Alicia Dwyer Cianciolo

Atmospheric risk assessment for the Mars Science Laboratory Entry, Descent, and Landing system

In 2012, the Mars Science Laboratory (MSL) mission will pioneer the next generation of robotic Entry, Descent, and Landing (EDL) systems, by delivering the largest and most capable rover to date to the surface of Mars.12 As with previous Mars landers, atmospheric conditions during entry, descent, and landing directly impact the performance of MSLs EDL system. While the vehicles novel guided entry system allows it to “fly out” a range of atmospheric uncertainties, its trajectory through the atmosphere creates a variety of atmospheric sensitivities not present on previous Mars entry systems and landers. Given the missions stringent landing capability requirements, understanding the atmosphere state and spacecraft sensitivities takes on heightened importance.

Human Mars lander design for NASA's evolvable mars campaign

Landing humans on Mars will require entry, descent, and landing capability beyond the current state of the art. Nearly twenty times more delivered payload and an order of magnitude improvement in precision landing capability will be necessary. To better assess entry, descent, and landing technology options and sensitivities to future human mission design variations, a series of design studies has been initiated. This paper describes the results of the first design study in the series of studies to be completed in 2016 and includes system and subsystem design details including mass and power estimates for a lander design using the Hypersonic Inflatable Aerodynamic Decelerator (HIAD) entry technology. Future design activities in this series will focus on other entry technology options.

Human Mars Entry, Descent, and Landing Architecture Study Overview

Landing humans on Mars will require entry, descent and landing (EDL) capability beyond the current state of the art. Nearly twenty times more delivered payload and an order of magnitude improvement in precision landing capability will be necessary. Several EDL technologies capable of meeting the human class payload delivery requirements are being considered. The EDL technologies considered include low lift-to-drag vehicles like Hypersonic Inflatable Aerodynamic Decelerators (HIAD), Adaptable Deployable Entry and Placement Technology (ADEPT), and mid range lift-to-drag vehicles like rigid aeroshell configurations. To better assess EDL technology options and sensitivities to future human mission design variations, a series of design studies has been conducted. The design studies incorporate EDL technologies with conceptual payload arrangements defined by the Evolvable Mars Campaign to evaluate the integrated system with higher fidelity than have been performed to date. This paper describes the results of the design studies for a lander design using the HIAD, ADEPT and rigid shell entry technologies and includes system and subsystem design details including mass and power estimates. This paper will review the point design for three entry configurations capable of delivering a 20 t human class payload to the surface of Mars.

Human Mars mission design study utilizing the adaptive deployable entry and placement technology

The Adaptive Deployable Entry and Placement Technology (ADEPT) is being considered as an entry, descent and landing (EDL) system to enable Human Mars class missions. ADEPT is a mechanically deployable decelerator that makes use of a 3 d woven carbon fabric as both heat shield and primary structure. The Human Mars Mission design study is focused, in part, on assessing the viability of ADEPT and identifying technical challenges, operational constraints, and critical risk mitigation activities. Study inputs included definition of the ground rules and assumptions, associated mission timelines and high level functional requirements. These inputs enabled the clarification of the concept of operations along with the design constraints and environments. Subsystem trades, mass sizing and integrated flight performance assessments enabled determination of a feasible mission architecture. Key outputs from the design study include a parametric mass model, driving requirements, key performance parameters and critical risks. These findings enable us to determine strategies for technical maturation and risk mitigation that can be assessed against resource and programmatic constraints to aid in advanced planning for human exploration of Mars.

Impacts of launch vehicle fairing size on human exploration architectures

Human missions to Mars, particularly to the Martian surface, are grand endeavors that place extensive demands on ground infrastructure, launch capabilities, and mission systems. The interplay of capabilities and limitations among these areas can have significant impacts on the costs and ability to conduct Mars missions and campaigns. From a mission and campaign perspective, decisions that affect element designs, including those based on launch vehicle and ground considerations, can create effects that ripple through all phases of the mission and have significant impact on the overall campaign. These effects result in impacts to element designs and performance, launch and surface manifesting, and mission operations. In current Evolvable Mars Campaign concepts, the NASA Space Launch System (SLS) is the primary launch vehicle for delivering crew and payloads to cis-lunar space. SLS is currently developing an 8.4m diameter cargo fairing, with a planned upgrade to a 10m diameter fairing in the future. Fairing diameter is a driving factor that impacts many aspects of system design, vehicle performance, and operational concepts. It creates a ripple effect that influences all aspects of a Mars mission, including: element designs, grounds operations, launch vehicle design, payload packaging on the lander, launch vehicle adapter design to meet structural launch requirements, control and thermal protection during entry and descent at Mars, landing stability, and surface operations. Analyses have been performed in each of these areas to assess and, where possible, quantify the impacts of fairing diameter selection on all aspects of a Mars mission. Several potential impacts of launch fairing diameter selection are identified in each of these areas, along with changes to system designs that result. Solutions for addressing these impacts generally result in increased systems mass and propellant needs, which can further exacerbate packaging and flight challenges. This paper presents the results of the analyses performed, the potential changes to mission architectures and campaigns that result, and the general trends that are more broadly applicable to any element design or mission planning for human exploration.

Large Mass, Entry, Descent and Landing Sensitivity Results for Environmental, Performance, and Design Parameters

Landing on Mars has been a challenging task. Past NASA missions have shown resilience to increases in spacecraft mass by scaling back requirements such as landing site altitude, landing site location and arrival time. Knowledge of the partials relating requirements to mass is critical for mission designers to understand so that the project can retain margin throughout the process. Looking forward to new missions that will land 1.5 metric tons or greater, the current level of technology is insufficient, and new technologies will need to be developed. Understanding the sensitivity of these new technologies to requirements is the purpose of this paper.

Effects of Atmospheric Phenomena on Mars Science Laboratory Entry Performance

At Earth during entry the shuttle has experienced what has come to be known as potholes in the sky or regions of the atmosphere where the density changes suddenly. Because of the small data set of atmospheric information where the Mars Science Laboratory (MSL) parachute deploys, the purpose of this study is to examine the effect similar atmospheric pothole characteristics, should they exist at Mars, would have on MSL entry performance. The study considers the sensitivity of entry design metrics, including altitude and range error at parachute deploy and propellant use, to pothole like density and wind phenomena.

Mission and design sensitivities for human Mars landers using Hypersonic Inflatable Aerodynamic Decelerators

Landing humans on Mars is one of NASAs long term goals. NASAs Evolvable Mars Campaign (EMC) is focused on evaluating architectural trade options to define the capabilities and elements needed to sustain human presence on the surface of Mars. The EMC study teams have considered a variety of in-space propulsion options and surface mission options. Understanding how these choices affect the performance of the lander will allow a balanced optimization of this complex system of systems problem. This paper presents the effects of mission and vehicle design options on lander mass and performance. Beginning with Earth launch, options include fairing size assumptions, co-manifesting elements with the lander, and Earth-Moon vicinity operations. Capturing into Mars orbit using either aerocapture or propulsive capture is assessed. For entry, descent, and landing both storable as well as oxygen and methane propellant combinations are considered, engine thrust level is assessed, and sensitivity to landed payload mass is presented. This paper focuses on lander designs using the Hypersonic Inflatable Aerodynamic Decelerators, one of several entry system technologies currently considered for human missions.

Sensitivity of simulated Martian atmospheric temperature to prescribed dust opacity distribution: Comparison of model results with reconstructed data from Mars Exploration Rover missions

We use the Mars Weather Research and Forecasting (MarsWRF) General Circulation Model to simulate the atmospheric structure corresponding to the landing location and time of the Mars Exploration Rovers (MER) Spirit (A) and Opportunity (B) in 2004. The multi-scale capability of MarsWRF facilitates high-resolution nested model runs centered near the landing site of each of the rovers. Dust opacity distributions based on measurements by Thermal Emission Spectrometer (TES) aboard the Mars Global Surveyor (MGS) spacecraft, and those from an old version of the Mars Climate Database (MCD v3.1 released in 2001) are used to study the sensitivity of the model temperature profile to variations in the dust prescription. The reconstructed Entry, Descent, and Landing (EDL) data from the rover missions are used for comparisons. We show that the model using dust opacity from TES limb and nadir data for the year of MER EDL, Mars Year 26 (MY26), yields temperature profiles in closer agreement with the reconstructed data than the pre-launch EDL simulations and models using other dust opacity specifications. The temperature at 100 Pa from the model (MY26) and the reconstruction are within 5° K. These results highlight the role of vertical dust opacity distribution in determining the atmospheric thermal structure. Similar studies involving data from past missions and models will be useful in understanding the extent to which atmospheric variability is captured by the models and in developing realistic pre-flight characterization required for future lander missions to Mars.

Autonomous Aerobraking Development Software: Phase One Performance Analysis at Mars, Venus, and Titan

When entering orbit about a planet or moon with an appreciable atmosphere, instead of using only the propulsion system to insert the spacecraft into its desired orbit, aerodynamic drag can be used after the initial orbit insertion to further decelerate the spacecraft. Several past NASA missions have used this aerobraking technique to reduce the fuel required to deliver a spacecraft into a desired orbit. Aerobraking was first demonstrated at Venus with Magellan in 1993 and then was used to achieve the science orbit of three Mars orbiters: Mars Global Surveyor in 1997, Mars Odyssey in 2001, and Mars Reconnaissance Orbiter in 2006. Although aerobraking itself reduces the propellant required to reach a final low period orbit, it does so at the expense of additional mission time to accommodate the aerobraking operations phase (typically 3-6 months), a large mission operations staff, and significant Deep Space Network (DSN) coverage. By automating ground based tasks and analyses associated with aerobraking and moving these onboard the spacecraft, a flight project could save millions of dollars in operations staffing and DSN costs (Ref. 1).