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Hannah S Alpert

Publications and source records attributed to Hannah S Alpert.

10 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS2022

Mars Entry Instrumentation Flight Data and Mars 2020 Entry Environments

On February 18th, 2021, the Mars 2020 entry vehicle delivered the Perseverance rover to the surface of Mars. The entry vehicle carried a set of instrumentation installed on the heatshield and backshell to measure aerodynamic and aerothermal performance, named the Mars Entry, Descent, and Landing Instrumentation 2. This set of instrumentation included pressure transducers, thermocouples, heatflux sensors, and a radiometer, as well as a dedicated sensor support electronics system. All MEDLI2 hardware operated as expected during cruise and entry. MEDLI2 sensors gathered accurate pressure measurements in hypersonic through supersonic regimes to reconstruct vehicle attitude and atmospheric profiles. MEDLI2 on the heatshield sensors indicated that surface temperatures, caused by turbulent heating beginning 70 seconds after entry, remained at or below 1430 °C, while heatshield bondline temperatures rose less than 45 °C. Backshell surface TPS temperatures peaked at 630 °C, which was caused primarily by radiative heating measured by several separate sensors. The MEDLI2 temperature and pressure measurements enabled further detailed characterization of the Mars 2020 entry performance, and the flight dataset will provide a wealth of information for the EDL community and future mission designers.

thermal protection system↗
NASA NTRSDate not supplied

Inverse Estimation of Mars 2020 Entry Aeroheating Environments Using MEDLI2 Flight Data

The Mars Entry, Descent, and Landing Instrumentation 2 (MEDLI2) sensor suite collected heating and pressure data during entry into Mars' atmosphere of the Mars 2020 Perseverance rover. MEDLI2 included thermocouples, heat flux sensors, and pressure transducers on both the heatshield and the backshell. This paper covers the inverse estimation of heatshield and backshell surface heating based on the MEDLI2 Instrumented Sensor Plugs (MISPs), a network of thermocouples embedded in thermal protection system plugs across the aeroshell. Monte Carlo analysis was conducted to assess the sensitivity of the surface heat rate and temperature to uncertainties in thermocouple depth and material properties such as density, specific heat capacity, and thermal conductivity. Data from each MISP was also used to estimate the local time of transition from laminar to turbulent flow at each plug location.

Hannah S Alpert↗
NASA NTRSDate not supplied

Mars 2020 Reconstructed Aerothermal Environments and Design Margins

The Mars 2020 aeroshell's thermal protection system was nearly identical to the Mars Science Laboratory system that successfully completed its mission for the Curiosity rover's landing in 2012. It was predicted that, like Mars Science Laboratory, the Mars 2020 heatshield would experience boundary layer transition and that the thermal protection system would provide sufficient material thickness margins against the aerothermal environments, even after adding radiative heating that was not included for Mars Science Laboratory. The Mars 2020 flight instrumentation suite included heatshield sensors similar to Mars Science Laboratory and new backshell measurements. A full set of surface pressure and in-depth temperature data were collected during atmospheric entry. This paper provides an initial analysis of the flight data and focuses on the reconstructed total surface heat flux inferred from the measured temperatures. Turbulent boundary layer conditions again were observed on the Mars 2020 heatshield, and the temperatures at all eleven heatshield and six backshell thermocouple in-depth sensor locations were well within system capabilities due to favorable entry conditions and conservative uncertainties. New computational fluid dynamics results on the reconstructed entry trajectory are compared to the measured surface pressures and reconstructed total heat fluxes. The predicted heatshield surface pressures at six locations match the data qualitatively and quantitatively well, as expected. Smooth-wall laminar heating predictions prior to boundary layer transition fall above the reconstructed heating on the heatshield. After the observed boundary layer transition time, total heat flux based on algebraic turbulence model calculations generally match the heat flux trends relative to one another. The convective heat flux was predicted to constitute the majority of the total heating. On the backshell, smooth-wall laminar total heat flux predictions generally exceed the reconstructed total heating at six locations. At each of these locations, it is estimated that radiative heating provided almost all of the total heating. The Mars 2020 as-flown aerothermal environments were well below the design levels for all measurement locations due to conservative design assumptions and a stressing design trajectory.

Mars 2020↗
NASA NTRSDate not supplied

MEDLI2: MISP Inferred Aerothermal Environment and Flow Transition Assessment

The Mars Entry, Descent, and Landing Instrumentation 2 (MEDLI2) sensor suite on the Mars2020 mission contained multiple sensors on the aeroshell to measure the aerothermal environment during entry into the Martian atmosphere. These sensors performed superbly and successfully returned forebody and aftbody heating measurements. Analysis of MEDLI2 data indicated flow transitioning from a laminar to turbulent state on the heatshield. No evidence of flow transition was observed on the backshell. Two methods were used to estimate flow transition times on the heatshield: (1) temperature gradient of near-surface thermocouple data and (2) heat flux gradient from an inverse reconstruction approach using thermocouple data and material response modeling. Both methods produced similar transition times with an estimated accuracy of ±1 s. To assess various transition criteria, transition parameters were evaluated at each sensor location using flow field solutions from computational fluid dynamics (CFD) simulations. The idea was to use conservative values inferred from MEDLI2 data as transition criteria for other Mars missions. To test this hypothesis, MEDLI data from the Mars Science Laboratory (MSL) mission was used to compare predicted vs. actual flow transition times. The comparisons suggest smooth wall transition criteria are not well-suited in modeling the rapid progression of a turbulent transition front. Transition criteria containing a roughness element parameter agreed better with the flight data. In summary, critical transition values derived from MEDLI2 data may be used as a starting point in constructing a flow transition model for future Mars missions.

Chun Y Tang↗
NASA NTRSDate not supplied

Overview and Performance of the LOFTID Instrumentation Suite

NASA’s Hypersonic Inflatable Aerodynamic Decelerator (HIAD) is an enabling technology that facilitates atmospheric entry of heavy payloads to planets such as Earth and Mars using a deployable aeroshell. The deployable nature of the HIAD technology allows it to overcome the size constraints imposed on current rigid aeroshell entry systems. This enables use of larger aeroshells resulting in increased entry system performance (e.g. higher payload mass and/or volume, higher landing altitude at Mars). On November 10th, 2022 the Low-Earth Orbit Flight Test of an Inflatable Decelerator (LOFTID) was launched out of Vandenberg Air Force Base as a secondary payload on an Atlas V rocket. After the primary payload was delivered to its orbit, the LOFTID reentry vehicle was inflated, positioned, and then separated to reenter Earth’s atmosphere at a velocity of 8.1km/s, ultimately splashing down safely in the Pacific Ocean. The flight successfully demonstrated a 6m diameter, 70-deg sphere-cone HIAD on a high-energy orbital reentry. This demonstration has provided invaluable fight data essential to characterize the vehicle performance and support the ongoing effort to further scale the HIAD technology to vehicles of 10m in diameter or greater. Aeroshells of this scale are applicable to near-term commercial applications and future NASA robotic and human exploration missions. LOFTID incorporated an extensive instrumentation suite totaling over 150 science measurements. This included thermocouples, total heat flux sensors, and a radiometer to characterize the aeroheating environment and aeroshell thermal response. An Inertial Measurement Unit (IMU), Global Positioning System (GPS), and flush air data system was included to allow post-flight reconstruction of the vehicle trajectory including a decoupling of the aerodynamics from the atmospheric density. Loadcells were used to measure HIAD structural response during entry, and cameras (both visual-spectrum and infrared) were mounted on the aft segment looking at the aeroshell to monitor structural deflection and surface temperature distribution. Finally, a single up-look camera was included which has provided a surprising amount of science potential from the spectacular footage. In addition to the primary instrumentation suite, a new Fiber Optic Sensing System (FOSS) was used to provide global temperature distributions as a technology demonstration. The system provided over 200 thermal measurements creating a thermal map for the backside of the nose Flexible Thermal Protection System. Another pair of FOSS cables were run along the rigid center structure and measured the temperature response to the vehicle wake environment. The LOFTID instrumentation suite leveraged Agency-wide expertise, with hardware development occurring at Ames Research Center, Langley Research Center, Marshall Space Flight Center, and Armstrong Flight Research Center. This paper will discuss the instrumentation selected for LOFTID, a summary of sensor in-flight performance, and will provide examples of data products from the post-flight analysis effort.

Gregory T Swanson↗
NASA NTRSDate not supplied

Evaluation of the LOFTID Flight Thermocouple Measurements

The Flexible Thermal Protection System (FTPS) on NASA’s Low-Earth Orbit Flight Test of an Inflatable Decelerator (LOFTID) vehicle was instrumented with thermocouples (TC) to measure the in-depth thermal response during entry into Earth’s atmosphere. Accurate flight temperature measurements are critical for verifying vehicle performance during the flight test and reducing uncertainties in the thermal models. The deployable nature of inflatable decelerator technology presents challenges, namely the TCs need to be compactable, cannot damage the FTPS nor the inflatable structure, and need to be able to withstand high temperatures and large thermal gradients over significant running lengths. As previously reported, an extensive ground-test campaign was conducted prior to flight to inform the selection of the LOFTID FTPS TCs which minimized all known potential error sources. Overall, the resulting flight FTPS TC measurements were exceptionally clean throughout the atmospheric entry heat pulse. However, a few of the FTPS TCs still exhibited electrical shorting or other anomalous behavior. This paper presents the flight measurements made by the LOFTID FTPS TCs, identifies anomalous behavior, and discusses potential explanations for the causes of the anomalous behavior. Lastly, recommendations to further mitigate TC measurement error on future Hypersonic Inflatable Aerodynamic Decelerator (HIAD) missions are provided.

Ruth A Miller↗
NASA NTRSDate not supplied

Overview and Performance of the LOFTID Instrumentation Suite

On November 10, 2022, NASA launched the Low-earth Orbit Flight Test of an Inflatable Decelerator (LOFTID) vehicle as a secondary payload mounted inside the launch vehicle adaptor on an Atlas V out of the Vandenburg Space Force Base (VSFB). The primary payload, Joint Polar Satellite System-2 (JPSS-2), was delivered successfully to a sun synchronous trajectory shortly after launch, at which point the Centaur upper stage performed a burn to de-orbit the system. Once on the desired trajectory to enter the atmosphere the payload adaptor canister was ejected by the payload adapter separation system to expose the packed LOFTID vehicle, then the LOFTID aeroshell, a Hypersonic Inflatable Aerodynamic Decelerator (HIAD), was deployed and inflated as planned. The Centaur pointed the LOFTID vehicle to the desired attitude to enter the atmosphere, spun the assembly to roughly three rpm, and separated the reentry vehicle. The LOFTID vehicle entered the atmosphere over Alaska at >8km/sec and decelerated as designed demonstrating stable flight from hypersonic entry through subsonic parachute deployment. On-board visible light cameras captured the reactions of the heatshield through all phases of flight, and co-located infrared light cameras captured the temperature distribution of the aft side of the heat shield anchored to a distribution of thermocouples on the inflatable structure (IS) in the field of view. Thermocouples were also embedded in the forward side of the aeroshell both in the Flexible Thermal Protection System (FTPS) as well as the IS. Heat Flux Gages and Pressure Transducers measured the heating rate and surface pressure experienced by the rigid nose of the reentry vehicle. Loadcells measured the interface loads between the IS and the rigid centerbody. This paper will discuss many of the key instruments flown on the successful LOFTID technology demonstration mission and will provide some high level results, while pointing to more detailed papers on the post-flight analyses.

Gregory T Swanson↗
NASA NTRSDate not supplied

Thermal Protection Systems (TPS) for High Velocity Earth Entry Missions

The fastest man-made object to re-enter into Earth’s atmosphere, with an entry velocity of 12.8 km/s, was the Stardust capsule, which returned samples from the comet Wild 2. Sample return missions offer high and long-lasting science yield; as such, they are discussed repeatedly in the latest planetary science decadal survey, titled Origins, Worlds, and Life. An advantage of sample return missions over in situ measurements is the ability to use state-of-the-art instruments that are not limited in power, size, or complexity; conversely, in situ measurements rely on instruments optimized to fit within the limitations of the spacecraft. Additionally, sample return provides the opportunity to revisit samples as analysis techniques improve as well as to further investigate unanticipated or ambiguous results. As launch vehicle and propulsion technologies continue to advance, it is becoming feasible to consider sample return missions from further out in the Solar System beyond Mars orbit. However, returning samples from increasingly distant destinations comes with high entry velocities, and thus requires ever more capable thermal protection systems (TPS). Evaluating the capabilities of existing mature TPS materials for a range of entry conditions and sample return aeroshell configurations is key to establishing the feasibility of future mission proposals. As such, this presentation describes the comprehensive trade study done to evaluate the feasibility of high velocity Earth entry missions with the currently available TPS materials.

Hannah S Alpert↗
NASA NTRSDate not supplied

A Green’s Function Sensor Fusion Approach for Evaluating Spacecraft Entry Heating From on-Board Thermal Instrumentation

During atmospheric entry, distributed thermal measurements are critical to enable the evaluation of heat loads on spacecraft thermal protection systems (TPS). In recent space exploration missions, Schmidt-Boelter-type heat flux gauges have been integrated into the TPS alongside conventional temperature measurement instrumentation to measure total (convective and radiative) and radiative heat transfer rates1,2. While direct heat flux sensors (HFS) are able to provide valuable information detailing the thermal loads experienced by spacecraft, the interpretation of these measurements in unsteady, convective environments requires a correction factor to account for local heating augmentations at the cold wall HFS surface2,3. Current efforts to estimate cold wall correction factors, and thus recover the hot wall TPS heat flux, rely on time-marching computational fluid dynamics (CFD) simulations2. Simulation-based methods are susceptible to large uncertainties, however, as they require estimations of vehicle trajectory, gas kinetics, wall catalysis models, and other flight conditions as input parameters2,4. Furthermore, CFD simulations are computationally expensive and cannot efficiently survey all possible entry scenarios, exacerbating the uncertainty of reconstructed hot wall heat flux values. These drawbacks motivate the development of alternative hot wall heat flux reconstruction methods that are not reliant on CFD-based correction factors.

Kenneth McAfee↗