Glenn T. Tsuyuki

Uncertainty Quantification in Estimating Critical Spacecraft Component Temperatures

A method for quantifying uncertainty in conceptual-level design via a computationally- efficient probabilistic method is presented. The investigated method is applied to estimating the maximum-expected temperature of several critical components on a spacecraft. The variables of the design are first classified and assigned appropriate probability density functions. To characterize the thermal control system of the spacecraft, Subset Simulation, an efficient simulation technique originally developed for reliability analysis of civil engineering structures, is used. The results of Subset Simulation are compared with traditional Monte Carlo simulation. The investigated method allows uncertainty in the maximum-expected temperatures to be quantified based on the risk tolerance of the decision maker. For the spacecraft thermal control problem presented, Subset Simulation successfully replicated Monte Carlo simulation results for estimating the maximum-expected temperatures of several critical components yet required significantly less computational effort, in particular for risk-averse decision makers. Nomenclature

The Hardware Challenges for the Mars Exploration Rover Heat Rejection System

The primary objective of the Mars Exploration Rover (MER) 2003 Project focused on the search for evidence of water on Mars. The launch of two identical flight systems occurred in June and July of 2003. The roving science vehicles are expected to land on the Martian surface in early and late January of 2004, respectively. The flight system design inherited many successfully features and approaches from the Mars Pathfinder Mission. This included the use of a mechanically‐pumped fluid loop, known as the Heat Rejection System (HRS), to transport heat from the Rover to radiators on the Cruise Stage during the quiescent trek to Mars. While the heritage of the HRS was evident, application of this system for MER presented unique and difficult challenges with respect to hardware implementation. We will discuss these hardware challenges in each HRS hardware element: the integrated pump assembly, cruise stage HRS, lander HRS, and Rover HRS. These challenges span the entire development cycle including fabrication, as...

Margin Determination in the Design and Development of a Thermal Control System

A method for determining margins in conceptual-level design via probabilistic methods is described. The goal of this research is to develop a rigorous foundation for determining design margins in complex multidisciplinary systems. As an example application, the investigated method is applied to conceptual-level design of the Mars Exploration Rover (MER) cruise stage thermal control system. The method begins with identifying a set of tradable system-level parameters. Models that determine each of these tradable parameters are then created. The variables of the design are classified and assigned appropriate probability density functions. To characterize the resulting system, a Monte Carlo simulation is used. Probabilistic methods can then be used to represent uncertainties in the relevant models. Lastly, results of this simulation are combined with the risk tolerance of thermal engineers to guide in the determination of margin levels. The method is repeated until the thermal engineers are satisfied with the balance of system-level parameter values. For the thermal control example presented, margins for maximum component temperatures, dry mass, maximum power required, schedule, and cost form the set of tradable system-level parameters. Use of this approach for the example presented yielded significant differences between the calculated design margins and the values assumed in the conceptual design of the MER cruise stage thermal control system.

Mars Exploration Rover: Thermal Design is a System Engineering Activity

The Mars Exploration Rovers (MER), were launched in June and July of 2003, respectively, and successfully landed on Mars in early and late January of 2004, respectively. The flight system architecture implemented many successful features of the Mars Pathfinder (MPF) system: A cruise stage that transported an entry vehicle that housed the Lander, which in turn, used airbags to cushion the Rover during the landing event. The initial thermal design approach focused on adopting the MPF design wherever possible, and then concentrating on the totally new Rover thermal design. Despite a fundamentally sound approach, there were several salient lessons learned. Some were due to differences from MPF, while others were caused by other means. These lessons sent a clear message: thermal design continues to be a system engineering activity. In each major flight system assembly, there were excellent examples of this recurring theme. From the cruise stage, the cascading impact of a propulsion fill and drain valve thermal design change after system level test is described. In addition, we present the interesting resolution of the sun sensor head thermal design (bare metal versus white paint). The final implementation went against best thermal engineering practices. For the entry vehicle consisting of the aeroshell and equipment mounted to it, an inertial measurement unit mounted on a shock-isolation fixture presented a particularly difficult design challenge. We initially believed that its operating time would be limited due to its relatively low mass and high power dissipation. We conclude with the evolution of the Rover actuator thermal design where the single-string warm-up heaters were employed. In this instance, fault protection requirements drove the final thermal design implementation, and in the case of Opportunity, proved to be critical for meeting primary mission lifetime.

Mars Exploration Rover Heat Rejection System Performance – Comparison of Ground and Flight Data

Mars Exploration Rover (MER) mission launched two spacecraft to Mars in June and July of 2003 and landed two rovers on Mars in January 2004. A Heat Rejection System (HRS) based on a mechanically pumped single-phase liquid cooling system was used to reject heat from electronics to space during the seven months cruise from Earth to Mars. Even though most of this HRS design was similar to the system used on Mars Pathfinder in 1996, several key modifications were made in the MER HRS design. These included the heat exchanger used in removing the heat from electronics, design of venting system used to vent the liquid prior to Mars entry, inclusion of pressure transducer in the HRS, and the spacecraft radiator design. Extensive thermal/fluids modeling and analysis were performed on the MER HRS design to verify the performance and reliability of the system. The HRS design and performance was verified during the spacecraft system thermal vacuum tests. Based on the analysis and the testing of the HRS system, operations of the HRS during launch, cruise and prior to the Martian entry were developed and implemented. The electronics and radiator temperatures were within the range of the predicted values. The HRS system pressure was maintained at the predicted levels indicating any liquid or gas leakages were within the predicted values. The venting system on both spacecraft performed flawlessly in January 2004 when the pyro-valves in the HRS were actuated before the spacecraft entered the Martian environment. The paper describes the various design modifications made on the MER HRS from that of Mars Pathfinder spacecraft. A description of the flight performance during the seven-month cruise of the spacecraft and a comparison of the performance on the ground and the flight is presented. Any significant deviation in the flight performance will be described.

Mars Exploration Rover Thermal Test Program Overview

In January 2004, two Mars Exploration Rovers (MER) landed on the surface of Mars to begin their mission as robotic geologists. A year prior to these historic landings, both rovers and the spacecraft that delivered them to Mars, were completing a series of environmental tests in facilities at the Jet Propulsion Laboratory. This paper describes the test program undertaken to validate the thermal design and verify the workmanship integrity of both rovers and the spacecraft. The spacecraft, which contained the rover within the aeroshell, were tested in a 7.5 m diameter thermal vacuum chamber. Thermal balance was performed for the near earth (hot case) condition and for the near Mars (cold case) condition. A solar simulator was used to provide the solar boundary condition on the solar array. IR lamps were used to simulate the solar heat load on the aeroshell for the off-sun attitudes experienced by the spacecraft during its cruise to Mars. Each rover was tested separately in a 3.0 m diameter thermal vacuum chamber over conditions simulating the warmest and coldest expected Mars diurnal temperature cycles. The environmental tests were conducted in a quiescent nitrogen atmosphere at a pressure of 8 to 10 Torr. In addition to thermal balance testing, the science instruments on board the rovers were tested successfully in the extreme environmental conditions anticipated for the mission. A solar simulator was not used in these tests.

The Thermal Design Evolution of the Phoenix Robotic Arm

An adequate warm-up heater approach was developed using analysis & previous Mars surface experience. In the worst-cold environment, all the actuators were able to warm to their minimum operating temperature within the required time. In the worst-hot environment, the maximum continuous actuator power dissipation during normal operations was determined. A mechanical thermostat provides best means to cope with a failed-on wrist heater & to ease RA operations planning.