ALBERT

Description: Experience with many spacecraft configurations boosted by a variety of launch vehicles indicates that the maximum loads experienced throughout most of the structure are inertial in origin. These loads arise from the dynamic elastic response of the flight vehicle to the transient disturbances of launch and flight, and are highly dependent on the dynamic characteristics of both the spacecraft and the launch vehicle. It has proved to be most advantageous, in the analysis of this critical dependency of loads upon vehicle dynamic properties, to establish a mathematical model in terms of normal mode characteristics. In this way, the vibration behavior of an elastomechanical structure (or substructure) can be described by means of the so-called modal or natural degrees of freedom. The conduct of a mode survey test and the use of a suitably test-verified model in loads analyses is essential to the flight worthiness certification process of space systems. The desirability of such tests is confirmed by the fact that, almost invariably, significant deficiencies in the analytical models are revealed by the results. Therefore, this experimental program was undertaken to determine those properties of a solid-propellant rocket motor (SRM) which are required to characterize a dynamic model. Random ambient-excited accelerations were measured at a series of stations along the motor for the purpose of identifying the motor beam-like stiffnesses in bending, shear, and torsion. From a system identification point of view, it is significant that stiffness properties of a subsystem (the motor) are determined from modes of the full system (motor/stand configuration) using mode shape data of the subsystem only. This contrasts with traditional system identification approaches which rely upon complete system mode shapes.

Description: The Texas A&M Nuclear and Aerospace engineering departments have worked on five different projects for the NASA/USRA Advanced Design Program during the 1987/88 year. The aerospace department worked on two types of lunar tunnelers that would create habitable space. The first design used a heated cone to melt the lunar regolith, and the second used a conventional drill to bore its way through the crust. Both used a dump truck to get rid of waste heat from the reactor as well as excess regolith from the tunneling operation. The nuclear engineering department worked on three separate projects. The NEPTUNE system is a manned, outer-planetary explorer designed with Jupiter exploration as the baseline mission. The lifetime requirement for both reactor and power-conversion systems was twenty years. The second project undertaken for the power supply was a Mars Sample Return Mission power supply. This was designed to produce 2 kW of electrical power for seven years. The design consisted of a General Purpose Heat Source (GPHS) utilizing a Stirling engine as the power conversion unit. A mass optimization was performed to aid in overall design. The last design was a reactor to provide power for propulsion to Mars and power on the surface. The requirements of 300 kW of electrical power output and a mass of less than 10,000 Rg were set. This allowed the reactor and power conversion unit to fit within the Space Shuttle cargo bay.

Description: The design and performance of a arcjet nuclear electric propulsion spacecraft, suitable for use in a space reactor power system (SRPS) flight experiment, are outlined. The vehicle design is based on a 92-kW ammonia arcjet system operating at a specific impulse of 1050 s and an efficiency of 45 percent. The arcjet/gimbal system, power processing unit, and propellant feed system are described. A 100-kW SRPS is assumed and the spacecraft mass is baselined at 5250 kg, excluding the propellant and propellant feed system. A radiation/arcjet efflux diagnostics package is included in the performance analysis. This spacecraft, assuming a Shuttle launch from Kennedy Space Center, can perform a 35-deg inclination change and reach a final orbit of 35,860 km with a 120-day trip time, thus providing a four-month active load for the SRPS. Alternatively, a Titan IV launch could provide a mass margin of 120 kg to a 1000km, 58-deg final orbit in 74 days.

Description: The major requirements and guidelines that affect the manned Space Station configuration and the power systems are explained. The evolution of the Space Station power system from the NASA program development feasibility phase through the current preliminary design phase is described. Several early station concepts are described and linked to the present concept. The recently completed phase B tradeoff study selections of photovoltaic system technologies are described. The present solar dynamic and power management and distribution systems are also summarized for completeness.

Description: A new concept for a Space Station power system is proposed which reduces the drag effect of the solar panels and eliminates eclipsing by the Earth. The solar generator is physically separated from the Space Station, and power transmitted to the station by a microwave beam. The power station can thus be placed high enough that drag is not a significant factor. For a resonant orbit where the ratio of periods s:p is a ratio of odd integers, and the orbital planes nearly perpendicular, an orbit can be chosen such that the line of sight is never blocked if the lower orbit has an altitude greater than calculatable mininum. For the 1:3 resonance, this minimum altitude is 0.5 r(e). Finally, by placing the power station into a sun-synchronous orbit, it can be made to avoid shadowing by the Earth, thus providing continuous power.

Description: The rudiments of a rocket thruster that receives its enthalpy from a remote energy source, a laser, are described. An experimental study is discussed that provides details of the physics for assessing the feasibility of using hydrogen plasmas for accepting and converting this energy to enthalpy. A plasma ignition scheme that uses a pulsed CO2 laser has been developed and the properties of the ignition spark documented, including breakdown intensities in hydrogen. A complete diagnostic system, capable of determining plasma temperature and the plasma absorptivity for subsequent steady-state absorption of a high-power CO2 laser beam, is developed and demonstrative use is discussed for the preliminary case study, a 2 atm laser-supported argon plasma.

Description: A two-year feasibility and test program to solve the problem of unburned confined hydrogen at the Vandenberg Space Launch Complex Six (SLC-6) during Space Shuttle Main Engine (SSME) firings is discussed. A novel steam inerting design was selected for development. Available sound suppression water is superheated to flash to steam at the duct entrance. Testing, analysis, and design during 1987 showed that the steam inerting system (SIS) solves the problem and meets other flight-critical system requirements. The SIS design is complete and available for installation at SLC-6 to support shuttle or derivative vehicles.

Description: The development of probabilistic structural analysis methods is a major part of the SSME Structural Durability Program and consists of three program elements: composite load spectra, probabilistic finite element structural analysis, and probabilistic structural analysis applications. Recent progress includes: (1) the effects of the uncertainties of several factors on the HPFP blade temperature pressure and torque, (2) the evaluation of the cumulative distribution function of structural response variables based on assumed uncertainties on primitive structural variables, and (3) evaluation of the failure probability. Collectively, the results obtained demonstrate that the structural durability of critical SSME components can be probabilistically evaluated.

Description: Vehicle/engine analysis studies have identified the High/Dual Mixture Ratio O2/H2 Engine cycle as a leading candidate for an advanced Single Stage to Orbit (SSTO) propulsion system. This cycle is designed to allow operation at a higher than normal O/F ratio of 12 during liftoff and then transition to a more optimum O/F ratio of 6 at altitude. While operation at high mixture ratios lowers specific impulse, the resultant high propellant bulk density and high power density combine to minimize the influence of atmospheric drag and low altitude gravitational forces. Transition to a lower mixture ratio at altitude then provides improved specific impulse relative to a single mixture ratio engine that must select a mixture ratio that is balanced for both low and high altitude operation. This combination of increased altitude specific impulse and high propellant bulk density more than offsets the compromised low altitude performance and results in an overall mission benefit. Two areas of technical concern relative to the execution of this dual mixture ratio cycle concept are addressed. First, actions required to transition from high to low mixture ratio are examined, including an assessment of the main chamber environment as the main chamber mixture ratio passes through stoichiometric. Secondly, two approaches to meet a requirement for high turbine power at high mixture ratio condition are examined. One approach uses high turbine temperature to produce the power and requires cooled turbines. The other approach incorporates an oxidizer-rich preburner to increase turbine work capability via increased turbine mass flow.