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To assure successful tracking of the four Pioneer Venus probes, it is necessary for the deep space station receiver operators to be familiar with the characteristics of the received signals. As an aid for training the operators for this mission, signal simulators have been developed which completely duplicate the probe signals. This article describes the method used for generating these test signals.
A 6.4-cm-diameter scale model of the Jovian entry vehicle is tested in an electric-arc-driven shock tube and a 5-cm-diameter sphere model is tested in a combustion-driven shock tube and in an electric-arc-driven shock tunnel. The radiative heat-transfer rate and pressure on the front and the base regions are measured in the absence of ablation with sensors imbedded in the models in a stream consisting of 10% hydrogen in a bath of either neon or argon. The measured radiative heat-transfer rates and pressures range to about 22 kW/sq cm and 12 atm, respectively, at the front stagnation point. The ratio of the radiative heat-transfer rate at the base stagnation point to that at the front stagnation point is found to be about 1/4 for the sphere at Mach 1.8, about 1/30 for the sphere at Mach 4.8, and about 1/6 for the scale model at Mach 1.7. The present experimental results agree well with the theoretical predictions of Park, thus indicating that Park's theory is valid.
Saturn Probe and Ice Giant Orbiter along with in-situ Probe Missions continue to be very high priority mis-sions. A result of the advocacy by OPAG and other Analysis Groups, the 3D Woven, Dual-Layer HEEET thermal protection system, mature at TRL 6, has closed the TPS gap for extreme environment missions. A mid density follow on to DL HEEET, developed to meet earth entry requirements, is a Single Layer vari-ant of the 3D woven TPS, which provides a mass effi-cient single layer 3D Mid-Density TPS (3MDCP) that has been baselined as the heatshield for MSR EEV. Continued development of 3MDCP will elevate it to TRL 6 by 2025. Conformal-PICA (C-PICA) devel-opment was pursued to establish a more efficient and robust alternate to PICA, and it is at TRL 4+. Our rationale for the next generation of TPS de-velopment is based on the on the missions needs of the next decade Outer Planet missions that are unique and more demanding than any other destinations. Taking advantage of the recent planned development of sev-eral materials, the next generation of TPS offers a much more mass efficient option for small, medium, and large class Outer Planet missions. Saturn Probes: While mission designers are gen-erally interested in shallow entry to maintain the g-load during entry around 50g, the heat-load for shallow entry can range between (100 kJ/cm2 – 300 kJ/cm2), two orders of magnitude higher than Venus or Sample Return missions. TPS must not only offer protection but must be mass efficient to perform reasonable sci-ence. TPS mass can quickly become 50% or more of the mass of the entire entry system. Recent analysis performed shows Saturn Probe missions could signifi-cantly benefit from the single layer HEEET (SL-HEEET/3MDCP). A DL-HEEET based heatshield mass could be ~ (40% -50%), SL-HEEET can provide additional (30% - 50%) mass savings. At the same time, C-PICA can provide (30% - 50%) mass savings on the backshell. The combined mass savings can be significant enough to carry an additional probe, if de-sired. Ice Giant Aerocapture Missions: Aerocapture mission architectures can provide significant ad-vantage over traditional propulsive insertion missions in multiple ways. 1) Reduced trip time ~ (4- 6) years (30% -40%), 2) Enables placing the orbiter, probe, and lander, all together and 3) Allows for greater science mass (probes and landers) due to mass efficiency. The delivery of a probe from orbit makes it easier and eliminates mission design constraints by HEEET for direct entry [4] of probes and allows for more targeted in-situ science once the Orbiter is able to collect data. In the past 20 years, progress made in GN&C for lift-guided entry missions (MSL, Orion EFT1, Mars 2020) and the expertise in blunt body aerodynamics at large scale (~ 5m) has led the EDL community to conclude that aerocapture is a “go do” engineering activity. An aerocapture mission that will deplete the excess energy of a fast arrival mission will require a mass efficient TPS that can handle extreme heat-load, ~ (100 kJ/cm2 – 500 kJ/cm2). Hence TPS, feasibility as well as mass efficiency requires assessment. Utilizing the recent developments, a comprehensive, bounding analysis was done to establish the potential for SOA (HEEET) system as well as emerging new TPS such as SL-HEEET and C-PICA. In this proposed poster, we will outline the process by which we establish bounding aerocapture trajecto-ries for hyperbolic excess velocities ranging from 27 km/s to 35 km/s, for low L/D (~ 0.4) configurations and determine conservative/bounding estimate of aer-othermal environment by using a combination of CFD simulations and stagnation point heating estimates [7]. This engineering approach allows us to assess the TPS need vs. TPS capability and determine the applicability of existing TPS. Once an applicable suite of TPS is determined, the TPS thickness and mass are comput-ed. We show that the TPS mass fraction can be as low as 5% to as high as 20%, depending on the use of ad-vanced TPS, while HEEET is sufficient but will require 50% of the entry mass.
Flight proven entry system and TPS technologies are critical for the successful execution of in-situ science missions at Venus. Emerging new technologies point to new possibilities and offer innovative approaches to delivering small satellites for orbital science. Venus entry can be very demanding and there are only a few flight proven TPS, some developed by Industry and others by NASA, capable of meeting the mission needs. NASA developed TPS has predominately been transferred to Industry and it is assumed industry will maintain the fabrication capability. However, lack of mission needs may result in obsolence of TSP fabrication capability if there is no money and no motivation. Even within NASA, its' expertise could be diverted to higher priority objectives and thereby the readiness for particular material systems can be impacted or lost. Atrophy of capabilities can come about in other ways as well such as changes to raw materials. Even small manufacturing process changes can demand requalification and TRL may be degraded. Carbon-Phenolic is a text book example. After a long period of absence of US Venus missions, VEXG and the Science community is making the case for future missions. It is insufficient to assume the TSP technologies will be there in 5 or 10 years without active and continual planning and assessment. After Galileo, Carbon-Phenolic materials and fabrication skills were allowed to atrophy. Then when missions needed it, in early 2000, it was no longer possible to make the heritage Carbon-Phenolic. What do we need to do? The first step is to advocate for the establishment of TPS readiness assess-ment. The assessment will involve understanding threats and opportunities, and the development of risk mitigation strategies. VEXAG needs to advocate for such an active monitoring of the needed capabilities, assessment of emerging risks and development of risk mitigation strategies with implementation plans. Such an approach reduces the threat of material obsolence and helps maintain the availability of entry system and TPS technology capabilities, both old and new. Venus probes, landers, balloons and other variable altitude missions, and skimmer missions such as "Cu-pid's Arrow" as well as aerocapture missions to deliver small spacecraft require qualified entry systems and ablative TPS. VEXAG advocated for HEEET in 2013/2014 and the community is well versed with the need to sustain it. But, other TPS that need to be sustained may not be apparent to VEXAG community. The following figure summarizes the ablative TPS capabilities vs Venus mission needs for both primary heatshield and backshell.
While the HEEET project established the dual-layer, 3-D Woven TPS at TRL 6 for extreme entry environment missions such as Saturn and Ice-Giant Probe missions, subsequent to the completion of HEEET project in 2018, further development was undertaken to assess the viability of a variant of the 3-D Woven, a single, insulating layer alone as a viable TPS for Mars Sample Return mission. MSR Earth Entry System (EES) baselined it and named the TPS as 3D, mid-density, carbon-phenolic (3MDCP). MSR program invested in building a new loom and the weaver T.E.A.M. Inc is currently weaving 80” wide 1.2” thick insulating layer. Unlike HEEET, where a set of tiles and seams are integrated to form the heat-shield, 3MDCP is a seamless heatshield by wrapping and shaping a flat panel into the heatshield shape and then rigidizing the preform through resin infusion. MSR EES is investing in technology maturation of 3MDCP for the current mission launch in 2027. This lighting talk will provide key developments and time line for 3MDCP technology maturation to show that 3MDCP will be ready to support future UOP or Saturn Probe mission. In addition, the limitations of the capability will be presented so that Saturn mission proposers as well as UOP flagship mission study can be fully informed as to the 3MDCP capability.
Aerosol Rapid Analysis Combined Entry Probe/Sonde Technology (AERACEPT) is a dual descent probe and in-situ sampler of cloud and aerosol particles, used as part of the Nephele mission concept to study Venus cloud layers. Various optical spectroscopic characterization methods, including Raman spectroscopy and Surface Enhanced Raman Spectroscopy (SERS), Laser induced Breakdown Spectroscopy (LIBS), and absorbance and fluorescence spectroscopy are being assessed for viability for compatibility with AERACEPT. Specifically, the detection limits, sensitivity, specificity, and analysis cadence of these techniques for a set of Venus aerosol analogs and their mixtures are tested. We report on preliminary optical spectroscopy results of the Venus aerosol analogs, as well as optical substrate survivability in heated and acidic conditions, substrate physical robustness, and analyte wettability on substrates.
In 2012, NASA's mission of landing the Mars Science Laboratory (MSL) on Mars was successful. MSL was protected by an ablative heatshield of tiled low density material known as PICA (phenolic impregnated carbon ablator). The heatshield was instrumented with MEDLI (MSL Entry Descent Landing Instrument) a suite of sensors & thermocouples at discrete locations in order to monitor the in-depth ablator temperature response and surface pressure. MEDLI was designed and developed by NASA Langley, in partnership with NASA Ames Research Center for the purposes of probing Mars and evaluating the performance of the spacecraft upon entry into the Martian atmosphere. The flight data reduces the uncertainty in engineer models for predicting the response of a spacecraft towards the extreme heating environment of an entry into the Martian atmosphere. MEDLI2 is a part of the Mars 2020 mission and is the next-generation sensor suite for entry, descent, and landing. This data will again help engineers validate their flight models. Additionally, the atmospheric data, can help us understand atmospheric density and winds. This is a critical study for reducing risks to both robotic and future human missions to Mars. Engineered models for Mars 2020 are dependent upon parameters related to the materials response to heating and radiation. In this work, the thermal properties and other measurements of various ablative materials are analyzed to achieve greater utility of the 2012 MEDLI flight data and more accurately determine the parameters being used in Mars 2020. The purpose of this study was to measure specific heat, thermal conductivity, char yield, reflectance as a function of wavelength, and other thermal parameters and optical properties of various ablative materials. CO2(g) at extreme temperatures emits radiation impacting MEDLI2 flight predictions. Emissivity & absorptivity as a function of temperature were calculated from FTIR and UV-Vis data as a means of investigating whether these materials might absorb CO2 radiation upon entry into the Mars atmosphere.
NASA's science mission directorate has put increasing emphasis on innovative, smaller, and lower cost missions to achieve their science objectives. One example of this was the recent call by the Planetary Science Division for cube and small satellite concepts expected to cost $100M or less, not including launch and weighing less than 180kg. Over 100 proposals were submitted suggesting that indeed this is a size of mission worthy of being considered in future planning. Nineteen missions were selected for study, one being a long-lived Venus mission called SAEVe, for Seismic and Atmospheric Exploration of Venus. The science objectives and relevance of SAEVe include: Is Venus seismically active? What can we learn about its crust (thickness and composition) and its interior (lithosphere, mantle, and core)? What can be learned about its evolutionary history or about the planet / atmosphere interactions? SAEVe begins to address these science questions with simple, but capable, instrumented probes that can survive on the surface of Venus and take temporal measurements over months something never attempted before. The data returned will further our understanding of the solar system and Earth, and aid in meeting the NASA Science Plan goal to ascertain the content, origin, and evolution of the solar system and the chemical and physical processes in our solar system. SAEVe is delivered to Venus as a ride-along on another mission to Venus. Its two small probes are placed into the Venus atmosphere via a single Stardust-like entry capsule, are ejected at different times, free fall, and decelerate in the thickening atmosphere to touchdown under 8 m/s2 or less. The probes will begin taking measurements and transmitting important parameters at or near the surface and will focus on measurements like seismic activity, heat flux, wind speed and direction, basic chemical abundances, temperature, and pressure. At preset intervals, the probes acquire the science measurements and beam the data to the orbiting host spacecraft. SAEVe will serve as a highly capable precursor and pave the way for larger and more complex lander missions to explore Venus.
Solutions are presented for the aerothermal heating environments about blunt probes entering the atmosphere of Jupiter. The analysis employs a novel prediction procedure in which finite differences versions of the thin layer fluid dynamic equations are solved simultaneously with, and coupled to, a fully detailed radiation flux model. Predictions are presented and compared with those of earlier investigators. Radiative blowing corrections are presented for on- and off-stagnation point conditions. The use of a blowing correlation based on the retained mass was found to be useful for off-stagnation point conditions.
Soon after Galileo Probe mission, capabilities to manufacture and test heritage carbon-phenolic (HCP) TPS atrophied. The impact of capability loss to Outer Planet probe missions, Venus in-situ and Mars Sample Return (MSR) missions was realized after two decades. MSR (1997 – 2006) initiated a search and recover of HCP. Decadal Survey (2003) recommended JPOP mission (Jupiter Polar Orbiter and Probe), but lack of HCP TPS resulted in Juno being proposed without a probe. Decadal Survey (2013) recognized and advised NASA to close the TPS gap. With the community support, NASA initiated a technology development effort to close this gap in 2019 and today, missions to Venus, Saturn, Uranus, and MSR are enabled by new 3-D Woven TPS. Jupiter, the most demanding mission from an entry and TPS perspective ever, still lacks a viable qualified TPS. While the Science community needs to make the case for going back to Jupiter with a Probe, this work is an assessment of emerging capabilities to support such a mission when needed. Recent SOA developments are very promising to close this technology gaps. Encourage the Science Community about the feasibility of future Jupiter Multi-Probes Mission Engage the Entry System/TPS Technology Community as to the emerging capabilities and need to preserve them Educate and excite the early career and next generation students interested in future Science and technology challenge.
Increased scientific interest in the structure and composition of Titan's atmosphere, clouds and surface have led to the study of mission options to the Saturnian system with the main goal of placing a probe into the atmosphere of the satellite. Two probe concepts have been studied by NASA: the first concept, a slightly modified Galileo probe capable of withstanding approximately 50 earth G during atmospheric entry heating and deceleration, would consist of a blunted 53 degree, 136-cm-diameter half-angle cone with a hemispherical afterbody, and a descent module containing scientific instruments and a parachute; the second concept, a system designed to provide in situ atmospheric measurements of Titan's organic haze layer, would consist of a probe using a 165-cm deployable graphite fabric decelerator, a 50-cm-diameter cylindrical descent module containing five instruments and a 2.5 m-diameter parachute and a 50-cm-radius spherical nose cap. Although the modified Galileo probe is feasible, its scientific drawback includes its inability to obtain in situ measurements above approximately 100 km.
Neptune is a prime destination for future exploration missions. This study designed the mission concept including the Flagship-class orbiter and atmospheric probe to the Neptune-Triton system. The results of this study will be used as the Decadal Survey considers on NASA’s planetary science priorities from 2022-2032.This poster will focus on the design of the entry and descent sequence for the atmospheric probe.
A coupled, trajectory-based flowfield and material thermal-response analysis is presented for the European Space Agency proposed Rosetta comet nucleus sample return vehicle. The probe returns to earth along a hyperbolic trajectory with an entry velocity of 16.5 km/s and requires an ablative heat shield on the forebody. Combined radiative and convective ablating flowfield analyses were performed for the significant heating portion of the shallow ballistic entry trajectory. Both quasisteady ablation and fully transient analyses were performed for a heat shield composed of carbon-phenolic ablative material. Quasisteady analysis was performed using the two-dimensional axisymmetric codes RASLE and BLIMPK. Transient computational results were obtained from the one-dimensional ablation/conduction code CMA. Results are presented for heating, temperature, and ablation rate distributions over the probe forebody for various trajectory points. Comparison of transient and quasisteady results indicates that, for the heating pulse encountered by this probe, the quasisteady approach is conservative from the standpoint of predicted surface recession.
A coupled, trajectory based flowfield and material thermal response analysis is presented for the European Space Agency (ESA) proposed Rosetta comet nucleus sample return vehicle. The probe returns to Earth along a hyperbolic trajectory with an entry velocity of 16.5 km/sec and requires an ablative heat shield on the forebody. Combined radiative and convective, ablating flowfield analyses were performed for the significant heating portion of the shallow ballistic entry trajectory. Both quasi-steady ablation and fully transient analyses were performed for a heat shield composed of carbon-phenolic ablative material. Quasi-steady analysis was performed using the two-dimensional, axisymmetric codes RASLE and BLIMPK. Transient computational results were obtained from the one-dimensional ablation/conduction code, CMA. Results are presented for heating, temperature and ablation rate distributions over the probe forebody for various trajectory points. Comparison of transient and quasi-steady results indicates that, for the heating pulse encountered by this probe, the quasi-static approach is conservative from the standpoint of predicted surface recession.
A detailed Neptune aerocapture systems analysis and spacecraft design study was performed to improve our understanding of the techonology requirement for such a hard mission. The primary objective was to engineer a point design based on blunt body aeroshell technology and quantitatively assess feasibility and performance. This paper reviews the launch vehicle, propulsion, and trajectory options to reach Neptune in the 2015-2020 time frame using aerocapture and all-propulsive vehicles. It establishes the range of entry conditions that would be consistent with delivering a - 1900 kg total entry vehicle maximum expected mass to Neptune including a - 790 kg orbiter maximum expected mass to the science orbit. Two Neptune probes would be also be delivered prior to the aerocapture maneuver. Results show that inertial entry velocities in the range of 28 to 30 km/s are to be expected for chemical and solar electric propulsion options with several gravity assists (combinations of Venus, Earth and Jupiter gravity assists). Trip times range from approximately 10-11 years for aerocapture orbiters to 15 years for all-propulsive vehicles. This paper shows that the use of aerocapture enables this mission given the payload to deliver around Neptune compared to an all-propulsive orbit insertion approach. However, an all-propulsive chemical insertion option is possible for lower payload masses than the one needed for this science mission. Both approaches require a Delta IV heavy class launch vehicle.
Design criteria of passive damping system for Mars entry vehicle - damping of angle of attack oscillation during entry trajectory
Application of a viscous shock-layer analysis to the calculation of nonequilibrium-flow species distributions in the plasma layer of a blunt-nosed vehicle at high altitudes. The theoretical electron-density results obtained are in good agreement with those measured in flight for a hemisphere-9 deg cone entry vehicle. The flight measurements were obtained using electrostatic probes that protruded well into the shock layer. In addition, the theoretically obtained heavy-particle translational temperatures appear to agree fairly well with the electron temperatures that were measured in the flight experiments using voltage-swept thin-wire electrostatic probes. The influence of the reaction rate coefficients on the calculated electron densities has been assessed and shown to be within the uncertainty in the flight data. The theoretical results demonstrate the importance of including in the chemical model the positive ions N2(+), O2(+), N(+), and O(+), in addition to NO(+), for the high altitudes and velocities considered.
The Pioneer Venus Orbiter and Multiprobe spacecraft design and operation enabled both remote and in-situ measurements of the Venusian environment from the outermost fringes of the atmosphere all the way to the surface. Both spacecraft were spin-stabilized and solar-cell powered from launch to Venus. Since orbit insertion, the Orbiter has been transmitting measurements from a highly elliptical 24-h orbit with periapsis altitudes down to about 150 km. Data rates up to 2048 bits/s have been utilized through a despun high-gain antenna transmitting at S-band frequency. Spacecraft attitudes, orbit periods, and periapsis altitudes are being maintained as required with a hydrazine propulsion system. The Multiprobe spacecraft (Bus with all four Probes attached) performed the necessary Probe checkouts and deployed the Probes to achieve the desired Probe and Bus targeting. Silver-zinc batteries provided the necessary power on each of the four Probes from separation from the Bus through the entry/descent sequence. Data rates of 256 and 128 bits/s on the Large Probe were maintained with 40-W radiated power, and 64 and 16 bits/s on the Small Probes were maintained with 10-W radiated power, through omni antennas directly to Earth-based stations. Each Probe's entry/descent sequence was controlled with a hardwired entry sequence programmer to achieve the desired scientific and spacecraft operations.