Jose A. Santos

Initial Assessment of Mars Science Laboratory Heatshield Instrumentation and Flight Data

The Mars Science Laboratory (MSL) Entry Descent and Landing Instrumentation (MEDLI) suite on MSL entry vehicle heatshield has successfully returned pressure, temperature, and thermal protection system (TPS) ablation data acquired during entry. This paper provides an initial assessment of MEDLI thermal instrumentation data that is comprised of in-depth temperatures in the TPS made of Phenolic-Impregnated Carbon Ablator (PICA). Temperatures are measured in-depth at seven different locations on the surface. The thermal sensor plugs are also characterized in arc jet facilities to quantify measurement uncertainties and biases. The assessment of flight data provides key insights into boundary layer transition to turbulence, surface recession, turbulent heating augmentation, stagnation point and apex laminar heating, and in-depth thermal response. A preliminary comparison with model results highlights inadequacies in our predictive capability. The peak temperature measured by near surface thermocouples was found to be 1049 C in the vicinity of apex region. Initial estimate of peak surface temperature with nominal model settings is about 1575 C. The peak heat flux was found to be on the leeside of the vehicle as predicted, but its value is sensitive to the recession model.

Development and Application of a TPS Ablation Sensor for Flight

Isotherm following sensors, which are sometimes referred to as recession sensors for historical reasons, have been resurrected and modified to quantify low density thermal protection system (TPS) material response. The modifications presented herein have dramatically improved reliability, accuracy, and data quality. Extensive research of TPS response on atmospheric entry vehicles for the Department of Defense was conducted in the 1960s, requiring the development of TPS recession sensor technology. The only NASA Earth entry vehicles whose TPS was heavily instrumented are the Apollo test vehicles and the Space Shuttle Orbiter, and only a limited number of sensors have been installed within the TPS of planetary atmospheric reentry probes. Despite advances in computational techniques, relatively large TPS sizing margins are still carried due to uncertainty in the flow environments, uncertainty in TPS material response, and the inability to fully simulate the flight environment in ground tests. TPS instrumentation provides critical engineering and science data that feeds back to aerothermal analyses and physics-based material response models. The acquisition of in-flight heatshield performance data will lead to improvements in the predictions obtained from these models. This, in turn, will greatly benefit future NASA missions by supporting the quantification of TPS risk from design with flight data.

Isotherm Sensor Calibration Program for Mars Science Laboratory Heat Shield Flight Data Analysis

Seven instrumented sensor plugs were installed on the Mars Science Laboratory heat shield in December 2008 as part of the Mars Science Laboratory Entry, Descent, and Landing Instrumentation (MEDLI) project. These sensor plugs contain four in-depth thermocouples and one Hollow aErothermal Ablation and Temperature (HEAT) sensor. The HEAT sensor follows the time progression of a 700 C isotherm through the thickness of a thermal protection system (TPS) material. The data can be used to infer char depth and, when analyzed in conjunction with the thermocouple data, the thermal gradient through the TPS material can also be determined. However, the uncertainty on the isotherm value is not well defined. To address this uncertainty, a team at NASA Ames Research Center is carrying out a HEAT sensor calibration test program. The scope of this test program is described, and initial results from experiments conducted in the laboratory to study the isotherm temperature of the HEAT sensor are presented. Data from the laboratory tests indicate an isotherm temperature of 720 C 60 C. An overview of near term arc jet testing is also given, including preliminary data from 30.48cm 30.48cm PICA panels instrumented with two MEDLI sensor plugs and tested in the NASA Ames Panel Test Facility. Forward work includes analysis of the arc jet test data, including an evaluation of the isotherm value based on the instant in time when it reaches a thermocouple depth.

Post-flight Analysis of Mars Science Laboratory Entry Aerothermal Environment and Thermal Protection System Response

The Mars Science Laboratory rover landed at Gale Crater on August 5 th , 2012. The rover was protected from the extreme heating environments of Martian atmospheric entry by an ablative heatshield. This tiled Phenolic Impregnated Carbon Ablator heatshield was instrumented with a suite of sensors that monitored the in-depth ablator temperature response and the surface pressure at discrete locations. This paper presents a comparison of the flight data with post-entry analysis at the discrete sensor locations. From the flight data, we postulate that the heatshield experienced roughness-induced turbulent transition due to roughness elements around the heatshield tile and sensor plugs. We find that the analytical ablator material model performs well and can be used directly with the in-depth temperature data. Finally, we assess the performance of the ablation sensors, and predict the bondline temperature rise. The flight data from the instrumentation, along with the successful landing of the rover, confirms the performance of the heatshield and the conservative heatshield design and margins process.

Proposed Analysis Process for Mars Science Laboratory Heat Shield Sensor Plug Flight Data

The Mars Science Laboratory (MSL) mission is scheduled to enter the Martian atmosphere in August 2012. Aboard the heatshield is the MSL Entry Descent and Landing Instrumentation (MEDLI) system that includes a series of embedded sensor plugs to measure in-depth response of the thermal protection system (TPS). The general objectives of the MEDLI system are to assess the TPS performance and reconstruct the aerothermal environment experienced during entry. Some specific objectives, such as measuring TPS temperature, can be addressed with direct measurements. Other objectives, such as determining surface heating, must be inferred using measurements combined with analytical tools. This paper describes the specific objectives, the expected sensor responses to the entry environment based on aerothermal and material response simulations, and the reconstruction analysis process being developed for the flight data.

Mars Science Laboratory Heat Shield Instrumentation and Arc Jet Characterization

The Mars Science Laboratory (MSL) Entry Descent and Landing Instrumentation (MEDLI) suite on MSL entry vehicle heatshield has returned pressure, temperature, and thermal protection system (TPS) performance data acquired during entry. This paper presents performance and characterization data of the MEDLI Integrated Sensor Plug (MISP) embedded in Phenolic-Impregnated Carbon Ablator (PICA) heatshield. The sensor is characterized in arc jet facilities at MSL flight relevant conditions. The performance of the Hollow aErothermal Ablation and Temperature (HEAT) sensor in tracking a moving temperature isotherm through the thickness is evaluated. A close agreement between HEAT sensor depth and measured char depth in arc jet samples is also found. The growth of a fence due to Room-Temperature Vulcanizing (RTV) bonding agent around MISP plugs, which has significant impact on aerothermal reconstruction, is also quantified. The data presented will be used for improved reconstruction of the aerothermal environment and TPS response using MISP flight data.

Volumetric Heat Flux Characterization Experiments in the Interaction Heating Facility at NASA Ames

Characterization of plasma flow conditions is an important link for any material test performed inside the Interaction Heating Facility (IHF) at NASA Ames Research Center. In a recent arc jet test campaign titled IHF 212, a series of sensors was deployed to study the volumetric heat flux distribution with the goal of improving the facility’s current ability to measure heat flux and to help understand physical phenomena through comparison with computational fluid dynamics. Further, potential asymmetry in the flow could be studied. In the test series, a pitot probe and a null point calorimeter were traversed across the arc jet flow stream; and a hemispherical slug calorimeter was used in a stagnation test configuration to give an additional data point at centerline. In addition, two 20° blunt-nosed, water-cooled, copper wedge assemblies, each holding ten slug calorimeters, were used to obtain shear flow heat flux measurements. Tests were conducted at facility maximum and minimum power settings. The shear flow heat flux measurements reveal a region of higher heat flux towards the West end of the facility by up to 9%. The heat flux data from the null point calorimeter sweeps show the same bias with the peak heat flux located 1.5 cm ± 0.5 cm from the centerline position. The pressure data from the pitot probe sweeps adds a third corroboration of the same observation as its peak is also towards the West end of the facility.

Challenge of developmental flight instrumentation for orion exploration flight test 1: Potential benefit of wireless technology for future Orion missions

The National Aeronautics and Space Administration has developed the Multi Purpose Crew Vehicle (MPCV) named Orion to readiness for its first test flight in 2014. The spacecraft is unique in its design to support deep space missions. In order to successfully man-rate the vehicle a series of two unmanned flight tests are scheduled (Exploration Flight Test 1, Exploration Mission 1), followed by the first crewed flight Exploration Mission 2. Accomplishing Flight Test Objectives (FTOs) for the flight tests requires a dedicated instrumentation system that will measure dynamic response of all vehicle subsystem performance during critical phases of the EFT-1 and EM-1 missions. These include structural response, and Thermal Protection System (TPS) response during atmospheric reentry. A suite of avionics data acquisition system electronics along with associated cabling supports the large number of channels. This paper shall discuss the architecture of the EFT-1 data system designed to meet FTOs at low risk, and the potential effect on mass with new technology using wireless applications. Also described will be an architecture that could decrease mass by a factor 10, or more with and without wireless capability.

Null Point Calorimeter Sweeps with Comparisons to Thermal FEA Model Predictions

This paper presents experimental data of null point calorimeter sweeps conducted in the 60 MW Interaction Heating Facility at NASA Ames Research Center. The test setup, test conditions, and current data reduction methods are described. Finite element analysis model computations of the null point cavity temperature are also presented to demonstrate the need for application of an inverse heat conduction model to the null point calorimeter geometry. Simulations have been performed with the commercial finite element analysis software package COMSOL Multiphysics using a two-dimensional axisymmetric geometry. Two distinct IHF test conditions—one at 433 W/cm 2 and the other at 802 W/cm 2 , as measured with a 10.16 cm diameter hemispherical slug calorimeter— are considered. The heat flux distribution on the exposed surface of the null point calorimeter has been numerically computed with CFD. This heating distribution is normalized by the area- averaged heat flux over the flat face of the null point sensor. The heat flux is scaled at each time step and imposed as a boundary condition in the finite element model in order to compute the temperature at the null point cavity. The computed results differ from the measured temperature by up to 35%—a deviation sufficiently large to encourage further characterization of the methods by which heat flux is computed from the null point temperature data.

Mars 2020 Entry, Descent and Landing Instrumentation (MEDLI2)

This paper will introduce Mars Entry Descent and Landing Instrumentation (MEDLI2) on NASAs Mars2020 mission. Mars2020 is a flagship NASA mission with science and technology objectives to help answer questions about possibility of life on Mars as well as to demonstrate technologies for future human expedition. Mars2020 is scheduled for launch in 2020. MEDLI2 is a suite of instruments embedded in the heatshield and backshell thermal protection systems of Mars2020 entry vehicle. The objectives of MEDLI2 are to gather critical aerodynamics, aerothermodynamics and TPS performance data during EDL phase of the mission. MEDLI2 builds up the success of MEDLI flight instrumentation on Mars Science Laboratory mission in 2012. MEDLI instrumentation suite measured surface pressure and TPS temperature on the heatshield during MSL entry into Mars. MEDLI data has since been used for unprecedented reconstruction of aerodynamic drag, vehicle attitude, in-situ atmospheric density, aerothermal heating, transition to turbulence, in-depth TPS performance and TPS ablation. [1,2] In addition to validating predictive models, MEDLI data has highlighted extra margin available in the MSL forebody TPS, which can potentially be used to reduce vehicle parasitic mass. MEDLI2 expands the scope of instrumentation by focusing on quantities of interest not addressed in MEDLI suite. The type the sensors are expanded and their layout on the TPS modified to meet these new objectives. The paper will provide key motivation and governing requirements that drive the choice and the implementation of the new sensor suite. The implementation considerations of sensor selection, qualification, and demonstration of minimal risk to the host mission will be described. The additional challenges associated with mechanical accommodation, electrical impact, data storage and retrieval for MEDLI2 system, which extends sensors to backshell will also be described.