Bernard Laub

Adaptive Deployable Entry and Placement Technology (ADEPT): A Feasibility Study for Human Missions to Mars

The present paper describes an innovative, semi-rigid, mechanically deployable hypersonic decelerator system for human missions to Mars. The approach taken in the present work builds upon previous architecture studies performed at NASA, and uses those findings as the foundation to perform analysis and trade studies. The broad objectives of the present work are: (i) to assess the viability of the concept for a heavy mass (landed mass ≈40 mT) Mars mission through system architecture studies; (ii) to contrast it with system studies previously performed by NASA; and (iii) to make the case for a Transformable Entry System Technology. The mechanically deployable concept at the heart of the proposed transformable architecture is akin to an umbrella, which in a stowed configuration meets launch requirements by conforming to the payload envelope in the launch shroud, and when deployed in earth orbit forms a large aerosurface designed to provide the necessary aerodynamic forces upon entry into the Martian atmosphere. The aerosurface is a thin skin draped over high-strength ribs; the thin skin or fabric with flexible material serves as the thermal protection system, and the ribs serve as the structure. A four-bar linkage mechanism allows for a reorientation of the aerosurface during aerocapture or during the entry and descent phases of atmospheric flight, thus providing a capability to navigate and control the vehicle and make possible precision landing. The actuators and mechanisms that are used to deploy the aerosurface are multi-functional in that they also allow for reorienting the

Mars Science Laboratory Entry Capsule Aerothermodynamics and Thermal Protection System

The mars science laboratory (MSL) spacecraft is being designed to carry a large rover (> 800 kg) to the surface of Mars using a blunt-body entry capsule as the primary decelerator. The spacecraft is being designed for launch in 2009 and arrival at mars in 2010. The combination of large mass and diameter with non-zero angle-of-attack for MSL will result in unprecedented convective heating environments caused by turbulence prior to peak heating. Navier-Stokes computations predict a large turbulent heating augmentation for which there are no supporting flight data1 and little ground data for validation. Consequently, an extensive experimental program has been established specifically for MSL to understand the level of turbulent augmentation expected in flight. The experimental data support the prediction of turbulent transition and have also uncovered phenomena that cannot be replicated with available computational methods. The result is that the flight aeroheating environments predictions must include larger uncertainties than are typically used2 for a mars entry capsule. Finally, the thermal protection system (TPS) being used for MSL has not been flown at the heat flux, pressure, and shear stress combinations expected in flight, so a test program has been established to obtain conditions relevant to flight. This paper summarizes the aerothermodynamic definition analysis and TPS development, focusing on the challenges that are unique to MSL.

Monte Carlo Analysis for Spacecraft Thermal Protection System Design

This paper demonstrates a Monte Carlo analysis technique to establish margins on sizing a thermal protection system and identifies the chief sources of uncertainty in the material response modeling. Monte Carlo sensitivity and uncertainty studies are performed for the thermal protection systems of the Stardust Sample Return Capsule, Mars Exploration Rovers, and X-37 wing leading edge, using the Fully Implicit Ablation and Thermal response code. The computation results are presented and discussed in detail. It shows that a Monte Carlo approach provides more insight than the traditional Root-Sum-Square method into the relationship between the thickness margin of a thermal protection system and the probability of maintaining the temperature of the underlying material within specified requirements. Nomenclature erfi = inverse error function N = total number of samples R = random number x = input parameters y = output of interests σ = standard deviation * σ = mean x

Defining Ablative Thermal Protection System Margins for Planetary Entry Vehicles

After over 50 years of spaceflight, the calculation of the necessary margin on the thermal protection system of an entering spacecraft remains a largely ad-hoc (in a tactical sense) process governed by engineering judgment. Over the past several years, NASA has explored a more methodical technique through the application of statistically derived factors to various elements of the design. The current paper first reviews the history of margin application for several prior NASA missions ranging from Apollo to the Mars Exploration Rovers, then discusses a new probabilistic margin process, which was developed for the Mars Science Laboratory and Orion programs. The process is very data-dependent, and requires validation via a carefully crafted and executed test program including both ground and flight data. Examples of the application of this probabilistic process are given for Mars Pathfinder, Stardust, and the Mars Science Laboratory. This process is still a work in progress, and the paper discusses some of the key elements of it, along with potential ways by which the overall methodology could be improved. The eventual goal is the development of a rigorous link between TPS thermal margin process and the corresponding contribution to the overall heatshield reliability.

New TPS materials for aerocapture

Many planetary probes, landers and aerocapture concepts are conceived for entry trajectories where peak convective heat flux is in the range 150–400 W/cm2. This may be too severe an environment for either reusable or low-density ablative materials. The high-density ablatives will work in such environments but the associated TPS weight requirements can be prohibitive. Unfortunately, there are few, if any, well-understood materials that provide reliable, predictable ablative performance for the 150–400 W/cm2 regime while still providing weight efficient TPS solutions. JPL has recently been evaluating an Earth aerocapture demonstration at an entry velocity of ≈10 km/s. TPS thickness and areal weight requirements were determined for current ablative TPS candidates (e.g., SLA-561V, PICA) where, for the large integrated heat loads associated with aerocapture, it is shown that some of these materials may not provide efficient thermal protection. A new concept, employing a low catalycity, high emissivity coating on a low-density ceramic tile is evaluated and shown to provide significant benefits for such missions.