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NASA NTRS1990

Examples of the nonlinear dynamics of ballistic capture and escape in the earth-moon system

An example of a trajectory is given which is initially captured in an elliptic resonant orbit about the earth and then ballistically escapes the earth-moon system. This is demonstrated by a numerical example in three-dimensions using a planetary ephemeris. Another example shows a mechanism of how an elliptic orbit about the earth can increase its energy by performing a complex nonlinear transition to an elliptic orbit of a larger semi-major axis. Capture is also considered. An application of ballistic capture at the moon via an unstable periodic orbit using the four-body sun-earth-moon-S/C interaction is described.

Belbruno, Edward A.↗

NASA NTRS2017

Entry, Descent, and Landing technological barriers and crewed MARSvehicle performance analysis

Mars has been explored historically only by robotic crafts, but a crewed mission encompasses several new engineering challenges – high ballistic coefficient entry, hypersonic decelerators, guided entry for reaching intended destinations within acceptable margins for error in the landing ellipse, and payload mass are all critical factors for evaluation. A comprehensive EDL parametric analysis has been conducted in support of a high mass landing architecture by evaluating three types of vehicles −70° Sphere Cone, Ellipsled and SpaceX hybrid architecture called Red Dragon as potential candidate options for crewed entry vehicles. Aerocapture at the Martian orbit of about 400 km and subsequent Entry-from-orbit scenarios were investigated at velocities of 6.75 km/s and 4 km/s respectively. A study on aerocapture corridor over a range of entry velocities (6–9 km/s) suggests that a hypersonic L/D of 0.3 is sufficient for a Martian aerocapture. Parametric studies conducted by varying aeroshell diameters from 10 m to 15 m for several entry masses up to 150 mt are summarized and results reveal that vehicles with entry masses in the range of about 40–80 mt are capable of delivering cargo with a mass on the order of 5–20 mt. For vehicles with an entry mass of 20 mt to 80 mt, probabilistic Monte Carlo analysis of 5000 cases for each vehicle were run to determine the final landing ellipse and to quantify the statistical uncertainties associated with the trajectory and attitude conditions during atmospheric entry. Strategies and current technological challenges for a human rated Entry, Descent, and Landing to the Martian surface are presented in this study.

Prabhakar Subrahmanyam↗

NASA NTRS2000

GRACE Mission Design: Impact of Uncertainties in Disturbance Environment and Satellite Force Models

The Gravity Recovery and Climate Experiment (GRACE) primary mission will be performed by making measurements of the inter-satellite range change between two co-planar, low altitude, near-polar orbiting satellites. Understanding the uncertainties in the disturbance environment, particularly the aerodynamic drag and torques, is critical in several mission areas. These include an accurate estimate of the spacecraft orbital lifetime, evaluation of spacecraft attitude control requirements, and estimation of the orbital maintenance maneuver frequency necessitated by differences in the drag forces acting on both satellites. The FREEMOL simulation software has been developed and utilized to analyze and suggest design modifications to the GRACE spacecraft. Aerodynamic accommodation bounding analyses were performed and worst-case envelopes were obtained for the aerodynamic torques and the differential ballistic coefficients between the leading and trailing GRACE spacecraft. These analyses demonstrate how spacecraft aerodynamic design and analysis can benefit from a better understanding of spacecraft surface accommodation properties, and the implications for mission design constraints such as formation spacing control.

Mazanek, Daniel D.↗

NASA NTRS2000

Aerothermal Analysis and Design of the Gravity Recovery and Climate Experiment (GRACE) Spacecraft

The Gravity Recovery and Climate Experiment (GRACE) primary mission will be performed by making measurements of the inter-satellite range change between two co-planar, low altitude near-polar orbiting satellites. Understanding the uncertainties in the disturbance environment, particularly the aerodynamic drag and torques, is critical in several mission areas. These include an accurate estimate of the spacecraft orbital lifetime, evaluation of spacecraft attitude control requirements, and estimation of the orbital maintenance maneuver frequency necessitated by differences in the drag forces acting on both satellites. The FREEMOL simulation software has been developed and utilized to analyze and suggest design modifications to the GRACE spacecraft. Aerodynamic accommodation bounding analyses were performed and worst-case envelopes were obtained for the aerodynamic torques and the differential ballistic coefficients between the leading and trailing GRACE spacecraft. These analyses demonstrate how spacecraft aerodynamic design and analysis can benefit from a better understanding of spacecraft surface accommodation properties, and the implications for mission design constraints such as formation spacing control.

Mazanek, Daniel D.↗

NASA NTRS2009

Honeycomb vs. Foam: Evaluating a Potential Upgrade to ISS Module Shielding for Micrometeoroids and Orbital Debris

The presence of a honeycomb core in a multi-wall shielding configuration for protection against micrometeoroid and orbital debris (MMOD) particle impacts at hypervelocity is generally considered to be detrimental as the cell walls act to restrict fragment cloud expansion, creating a more concentrated load on the shield rear wall. However, mission requirements often prevent the inclusion of a dedicated MMOD shield, and as such, structural honeycomb sandwich panels are amongst the most prevalent shield types. Open cell metallic foams are a relatively new material with novel mechanical and thermal properties that have shown promising results in preliminary hypervelocity impact shielding evaluations. In this study, an ISS-representative MMOD shielding configuration has been modified to evaluate the potential performance enhancement gained through the substitution of honeycomb for open cell foam. The baseline shielding configuration consists of a double mesh outer layer, two honeycomb sandwich panels, and an aluminum rear wall. In the modified configuration the two honeycomb cores are replaced by open-cell foam. To compensate for the heavier core material, facesheets have been removed from the second sandwich panel in the modified configuration. A total of 19 tests on the double layer honeycomb and double layer foam configurations are reported. For comparable mechanical and thermal performance, the foam modifications were shown to provide a 15% improvement in critical projectile diameter at low velocities (i.e. 3 km/s) and a 3% increase at high velocities (i.e. 7 km/s) for normal impact. With increasing obliquity, the performance enhancement was predicted to increase, up to a 29% improvement at 60 (low velocity). Ballistic limit equations have been developed for the new configuration, and consider the mass of each individual shield component in order to maintain validity in the event of minor configuration modifications. Previously identified weaknesses of open cell foams for hypervelocity impact shielding such as large projectile diameters, low velocities, and high degrees of impact obliquity have all been investigated, and found to be negligible for the double-layer configuration.

Ryan, Shannon↗

NASA NTRS2011

Orion Parachute Riser Cutter Development

This paper presents the tests and analytical approach used on the development of a steel riser cutter for the CEV Parachute Assembly System (CPAS) used on the Orion crew module. Figure 1 shows the riser cutter and the steel riser bundle which consists of six individual cables. Due to the highly compressed schedule, initial unavailability of the riser material and the Orion Forward Bay mechanical constraints, JSC primarily relied on a combination of internal ballistics analysis and LS-DYNA simulation for this project. Various one dimensional internal ballistics codes that use standard equation of state and conservation of energy have commonly used in the development of CAD devices for initial first order estimates and as an enhancement to the test program. While these codes are very accurate for propellant performance prediction, they usually lack a fully defined kinematic model for dynamic predictions. A simple piston device can easily and accurately be modeled using an equation of motion. However, the accuracy of analytical models is greatly reduced on more complicated devices with complex external loads, nonlinear trajectories or unique unlocking features. A 3D finite element model of CAD device with all critical features included can vastly improve the analytical ballistic predictions when it is used as a supplement to the ballistic code. During this project, LS-DYNA structural 3D model was used to predict the riser resisting load that was needed for the ballistic code. A Lagrangian model with eroding elements shown in Figure 2 was used for the blade, steel riser and the anvil. The riser material failure strain was fine tuned by matching the dent depth on the anvil with the actual test data. LS-DYNA model was also utilized to optimize the blade tip design for the most efficient cut. In parallel, the propellant type and the amount were determined by using CADPROG internal ballistics code. Initial test results showed a good match with LS-DYNA and CADPROG simulations. Final paper will present a detailed roadmap from initial ballistic modeling and LS-DYNA simulation to the performance testing. Blade shape optimization study will also be presented.

Oguz, Sirri↗

NASA NTRS1992

Space aging of solid rocket materials

Solid rocket propellant and rocket motor components were aged in a vented container on the interior of the LDEF. This paper will present the results of aging the Improved Performance Space Motor-II/Payload Assist Module-Delta II (IPSM-II/PAM-DII) space motor components. Ballistic and mechanical properties of the space aged main propellant, igniter propellant, and ignition system were compared with similar data from preflight and ground aged samples. Mechanical properties of the composite materials and bonded joints used in the motor case, insulation, liner, nozzle, exit cone, and skirt were similarly evaluated. The space aging results will be compared to data collected in a ground based vacuum aging program on similar components. The operation of the vacuum actuated venting valve and pressure actuated resealing of the container will also be addressed. The materials tested showed no significant changes due to space aging. These results indicate that properly designed solid rocket motors can be expected to perform reliably after extended periods of exposure to a space environment.

Dean M Lester↗

NASA NTRSDate not supplied

Aerodynamic Performance of the 2018 InSight Mars Lander

InSight touched down in Elysium Planitia on 26 November 2018, becoming NASA’s eighth successful entry, descent, and landing (EDL) at Mars. InSight was a build-to-print of the successful 2008 Phoenix EDL system, flying a non-spinning, ballistic trajectory with a 70-degree sphere-cone aeroshell (2.65-meter diameter), disk-gap-band parachute, and pulsed terminal descent and landing engines. This work discusses entry aerodynamic performance for InSight up to parachute deployment, including pre-flight aerodynamics predictions and comparisons with post-flight reconstruction, as well as comparisons with the Phoenix reconstruction.

Ashley M Korzun↗

NASA NTRS2017

Targeting Cislunar Near Rectilinear Halo Orbits for Human Space Exploration

Part of the challenge of charting a human exploration space architecture is finding locations to stage missions to multiple destinations. To that end, a specific subset of Earth-Moon halo orbits, known as Near Rectilinear Halo Orbits (NRHOs) are evaluated. In this paper, a systematic process for generating full ephemeris based ballistic NRHOs is outlined, different size NRHOs are examined for their favorability to avoid eclipses, the performance requirements for missions to and from NRHOs are calculated, and disposal options are evaluated. Combined, these studies confirm the feasibility of cislunar NRHOs to enable human exploration in the cislunar proving ground.

Williams, Jacob↗

NASA NTRS2018

NASA ACC High Energy Dynamic Impact Methodology and Outcomes

High fidelity analysis methods known as Progressive Damage Analysis (PDA) methods, capable of reliably predicting the onset and progression of damage in composite materials, are being developed for HEDI event simulation. Four state-of-the-art PDA methods are being investigated: LS-DYNA MAT162, LS-DYNA MAT261, Smooth-Particle Galerkin (SPG), and EMU Peridynamics. Validation of these PDA models aims to follow a building block approach, starting with coupon- and component-level development. As part of the overall effort, material characterization testing was performed to bridge gaps between existing experimental data and the material property inputs required to predict ballistic impact behavior at the component-level. The material models were updated with the results of the coupon-level testing and were then used to predict panel behavior and damage in ballistic impact events. To assess the accuracy of the PDA methods, these pre-test predictions were compared to the results of ballistic impact testing.

Hunziker, Kenneth J.↗

NASA NTRS2018

Studies in Support of Venus Aerocapture Utilizing Drag Modulation

Aerocapture has been extensively studied and these studies have shown the benefit for planetary exploration missions. While the traditional approach to aerocapture with lifting configurations and lift-guided modulations have been assessed to be technologically feasible, aerocapture using purely drag modulation was proposed and studied by Prof. Braun and his students. These studies show that if one can assess the feasibility of aerocapture using drag modulation at Venus, and develop tall pole technologies needed at Venus, then this concept is much easier to execute at all other relevant destinations. Based on the above finding, partnered proposals were submitted by Adam Nelessen at JPL and Ethiraj Venkatapathy at Ames in collaboration with Prof. Braun at the University of Colorado, Boulder (UCB). Under this partnership, Ames Research Center (ARC) is working to address some of the key entry technology challenges associated with drag modulation aerocapture at Venus. Drag modulation aerocapture is a simple, scalable, and likely cost-effective way to enhance planetary science missions. The approach envisioned is to design a small spacecraft, that would most likely be a secondary payload, with a removable drag skirt. The vehicle would enter the atmosphere at Venus with a low ballistic coefficient, decelerate rapidly, drop the skirt resulting in a smaller vehicle with a higher ballistic coefficient which would skip out of the atmosphere and enter into a desired orbit. ARC's role in this collaboration is multifold. First of which is to perform design studies on various pre- and post-jettison geometries utilizing a 3-DOF trajectory code to determine the aerodynamics and aerothermodynamics of the vehicles and evaluate viable thermal protection material system designs. Once these design studies are complete, Ames will then perform higher fidelity CFD and TPS sizing to further design the vehicles. Second, the multi-body separation dynamics of the drag modulation event will be explored using both CFD simulations (CART3-D and US3D) as well as possible ballistic range testing. ARC's tools and expertise have been used to assess and advise on the selection of the separating configuration. In addition to the preliminary evaluation, ARC will provide tools and expertise to UCB team members to further assess aerodynamic interactions between the separating bodies and provide guidance as to the feasibility of stable transition.

Beck, Robin↗

NASA NTRS2012

Aero-Assisted Pre-Stage for Ballistic and Aero-Assisted Launch Vehicles

A concept of an aero-assisted pre-stage is proposed, which enables launch of both ballistic and aero-assisted launch vehicles from conventional runways. The pre-stage can be implemented as a delta-wing with a suitable undercarriage, which is mated with the launch vehicle, so that their flight directions are coaligned. The ample wing area of the pre-stage combined with the thrust of the launch vehicle ensure prompt roll-out and take-off of the stack at airspeeds typical for a conventional jet airliner. The launch vehicle is separated from the pre-stage as soon as safe altitude is achieved, and the desired ascent trajectory is reached. Nominally, the pre-stage is non-powered. As an option, to save the propellant of the launch vehicle, the pre-stage may have its own short-burn propulsion system, whereas the propulsion system of the launch vehicle is activated at the separation point. A general non-dimensional analysis of performance of the pre-stage from roll-out to separation is carried out and applications to existing ballistic launch vehicle and hypothetical aero-assisted vehicles (spaceplanes) are considered.

access to space↗

NASA NTRS1977

Aerodynamic behavior of the Viking entry vehicle - Ground test and flight results

An extensive series of tests of the Viking entry vehicle flying in pure CO2 was conducted in a ballistic range at Ames Research Center. The primary purpose of these tests was to calibrate the aerodynamic lift and drag characteristics in order to allow the density, pressure, and temperature profiles of the Martian atmosphere to be determined from onboard instrumentation carried on Viking. Both the Viking 1 and Viking 2 entry vehicles performed flawlessly during entry and descent, and the atmosphere structure was deduced to an altitude of about 120 km. A description is given of the ballistic range tests and of the aerodynamic behavior of the full scale entry vehicles during entry into the Martian atmosphere. Some comparisons between ground test and flight results are shown.

Kirk, D. B.↗

NASA NTRS1972

Potential advantages of solar electric propulsion for outer planet orbiters.

Past studies of solar electric propulsion for outer planet orbiters have generally emphasized the advantages of flight time reduction and payload increases. However, several subtle advantages exist, which may become important in an environment of increasingly difficult requirements as ways to extend current technology are sought. These advantages accrue primarily because of the inherent capability, unique to electric propulsion, to efficiently shape a trajectory while enroute. Stressed in this paper are: the ability to meet orbital constraints due to assumed radiation belts, science flexibility in a dual launch program, increased numbers of observational passes, and the lengthening of launch periods. These are examined for years representative of relatively easy and difficult ballistic missions. The results indicate that an early investment in solar electric technology will provide a strong performance foundation for a long range outer planet exploration program which evolves from current spacecraft technology.

Sauer, C. G.↗

NASA NTRS2017

Spacecraft Re-Entry Impact Point Targeting Using Aerodynamic Drag

The ability to re-enter the atmosphere at a desired location is important for spacecraft containing components that may survive re-entry. While impact point targeting has traditionally been initiated through impulsive burns with chemical thrusters on large vehicles such as the Space Shuttle, and the Soyuz and Apollo capsules, many small spacecraft do not host thrusters and require an alternative means of impact point targeting to ensure that falling debris do not cause harm to persons or property. This paper discusses the use of solely aerodynamic drag force to perform this targeting. It is shown that by deploying and retracting a drag device to vary the ballistic coefficient of the spacecraft, any desired longitude and latitude on the ground can be targeted provided that the maneuvering begins early enough and the latitude is less than the inclination of the orbit. An analytical solution based on perturbations from a numerically propagated trajectory is developed to map the initial state and ballistic coefficient profile of a spacecraft to its impact point. This allows the ballistic coefficient profile necessary to reach a given target point to be rapidly calculated, making it feasible to generate the guidance for the decay trajectory onboard the spacecraft. The ability to target an impact point using aerodynamic drag will enhance the capabilities of small spacecraft and will enable larger space vehicles containing thrusters to save fuel by more effectively leveraging the available aerodynamic drag.

propagation↗

NASA NTRS2007

Space Shuttle MMOD Threat Mitigation Techniques

Prior to each shuttle mission, threat assessments are performed to determine the risk of critical penetration, payload bay door radiator tube leak and crew module window replacement from Micrometeoroid and Orbital Debris (MMOD). Mission parameters, such as vehicle attitude, exposure time and altitude are used as inputs for the analysis. Ballistic limit equations, based on hypervelocity impact testing of shuttle materials are used to estimate the critical particle diameters of the outer surfaces of the vehicle. The assessments are performed using the BUMPER computer code at the NASA/JSC Hypervelocity Impact Technology Facility (HITF). The most critical involves the calculation of Loss of Crew and Vehicle (LOCV) risk. In recent years, NASA has implemented several techniques to reduce the risk to the Shuttle from MMOD impacts. This paper will describe on-orbit inspection of the reinforced carbon-carbon (RCC) regions and the methods used discern hypervelocity impact damage. Impact damage contingency plans and on-orbit repair techniques will also be discussed. The wing leading edge impact detection system (WLEIDS) and it's role in the reduction of on-orbit risk reduction will be presented. Finally, an analysis of the effectivity of alternative shuttle flight attitudes on MMOD risk will be demonstrated.

Hyde, Justin L.↗

NASA NTRS2008

Space Shuttle MMOD Threat Mitigation Techniques

Prior to each shuttle mission, threat assessments are performed to determine the risk of critical penetration, payload bay door radiator tube leak and crew module window replacement from Micrometeoroid and Orbital Debris (MMOD). Mission parameters, such as vehicle attitude, exposure time and altitude are used as inputs for the analysis. Ballistic limit equations, based on hypervelocity impact testing of shuttle materials are used to estimate the critical particle diameters of the outer surfaces of the vehicle. The assessments are performed using the BUMPER computer code at the NASA/JSC Hypervelocity Impact Technology Facility (HITF). The most critical involves the calculation of Loss of Crew and Vehicle (LOCV) risk. An overview of significant MMOD impacts on the Payload Bay Door radiators, wing leading edge reinforced carbon-carbon (RCC) panels and crew module windows will be presented, along with a discussion of the techniques NASA has implemented to reduce the risk from MMOD impacts. This paper will describe on-orbit inspection of the RCC regions and the methods used discern hypervelocity impact damage. Impact damage contingency plans and on-orbit repair techniques will also be discussed. The wing leading edge impact detection system (WLEIDS) and it s role in the reduction of on-orbit risk reduction will be presented. Finally, an analysis of alternative shuttle flight attitudes on MMOD risk will be demonstrated.

Hyde, J. L.↗

NASA NTRS2013

Atmospheric Entry Studies for Venus Missions: 45 Sphere-Cone Rigid Aeroshells and Ballistic Entries

The present study considers direct ballistic entries into the atmosphere of Venus using a 45deg sphere-cone rigid aeroshell, a legacy shape that has been used successfully in the past in the Pioneer Venus Multiprobe Mission. For a number of entry mass and heatshield diameter combinations (i.e., various ballistic coefficients) and entry velocities, the trajectory space in terms of entry flight path angles between skip out and -30deg is explored with a 3DoF trajectory code, TRAJ. From these trajectories, the viable entry flight path angle space is determined through the use of mechanical and thermal performance limits on the thermal protection material and science payload; the thermal protection material of choice is entry-grade carbon phenolic, for which a material thermal response model is available. For mechanical performance, a 200 g limit is placed on the peak deceleration load experienced by the science instruments, and 10 bar is assumed as the pressure limit for entry-grade carbon-phenolic material. For thermal performance, inflection points in the total heat load distribution are used as cut off criteria. Analysis of the results shows the existence of a range of critical ballistic coefficients beyond which the steepest possible entries are determined by the pressure limit of the material rather than the deceleration load limit.

Prabhu, Dinesh K.↗