operation analysis methodology, and /3/ systems effectiveness analysis technique.
The prediction of flight loads and their potential reduction, using various control system logics for the space shuttle vehicles, is discussed. Some factors not found on previous launch vehicles that increase the complexity are large lifting surfaces, unsymmetrical structure, unsymmetrical aerodynamics, trajectory control system coupling, and large aeroelastic effects. These load-producing factors and load-reducing techniques are analyzed.
This research has focused on improving the mathematical models of the air-breathing propulsion systems, which can be mated with the rocket engine model and incorporated in trajectory optimization codes. Improved engine simulations provided accurate representation of the complex cycles proposed for advanced launch vehicles, thereby increasing the confidence in propellant use and payload calculations. The versatile QNEP (Quick Navy Engine Program) was modified to allow treatment of advanced turboaccelerator cycles using hydrogen or hydrocarbon fuels and operating in the vehicle flow field.
Platforms for polar and 28.5 deg orbits were studied to determine the platform requirements and characteristics necessary to support the science objectives. Large platforms supporting the Earth-Observing System (EOS) were initially studied. Co-orbiting platforms were derived from these designs. Because cost estimates indicated that the large platform approach was likely to be too expensive, require several launches, and generally be excessively complex, studies of small platforms were undertaken. Results of these studies show the small platform approach to be technically feasible at lower overall cost. All designs maximized hardware inheritance from the Space Station program to reduce costs. Science objectives as defined at the time of these studies are largely achievable.