Earth Entry Vehicle Research Articles

Infrared Optical Observability of an Earth Entry
 Orbital Test Vehicle Using Ground‐Based
 Remote Sensors

Optical design parameters for a ground-based infrared sensor rely strongly on the target’s optical radiation properties. Infrared (IR) optical observability and imaging simulations of an Earth entry vehicle were evaluated using a comprehensive numerical model. Based on a ground-based IR detection system, this model considered many physical mechanisms including thermochemical nonequilibrium reacting flow, radiative properties, optical propagation, detection range, atmospheric transmittance, and imaging processes. An orbital test vehicle (OTV) was selected as the research object for analysis of its observability using a ground-based infrared system. IR radiance contours, maximum detecting range (MDR), and thermal infrared (TIR) pixel arrangement were modeled. The results show that the distribution of IR radiance is strongly dependent on the angle of observation and the spectral band. Several special phenomena, including a strong receiving region (SRR), a characteristic attitude, a blind zone, and an equivalent zone, are all found in the varying altitude MDR distributions of mid-wavelength infrared (MWIR) and long-wavelength infrared (LWIR) irradiances. In addition, the possible increase in detectivity can greatly improve the MDR at high altitudes, especially for the backward and forward views. The difference in the peak radiance of the LWIR images is within one order of magnitude, but the difference in that of the MWIR images varies greatly. Analyses and results indicate that this model can provide guidance in the design of remote ground-based detection systems.

Feasible zone for planetary entry vehicles

Understanding the coverage of ground track is conducive to evaluating an entry vehicle's maneuver capability, especially for an entry mission with waypoint constraint. This paper investigates a feasible zone which describes the lateral corridor of the entry vehicle's ground track. Any point inside the feasible zone is supposed to be reachable from the initial condition and controllable to the terminal condition. Analyses show that the zone boundary consists of two or three phases. For each phase, the zone boundary is generated by solving one or more trajectory optimization problems. The feasible zone can be employed to judge whether or not a waypoint is feasible to the vehicle. Moreover, to evaluate the feasible level of the waypoint, an evolved feasible zone with pass index is defined and calculated. The feasible zone is simulated for an Earth entry vehicle in various missions. Dispersions are considered in the simulation to analyze the influence of aerodynamic uncertainties.

IMARS Phase 2A Draft Mission Architecture and Science Management Plan for the Return of Samples from Mars Phase 2 Report of the International Mars Architecture for the Return of Samples (iMARS) Working Group.

iMARS Phase 2A Draft Mission Architecture and Science Management Plan for the Return of Samples from Mars Phase 2 Report of the International Mars Architecture for the Return of Samples (iMARS) Working Group.

Impact of state-specific flowfield modeling on atomic nitrogen radiation

Removing typically-applied approximations (quasi-steady-state electronic levels and local escape factors) noticeably impacts the predicted shock-layer radiative heating to the afterbody region of Earth entry vehicles.

Contemporary Impact Analysis Methodology for Planetary Sample Return Missions

Development of an Earth entry vehicle and the methodology created to evaluate the vehicle’s impact landing response when returning to Earth is reported. NASA’s future Mars Sample Return Mission requires a robust vehicle to return Martian samples back to Earth for analysis. The Earth entry vehicle is a proposed solution to this Mars mission requirement. During Earth reentry, the vehicle slows within the atmosphere and then impacts the ground at its terminal velocity. To protect the Martian samples, a spherical energy absorber called an impact sphere is under development. The impact sphere is composed of hybrid composite and crushable foam elements that endure large plastic deformations during impact and cause a highly nonlinear vehicle response. The developed analysis methodology captures a range of complex structural interactions and much of the failure physics that occurs during impact. Numerical models were created and benchmarked against experimental tests conducted at NASA Langley Research Center. The postimpact structural damage assessment showed close correlation between simulation predictions and experimental results. Acceleration, velocity, displacement, damage modes, and failure mechanisms were all effectively captured. These investigations demonstrate that the Earth entry vehicle has great potential in facilitating future sample return missions.

Products from NASA's in-space propulsion technology program applicable to low-cost planetary missions

Since September 2001, NASA's In-Space Propulsion Technology (ISPT) program has been developing technologies for lowering the cost of planetary science missions. Recently completed is the high-temperature Advanced Material Bipropellant Rocket (AMBR) engine providing higher performance for lower cost. Two other cost saving technologies nearing completion are the NEXT ion thruster and the Aerocapture technology project. Under development are several technologies for low-cost sample return missions. These include a low-cost Hall-effect thruster (HIVHAC) which will be completed in 2011, light-weight propellant tanks, and a Multi-Mission Earth Entry Vehicle (MMEEV). This paper will discuss the status of the technology development, the cost savings or performance benefits, and applicability of these in-space propulsion technologies to NASA's future Discovery, and New Frontiers missions, as well as their relevance for sample return missions.

Multiobjective Optimization of Earth-Entry Vehicle Heat Shields

A differential evolutionary algorithm has been executed to optimize the hypersonic aerodynamic and stagnationpoint heat transfer performance of Earth-entry heat shields for manned and unmanned missions. Objective functions comprise maximizing cross range, minimizing heat flux, and minimizing heat load. Each considered heatshield geometry is composed of an axial profile tailored to fit a base cross section. Axial profiles consist of spherical segments, spherically blunted cones, and power law geometries. Heat-shield cross sections include oblate and prolate ellipses, rounded-edge parallelograms, and blendings of the two. Aerothermodynamicmodels are based onmodified Newtonian impact theory with semiempirical correlations for convection and radiation. Entry velocities of 11 and 15 km=s are used to simulate atmospheric entry conditions at lunar and Mars return conditions, respectively. Results indicate that skip trajectories allow for vehicles with a low lift-to-drag ratio of 0.25 to achieve 1000 km cross range, a factor of 4 increase in capability over direct entry. For 11 km=s entry with a 6g deceleration limit, the spherical segment provides optimal performance. For 15 km=s entry with a 12g deceleration limit, the spherically blunted cone produces an 8% lower heat flux when compared with spherical segments with similar aerodynamic characteristics. This result is attributed to the blunted cone’s smaller shock standoff, which reduces the heat load generated by thermal radiation, the dominant heat transfer mode.

Planetary protection for Mars sample return

A Mars sample return (MSR) mission must deal with two new categories of planetary protection requirements: (1) assuring a very low probability of inadvertent release of the sample in order to provide extra protection against the extremely unlikely possibility of biological hazards in the returned material and (2) keeping the samples free of round-trip Earth organisms to facilitate confirmation of safety after return to Earth. This paper describes the planetary-protection-related technical challenges awaiting an MSR mission as well as the range of capabilities being developed to meet these challenges. Containment assurance requires breaking the chain of contact with Mars: the exterior of the sample container must not be contaminated with Mars material. The sample container and its seals must be designed to survive Earth impacts corresponding to the candidate mission profiles, and the Earth entry vehicle must be designed to withstand the thermal and structural rigors of Earth atmosphere entry (all with an unprecedented degree of confidence). Sample contamination must be avoided by sterilizing the entire spacecraft, a challenge with modern avionics, or by sterilizing the sample collection and containment gear and then isolating it from other parts of the spacecraft.

Probabilistic Risk Assessment Strategy for Damage-Tolerant Composite Spacecraft Component Structural Design

Incorporating risk-based design as an integral part of spacecraft development is becoming more and more common. Assessment of uncertainties associated with design parameters and environmental aspects such as loading provides increased understanding of the design and its performance. Results of such studies can contribute to mitigating risk through a system-level assessment. Understanding the risk of an event occurring, the probability of its occurrence, and the consequence(s) of its occurrence can lead to robust, reliable designs. An approach to risk-based structural design incorporating damage-tolerance analysis is described. The application of this approach to a candidate Earth-entry vehicle is also described. The emphasis is on describing an approach for establishing damage-tolerant structural response inputs to a system-level probabilistic risk assessment.

The mars sample return project

The Mars Sample Return (MSR) Project is underway. A 2003 mission to be launched on a Delta III Class vehicle and a 2005 mission launched on an Ariane 5 will culminate in carefully selected Mars samples arriving on Earth in 2008. NASA is the lead agency and will provide the Mars landed elements, namely, landers, rovers, and Mars ascent vehicles (MAVs). The French Space Agency CNES is the largest international partner and will provide for the joint NASA/CNES 2005 Mission the Ariane 5 launch and the Earth Return Mars Orbiter that will capture the sample canisters from the Mars parking orbits the MAVs place them in. The sample canisters will be returned to Earth aboard the CNES Orbiter in the Earth Entry Vehicles provided by NASA. Other national space agencies are also expected to participate in substantial roles. Italy is planning to provide a drill that will operate from the Landers to provide subsurface samples. Other experiments in addition to the MSR payload will also be carried on the Landers. This Papepwill present the current status of the design of the MSR missions and flight articles.

Aerodynamics of Stardust Sample Return Capsule

Successful return of interstellar dust and cometary material by the Stardust Sample Return Capsule requires an accurate description of the Earth entry vehicle''s aerodynamics. This desciption must span the hypersonic-rarefied, hypersonic-continuum, supersonic, transonic, and subsonic flow regimes. Data from numerous sources are compiled to accomplish this objective. These include Direct Simulation Monte Carlo analyses, thermochemical nonequilibrium computational fluid dynamics, transonic computational fluid dynamics, existing wind tunnel data, and new wind tunnel data. Four observations are highlighted: 1) a static instability is revealed in the free-molecular and early transitional-flow regime due to aft location of the vehicle''s center-of-gravity, 2) the aerodynamics across the hypersonic regime are compared with the Newtonian flow approximation and a correlation between the accuracy of the Newtonian flow assumption and the sonic line position is noted, 3) the primary effect of shape change due to ablation is shown to be a reduction in drag, and 4) a subsonic dynamic instability is revealed which will necessitate either a change in the vehicle''s center-of-gravity location or the use of a stabilizing drogue parachute.

Technology and Operational Considerations for Low-Heat-Rate Entry Trajectories

A broad parametric study which examines several critical aspects of low-heat-rate entry trajectories is performed. Low planform loadings associated with future winged Earth-entry vehicles coupled with the potential application of metallic thermal protection systems suggest that such trajectories are of particular interest. Some aspects of these trajectories included in this study are: three heating conditions, i.e., reference, stagnation, and windward centerline (for both laminar and turbulent flow); configuration-related factors including planform loading and hypersonic angle of attack; and mission-related factors such as crossrange and orbit inclination. Results indicate benefits in the design of thermal protection systems are to be gained by utilizing moderate angles of attack, as opposed to high lift coefficients coupled with high angles of attack, during entry. An assessment of design and technology implications is made. Nomenclature e.g. = center of gravity in percent of body reference length measured from the disappearing apex of the nose CL . = lift coefficient cp —

Entry vehicles for space programs

Earth and planetary entry vehicle designs for manned and unmanned space missions, noting thermal protection, environmental control requirements and communication systems

Development and Characterization of a Radio Frequency-Transparent Ablator

Abstract Earth and planetary entry vehicles require heat shield materials whose transparency to radio frequency (rf) signals is unaffected by the ablation process. Teflon or fused silica are currently used for this purpose, but these materials are heavy and difficult to attach. For applications such as Mars entry vehicles or lifting bodies where low heating rates prevail, low-density, rf-transparent ablators are desired. To fill this need, Martin Marietta Corporation has developed a rf-transparent composition designated SLA-220. This ablator, a silica-filled elastomeric silicone, has a density of 0.25 g/cm3 and forms a siliceous rather than a carbonaceous char. Pyrolysis of SLA-220 at slow heating yields a pure silica residue. As heating rate increases, the volatile pyrolysis products tend to undergo secondary decomposition and to deposit carbon while passing through the hot char. Therefore, rf signal attenuation increases with heating rate. Electrical properties of SLA-220 (dielectric constant, signal attenuation) have been measured before and after ablation. Ablation testing has been conducted at heating rates ranging from 6 to 60 Btu/ft2-sec. Thermal properties (conductivity, specific heat, expansion) and mechanical properties (strength, elastic modulus, elongation) have been determined. The properties of SLA-220 are compared with those of 0.50 and 0.70 g/cm3 porous Teflon ablators to demonstrate the advantages of SLA-220 over other low-density rf-transparent ablative systems.