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NASA NTRS1958

The Explorer Rocket Research Program

Since September of 1956, nine Jupiter-C missiles have been launched from the firing pad at Cape Canaveral. The first Jupiter-C firing tested the propulsion system, air frame, and guidance components of the missile, and the second and third firings tested a model of the Jupiter nose cone under realistic re-entry conditions. The remaining six Jupiter-C missiles were used as the launching vehicles for EXPLORER satellites I through VI (Fig. 1). Of the six satellite firings, EXPLORERs I, III, and IV achieved satisfactory orbits. The Jupiter-C missile was designed and developed as a joint program under the technical direction of the Jet Propulsion Laboratory and the Army Ballistic Missile Agency. The Jet Propulsion Laboratory developed the three high-speed stages, and the Army Ballistic Missile Agency handled the development, construction, and operation of the first-stage booster rocket and the guidance system. Many other organizations have contributed to the success of the EXPLORER satellite program, most notably the State University of Iowa, the Air Force Cambridge Research Center, and the satellite tracking teams of the Vanguard organization.

Robillard, G.↗

NASA NTRS2016

Results of Outdoor to Indoor Propagation Measurements from 5-32GHz

The demand for wireless services has increased exponentially in the last few years and shows no signs of slowing in the near future. In order for the next generation wireless to provide seamless access, whether the user is indoors or out, a thorough understanding and validation of models describing the impact of building entry loss (BEL) is required. This information is currently lacking and presents a challenge for most system designers. For this reason empirical data is needed to assess the impact of BEL at frequencies that are being explored for future mobile broadband applications This paper present the results of measurements of outdoor-to-indoor propagation from 5-32 GHz in three different buildings. The first is a newer building that is similar in construction to modern residential home. The second is an older commercial office building. The last building is a very new commercial office building built using modern green building techniques. These three buildings allow for the measurement of propagation losses through both modern and older materials; such as glass windows and exterior block and siding. Initial results found that at particular spatial locations the BEL could be less than 1dB or more than 70dB with free space losses discounted (this is likely influenced by multipath). Additionally, it was observed that the PDF distributions of a majority of the measurements trended toward log-normal with means and standard deviations ranging from 8-38dB and 6-14dB, respectively.

building entry loss↗

NASA NTRS2022

Reconstructed Performance of the Mars InSight Lander’s Supersonic Parachute & Comparison with the Phoenix Lander

On November 26th, 2018 the Mars InSight landersuccessfully touched down at Elysium Planitia. InSight’s Entry,Descent, and Landing sequence included an 11.8-m supersonicallydeployed Disk-Gap-Band parachute that was a largelybuild-to-print version of the parachute used to successfully landthe Phoenix lander a decade earlier. This paper describes thereconstructed performance of InSight’s supersonic parachuteat Mars, highlighting the differences and similarities with theperformance of Phoenix’s parachute. Measurements from theonboard inertial measurement unit along with pre-launch measurementsof the parachute system and spacecraft, assumptionsabout the vehicle’s aerodynamics, and models for the Martianatmosphere were used to reconstruct the spacecraft’s trajectoryand the parachute system’s performance. The reconstruction resultswere compared against pre-flight predictions. Reconstructionof the InSight trajectory leading up to parachute deploymentshowed that the vehicle trimmed in a lift-down orientationduring entry and thus experienced greater deceleration thanexpected by most pre-flight simulations. This led to parachutedeployment conditions that diverged from the nominal preflightpredictions. The parachute was mortar-deployed at aMach number of approximately 1.5, below the nominal preflightexpectation of 1.66. The approximate dynamic pressureat mortar fire was between 518 Pa and 546 Pa, which agreedwell with pre-flight expectations. The mortar system performednominally, and the system’s deployment (0.755 s) and inflation(1.1 s) times were in line with pre-flight modeling. The peakinflation load was 45 kN, well below the parachute’s 67 kNdesign limit load. Following deployment of the parachute,the vehicle’s rotational rates and the dynamics of the systemwere in excellent agreement with pre-flight expectations. Theperformance of the InSight parachute system was also foundto agree well with that of the Phoenix parachute system, asexpected given the similarities between the two systems.

Karlgaard, Chris↗

NASA NTRS2013

Parachute Models Used in the Mars Science Laboratory Entry, Descent, and Landing Simulation

An end-to-end simulation of the Mars Science Laboratory (MSL) entry, descent, and landing (EDL) sequence was created at the NASA Langley Research Center using the Program to Optimize Simulated Trajectories II (POST2). This simulation is capable of providing numerous MSL system and flight software responses, including Monte Carlo-derived statistics of these responses. The MSL POST2 simulation includes models of EDL system elements, including those related to the parachute system. Among these there are models for the parachute geometry, mass properties, deployment, inflation, opening force, area oscillations, aerodynamic coefficients, apparent mass, interaction with the main landing engines, and off-loading. These models were kept as simple as possible, considering the overall objectives of the simulation. The main purpose of this paper is to describe these parachute system models to the extent necessary to understand how they work and some of their limitations. A list of lessons learned during the development of the models and simulation is provided. Future improvements to the parachute system models are proposed.

Cruz, Juan R.↗

NASA NTRS2007

Ablation Modeling of Ares-I Upper State Thermal Protection System Using Thermal Desktop

The thermal protection system (TPS) for the Ares-I Upper Stage will be based on Space Transportation System External Tank (ET) and Solid Rocket Booster (SRB) heritage materials. These TPS materials were qualified via hot gas testing that simulated ascent and re-entry aerothermodynamic convective heating environments. From this data, the recession rates due to ablation were characterized and used in thermal modeling for sizing the thickness required to maintain structural substrate temperatures. At Marshall Space Flight Center (MSFC), the in-house code ABL is currently used to predict TPS ablation and substrate temperatures as a FORTRAN application integrated within SINDA/G. This paper describes a comparison of the new ablation utility in Thermal Desktop and SINDA/FLUINT with the heritage ABL code and empirical test data which serves as the validation of the Thermal Desktop software for use on the design of the Ares-I Upper Stage project.

Sharp, John R.↗

NASA NTRS2006

Global Reference Atmospheric Models, Including Thermospheres, for Mars, Venus and Earth

This document is the viewgraph slides of the presentation. Marshall Space Flight Center's Natural Environments Branch has developed Global Reference Atmospheric Models (GRAMs) for Mars, Venus, Earth, and other solar system destinations. Mars-GRAM has been widely used for engineering applications including systems design, performance analysis, and operations planning for aerobraking, entry descent and landing, and aerocapture. Preliminary results are presented, comparing Mars-GRAM with measurements from Mars Reconnaissance Orbiter (MRO) during its aerobraking in Mars thermosphere. Venus-GRAM is based on the Committee on Space Research (COSPAR) Venus International Reference Atmosphere (VIRA), and is suitable for similar engineering applications in the thermosphere or other altitude regions of the atmosphere of Venus. Until recently, the thermosphere in Earth-GRAM has been represented by the Marshall Engineering Thermosphere (MET) model. Earth-GRAM has recently been revised. In addition to including an updated version of MET, it now includes an option to use the Naval Research Laboratory Mass Spectrometer Incoherent Scatter Radar Extended Model (NRLMSISE-00) as an alternate thermospheric model. Some characteristics and results from Venus-GRAM and Earth-GRAM thermospheres are also presented.

Justh, Hilary L.↗

NASA NTRS2018

New Metadata Capabilities within NASA's Common Metadata Repository (CMR)

This talk will convey the new capabilities and features of the UMM-Variables and the UMM-Services metadata model within the Common Metadata Repository (CMR) and how community engagement through the ESDIS Standards Office (ESO) review shaped the new versions of the models. The Unified Metadata Model (UMM) provides a common metadata model to unify legacy systems (i.e. GCMD (Global Change Master Directory), ECHO (Earth Observing System (EOS) Clearinghouse)) with new systems (i.e. CMR). The rationale and migration process of the Service Entry Resource Format (SERF) to the UMM-S will also be conveyed. The talk will conclude with discussing issues and lessons learned from the review process and how future reviews will be conducted to ensure a more targeted, meaningful review with a faster turn-around time for triage and implementation.

Earth Observing System↗

NASA NTRS1973

Ocean recovery of Shuttle Solid Rocket Boosters.

Cost effective recovery of the expended Space-Shuttle Solid Rocket Boosters (SRB) from the ocean will result in significant overall cost savings to the Space Shuttle Program. The ocean recovery mission begins with the dispatching of the recovery team to the predicted splashdown area. The SRBs, drogue parachutes and main parachutes must be tracked, located, retrieved, and transported to land where they will be refurbished and recycled for reuse. Trade studies to be conducted will consider the recovery mission requirements and weigh the advantages, disadvantages and costs of various candidate recovery systems. Major parameters effecting the selection of the final system will ensure that the system will meet overall objectives. Large- and small-scale SRB model testing has been conducted to establish characteristics of SRBs during water entry, floating free and under tow.

Junker, L. J.↗

NASA NTRS2011

Thermal protection system ablation sensor

An isotherm sensor tracks space vehicle temperatures by a thermal protection system (TPS) material during vehicle re-entry as a function of time, and surface recession through calibration, calculation, analysis and exposed surface modeling. Sensor design includes: two resistive conductors, wound around a tube, with a first end of each conductor connected to a constant current source, and second ends electrically insulated from each other by a selected material that becomes an electrically conductive char at higher temperatures to thereby complete an electrical circuit. The sensor conductors become shorter as ablation proceeds and reduced resistance in the completed electrical circuit (proportional to conductor length) is continually monitored, using measured end-to-end voltage change or current in the circuit. Thermocouple and/or piezoelectric measurements provide consistency checks on local temperatures.

Gorbunov, Sergey↗

NASA NTRS2017

Estimates for the Aerodynamic Coefficients of Ringsail and Disk-Gap-Band Parachutes Operating on Mars

The creation of flight dynamics simulations is often part of the design, development, and flight operations of missions that include entry, descent, and landing (EDL) in their concept of operation. Such flight dynamics simulations include models for many of the EDL sub-systems. For EDL systems that include a parachute, models of the parachute’s aerodynamic characteristics must be created for inclusion in the flight dynamics simulations. Models are presented for the aerodynamic coefficients (i.e., C(D),C(Tot),C(T),C(N), and C(m,SLCP)) of Supersonic Ringsail (SSRS) and Disk-Gap-Band (DGB) parachutes as functions of total porosity, λ(T), Mach number, M, and total angle of attack, λ(T) (when necessary). The source aerodynamic coefficients data used for creating these models were obtained from a wind tunnel test of subscale parachutes. In this wind tunnel test, subscale parachutes of both parachute types were fabricated from two different fabrics with very different permeabilities. By varying the fabric permeability, while maintaining the parachute geometry constant, it was possible to vary λ(T). The fabric permeability test data necessary for the calculation of λ(T) were obtained from samples of the same fabrics used to fabricate the subscale parachutes. Although the models for the aerodynamic coefficients are simple polynomial functions of λ(T) and M, they are capable of producing good reproductions of the source data. The (λ(T), M) domains over which these models are applicable are clearly defined. The models and their domains are applicable to flight operations on Mars. The models presented in this report are not unique – other models could be created from the source data. To make the creation of alternative numerical models, the aerodynamic coefficients source data are included in this report.

Cruz, Juan R.↗

NASA NTRSDate not supplied

Coupling CFD and Material Response for Analysis of Mars Entry

In computing the response of an ablating thermal protection system during atmospheric entry, the aerothermal environment and material response are generally computed separately with a blowing correction term in the material response model to account for the blowing of char and pyrolysis gases [1]. In this work, we apply a coupled approach in which pyrolysis blowing gases, computed in the PATO material response code [2], are used with a blowing boundary condition in the DPLR hypersonic CFD code [3]. This leads to an iterative method in which blowing products from PATO are input into DPLR to update surface heating estimates. The full iterative method, with the addition of radiative heating estimates using the NEQAIR radiation solver [4], is shown in Fig. 1. The method is demonstrated on a sphere case with the environment and material properties based on the Mars Science Laboratory entry. Future work includes utilizing this method in computing full 3D material response during the Mars 2020 entry.

Heat Transfer↗

NASA NTRSDate not supplied

Coupling CFD and Material Response for Analysis of Mars Entry

In computing the response of an ablating thermal protection system during atmospheric entry, the aerothermal environment and material response are generally computed separately with a blowing correction term in the material response model to account for the blowing of char and pyrolysis gases. In this work, we apply a coupled approach in which pyrolysis blowing gases, computed in the PATO material response code, are used with a blowing boundary condition in the DPLR hypersonic CFD code. This leads to an iterative method in which blowing products from PATO are input into DPLR to update surface heating estimates. The full iterative method, with the addition of radiative heating estimates using the NEQAIR radiation solver, is shown in Fig. 1. The method is demonstrated on a sphere case with the environment and material properties based on the Mars Science Laboratory entry. Future work includes utilizing this method in computing full 3D material response during the Mars 2020 entry.

Ablation↗

NASA NTRSDate not supplied

Progress Toward Evaluating Roughness-Resolved Aerothermodynamic Simulations for the Mars Sample Return Earth Entry System

This work pursues the efficient Navier-Stokes simulation, using NASA’s standard LAURA and DPLR simulation tools, of the flowfield and convective heating resulting from the patterned-roughness surface of the Mars Sample Return (MSR) Earth Entry System (EES). This type of simulation has been infeasible for past NASA missions. However, for EES, the combination of a patterned roughness heatshield approximated with a sinusoidal pattern, the zero degrees angle of attack nominal flow, and the supersonic boundary layer edge Mach number, all combine to make these roughness resolved grid simulations feasible. The patterned roughness and zero degrees angle of attack nominal flow allow the simulation domain to be reduced to a thin axial flowfield slice that captures a minimum of single period of the roughness pattern. With a periodic boundary condition applied along the boundaries of this slice, simulation on the reduced domain is identical to the full domain but with significant cost saving. The supersonic boundary layer edge Mach number prevents the upstream influence, which allows the smooth stagnation region to be frozen. This freezing of the stagnation region flow is required because of the numerical instability resulting from the tight spacing of the grid in the axial direction as the stagnation point is approached. Applying the developed analysis approach to a peak heating EES case results in a roughness augmentation factor for the convective heating of 1.32, which agrees well with correlations that predict values between 1.30 and 1.37. Analysis approach developed in this work provides the framework for future studies of turbulence modeling impacts and the influence of an ablating surface.

Earth Entry System↗

NASA NTRSDate not supplied

Parallelized Carbon Oxidation and Surface Recession Model in Direct-Simulation Monte Carlo Code, SPARTA

Ablative thermal protection system (TPS) materials for spacecraft are composites that often consist of a carbon-based reinforcement and a polymeric matrix. During Earth re-entry, they endure high-temperature oxidation and surface recession. Oxidation is an important mechanism for ablation, sometimes leading to the weakening, spallation, or failure of the oxidized fibers at the surface and in the char layer. However, more details are required including accurate material properties of the fiber microstructure, whether the fibers recede homogenously or localized at pits, and the role of pyrolysis outgassing in order to evaluate the role oxidation plays in the degradation and failure mechanisms of these materials. In this work, we demonstrate a parallelizable oxidation-driven ablation model developed for detailed, large-scale simulations in the DSMC code SPARTA. We also develop robust algorithms which handle the conservation of the surface state after an ablation step. Finally, we verify our model for both simple and more complex chemistry as well as microstructures with reaction statistics, oxidation depth calculations, and detailed visualization of oxidation-induced material recession at the microscale.

V Arias↗

NASA NTRS2018

Multi-Scale Modeling of Low-Density Carbon-Phenolic Ablators

Protecting a spacecraft during atmospheric entry is one of highest risk factors that needs to be mitigated during design of a space exploration mission. At entry speeds from space, air turns into high-temperature plasma, and spacecraft Thermal Protection Systems (TPS) are needed to protect the vehicle payload. Modern successful material architectures of spacecraft shields use a porous carbon fiber substrate impregnated with phenolic as an ablator material. In the lecture, efforts to build a Predictive Material Modeling framework for porous ablators from micro-scale to macro-scale will be presented. Several numerical methods and techniques will be summarized that use voxelized images to compute geometrical properties of the porous substrate. These computed properties include porosity, specific surface area and tortuosity that are otherwise indirectly measured through experimental techniques. Direct simulation Monte Carlo (DSMC), a particle-based method for approximating the Boltzmann equation, is used to compute the permeability coefficient of the porous substrate based on its digitized representation. The method computes the flow within the microstructure, where the size of the pores may approach the mean-free-path of the flow. Finally, a high-fidelity model implemented in PATO (Porous-material Analysis Toolbox) is discussed, and some examples of ablative material response are presented, including for the first time 3D simulations of the full tiled heat shield for the Mars Science Laboratory (MSL) capsule.

Ablators↗

NASA NTRS2019

Technology Readiness Assessment for HEEET TPS

The HEEET project was conceived to develop a heatshield with a high performance ablative thermal protection material that can withstand the extreme entry environment produced as a result of rapid deceleration during high speed entry into Venus, Saturn, Uranus or higher speed entry into Earth's atmosphere. Successful maturation of HEEET supports future New Frontiers and Discovery AO's, as well as Flagship and directed missions in the longer term. In addition, HEEET has the potential to evolve and to support re-entry to Earth, for missions such as Mars Sample Return.The primary goal of the HEEET Project was to develop an ablative TPS heat-shield based on woven TPS technology to Technology Readiness Level (TRL) 6. Key evidence to support the TRL evaluation includes: Demonstration of reproducible manufacturing of a dual layer material over a range of thicknesses and integrated on to a heatshield engineering test unit at a scale that is applicable to near term Discovery as the highest priority and future NF missions as secondary priority set of missions. Demonstration of predictable and stable performance of the dual layer TPS over a range of entry environments that are applicable to near term Discovery and NF missions of interest to SMD.Includes completion of coupon arc jet and laser testing and development of a mid-fidelity thermal response model that correlates with test results. Demonstration of flight heatshield system design for a range of sizes and loads that are relevant to near term Discovery and NF missions of interest to SMD. Includes completion of structural testing to validate analytic thermal/structural models and development of a material property database. Includes structural testing of a ~1m Engineering Test Unit under relevant entry loads.

Gage, Peter↗

NASA NTRS1995

Large-Scale Structure of Magnetospheric Plasma

Recent investigations of magnetospheric plasma structure are summarized under the broad categories of empirical models, transport across boundaries, formation, and dynamics of the plasma sheet. This report reviews work in these areas during the period 1991 to 1993. Fully three-dimensional empirical models and simulations have become important contributors to our understanding of the magnetospheric system. Some new structural concepts have appeared in the literature: the 'entry boundary' and 'geo-pause', the plasma sheet 'region 1 vortices', the 'low-energy layer', the 'adia-baticity boundary' or 'wall region', and a region in the tail to which we refer as the 'injection port'. Traditional structural concepts have also been the subject of recent study, notably the plasmapause, the magnetopause, and the plasma sheet. Significant progress has been made in understanding the nature of plasma sheet formation and dynamics, but the acceleration of electrons to high energy remains somewhat mysterious.

Moore, T. E.↗

NASA NTRSDate not supplied

Thermal Protection Systems and Materials

Thermal protection systems protect space vehicles from aerodynamic heating during entry into planetary atmospheres. The Thermal Protection Materials Branch combines experiment, fabrication, characterization, evaluation, analysis and modeling to support NASA missions and collaborates externally to technology transfer thermal protection materials and tools to commercial or other agency partners.

Modeling↗