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All sourcesNASA NTRSDOE OSTI
NASA NTRS2005

An Experimental Study of n-Heptane and JP-7 Extinction Limits in an Opposed Jet Burner

Propulsion engine combustor design and analysis requires experimentally verified data on the chemical kinetics of fuel. Among the important data is the combustion extinction limit as measured by observed maximum flame strain rate. The extinction limit relates to the ability to maintain a flame in a combustor during operation. Extinction limit data can be obtained for a given fuel by means of a laminar flame experiment using an opposed jet burner (OJB). Laminar extinction limit data can be applied to the turbulent application of a combustor via laminar flamelet modeling. The OJB consists of two axi-symmetric tubes (one for fuel and one for oxidizer), which produce a flat, disk-like counter-flow diffusion flame. This paper presents results of experiments to measure extinction limits for n-heptane and the military specification fuel JP-7, obtained from an OJB. JP-7 is an Air Force-developed fuel that continues to be important in the area of hypersonics. Because of its distinct properties it is currently the hydrocarbon fuel of choice for use in Scramjet engines. This study provides much-desired data for JP-7, for which very little information previously existed. The interest in n-heptane is twofold. First, there has been a significant amount of previous extinction limit study and resulting data with this fuel. Second, n-heptane (C7H16) is a pure substance, and therefore does not vary in composition as does JP-7, which is a mixture of several different hydrocarbons. These two facts allow for a baseline to be established by comparing the new OJB results to those previously taken. Additionally, the data set for n-heptane, which previously existed for mixtures up to 26 mole percent in nitrogen, is completed up to 100% n-heptane. The extinction limit data for the two fuels are compared, and complete experimental results are included.

Convery, Janet L.↗
NASA NTRS2016

NASA's Hydrogen Outpost: The Rocket Systems Area at Plum Brook Station

"There was pretty much a general knowledge about hydrogen and its capabilities," recalled former researcher Robert Graham. "The question was, could you use it in a rocket engine? Do we have the technology to handle it? How will it cool? Will it produce so much heat release that we can't cool the engine? These were the questions that we had to address." The National Aeronautics and Space Administration's (NASA) Glenn Research Center, referred to historically as the Lewis Research Center, made a concerted effort to answer these and related questions in the 1950s and 1960s. The center played a critical role transforming hydrogen's theoretical potential into a flight-ready propellant. Since then NASA has utilized liquid hydrogen to send humans and robots to the Moon, propel dozens of spacecraft across the universe, orbit scores of satellite systems, and power 135 space shuttle flights. Rocket pioneers had recognized hydrogen's potential early on, but its extremely low boiling temperature and low density made it impracticable as a fuel. The Lewis laboratory first demonstrated that liquid hydrogen could be safely utilized in rocket and aircraft propulsion systems, then perfected techniques to store, pump, and cleanly burn the fuel, as well as use it to cool the engine. The Rocket Systems Area at Lewis's remote testing area, Plum Brook Station, played a little known, but important role in the center's hydrogen research efforts. This publication focuses on the activities at the Rocket Systems Area, but it also discusses hydrogen's role in NASA's space program and Lewis's overall hydrogen work. The Rocket Systems Area included nine physically modest test sites and three test stands dedicated to liquid-hydrogen-related research. In 1962 Cleveland Plain Dealer reporter Karl Abram claimed, "The rocket facility looks more like a petroleum refinery. Its test rigs sprout pipes, valves and tanks. During the night test runs, excess hydrogen is burned from special stacks in the best Oklahoma oil field tradition." Besides the Rocket Systems Area, Plum Brook Station also included a nuclear test reactor, a large vacuum tank, a hypersonic wind tunnel, and a full-scale upper-stage rocket stand. The Rocket Systems Area operated from 1961 until NASA shut down all of Plum Brook in 1974. The center reopened Plum Brook in the late 1980s and continues to use several test facilities. The Rocket Systems Area, however, was not restored. Today Plum Brook resembles a nature preserve more than an oil refinery. Lush fields and forests separate the large test facilities. Until recently, the abandoned Rocket Systems Area structures and equipment were visible amongst the greenery. These space-age ruins, particularly the three towers, stood as silent sentinels over the sparsely populated reservation. Few knew the story of these mysterious facilities when NASA removed them in the late 2000s.

Rocket Systems↗
NASA NTRS2006

SACD's Support of the Hyper-X Program

NASA s highly successful Hyper-X program demonstrated numerous hypersonic air-breathing vehicle related technologies including scramjet performance, advanced materials and hot structures, GN&C, and integrated vehicle performance resulting in, for the first time ever, acceleration of a vehicle powered by a scramjet engine. The Systems Analysis and Concepts Directorate (SACD) at NASA s Langley Research Center played a major role in the integrated team providing critical support, analysis, and leadership to the Hyper-X Program throughout the program s entire life and were key to its ultimate success. Engineers in SACD s Vehicle Analysis Branch (VAB) were involved in all stages and aspects of the program, from conceptual design prior to contract award, through preliminary design and hardware development, and in to, during, and after each of the three flights. Working closely with other engineers at Langley and Dryden, as well as industry partners, roughly 20 members of SACD were involved throughout the evolution of the Hyper-X program in nearly all disciplines, including lead roles in several areas. Engineers from VAB led the aerodynamic database development, the propulsion database development, and the stage separation analysis and database development effort. Others played major roles in structures, aerothermal, GN&C, trajectory analysis and flight simulation, as well as providing CFD support for aerodynamic, propulsion, and aerothermal analysis.

Robinson, Jeffrey S.↗
NASA NTRS2007

The X-43A Flight Research Program: Lessons Learned on the Road to Mach 10

During an aerospace engineer's undergraduate studies, he or she will attend classes in aerodynamics, thermodynamics, structures, stability and control, dynamics, design, propulsion, and computer science, along with the related courses in mathematics, physics, statistics, and chemistry required to understand the material. Upon graduation, the new engineer will have acquired a basic knowledge of how to build an aerospace vehicle. What only comes through experience, however, is the understanding of the inevitable imperfect process through which an aerospace vehicle is built. This is the adventure of turning a basic concept into functional hardware. Engineers working on a project must often deal with ambiguous situations. They are routinely asked by management to provide risk assessments of a project, yet even after careful analysis uncertainties remain. The project must be accomplished within finite limits of time and money. The question an engineer answers is whether the solution to potential problem is worth the cost and schedule delay, or if the solution might actually be worse than the problem it is meant to solve. Review protocols are established to ensure that an unknown has not been overlooked. But these cannot protect against an unknown unknown. Examples of these situations can be found in the history of the X-43A Hyper-X (Hypersonic Experiment) program. In this NASA project, a supersonic combustion ramjet (scramjet) engine was flight tested on a subscale vehicle. The X-43A Hyper-X Research Vehicle (HXRV) was launched from a B-52B mothership, then boosted to the test speed by a modified Pegasus rocket first stage, called the Hyper-X Launch Vehicle (HXLV). Once at the proper speed and altitude, the X-43A separated from the booster, stabilized itself, and then the engine test began. Although wind-tunnel scramjet engine tests had begun in the late 1950s, before the Hyper-X program there had never been an actual in-flight test of such an engine integrated with an appropriate airframe. Thus, while the scramjet had successfully operated in the artificial airflow of wind tunnels, the concept had yet to be proven in real air. These conditions meant changes in density and temperature, as well as changes in angle of attack and sideslip of a free-flying vehicle. A wind tunnel is limited in its ability to simulate these subtle factures, which have a major impact on almost any vehicle, but especially that of a scramjet's performance. The Hyper-X project was to provide a real-world benchmark of the ground test data. The full scale X-43A engine would be operated in the wind tunnel, and then flown, and the data from its operation would then be compared with projections. If these matched, the wind tunnel data would be considered a reliable design tool for future scramjet. If there were significant differences, the reasons for these would have to be identified. Until such information was available, scramjets would lack the technological maturity to be considered for future space launch or high-speed atmospheric flight vehicles.

Peebles, Curtis↗
NASA NTRS1994

Particle kinetic simulation of high altitude hypervelocity flight

Rarefied flows about hypersonic vehicles entering the upper atmosphere or through nozzles expanding into a near vacuum may only be simulated accurately with a direct simulation Monte Carlo (DSMC) method. Under this grant, researchers enhanced the models employed in the DSMC method and performed simulations in support of existing NASA projects or missions. DSMC models were developed and validated for simulating rotational, vibrational, and chemical relaxation in high-temperature flows, including effects of quantized anharmonic oscillators and temperature-dependent relaxation rates. State-of-the-art advancements were made in simulating coupled vibration - dissociation - recombination for post-shock flows. Models were also developed to compute vehicle surface temperatures directly in the code rather than requiring isothermal estimates. These codes were instrumental in simulating aerobraking of NASA's Magellan spacecraft during orbital maneuvers to assess heat transfer and aerodynamic properties of the delicate satellite. NASA also depended upon simulations of entry of the Galileo probe into the atmosphere of Jupiter to provide drag and flow field information essential for accurate interpretation of an onboard experiment. Finally, the codes have been used extensively to simulate expanding nozzle flows in low-power thrusters in support of propulsion activities at NASA-Lewis. Detailed comparisons between continuum calculations and DSMC results helped to quantify the limitations of continuum CFD codes in rarefied applications.

Boyd, Iain↗
NASA NTRS2000

Navigation Strategy for the Mars 2001 Lander Mission

The Mars Surveyor Program (MSP) is an ongoing series of missions designed to robotically study, map and search for signs of life on the planet Mars. The MSP 2001 project will advance the effort by sending an orbiter, a lander and a rover to the red planet in the 2001 opportunity. Each vehicle will carry a science payload that will Investigate the Martian environment on both a global and on a local scale. Although this mission will not directly search for signs of life, or cache samples to be returned to Earth, it will demonstrate certain enabling technologies that will be utilized by the future Mars Sample Return missions. One technology that is needed for the Sample Return mission is the capability to place a vehicle on the surface within several kilometers of the targeted landing site. The MSP'01 Lander will take the first major step towards this type of precision landing at Mars. Significant reduction of the landed footprint will be achieved through two technology advances. The first, and most dramatic, is hypersonic aeromaneuvering; the second is improved approach navigation. As a result, the guided entry will produce in a footprint that is only tens of kilometers, which is an order of magnitude improvement over the Pathfinder and Mars Polar Lander ballistic entries. This reduction will significantly enhance scientific return by enabling the potential selection of otherwise unreachable landing sites with unique geologic interest and public appeal. A landed footprint reduction from hundreds to tens of kilometers is also a milestone on the path towards human exploration of Mars, where the desire is to place multiple vehicles within several hundred meters of the planned landing site. Hypersonic aeromaneuvering is an extension of the atmospheric flight goals of the previous landed missions, Pathfinder and Mars Polar Lander (MPL), that utilizes aerodynamic lift and an autonomous guidance algorithm while in the upper atmosphere. The onboard guidance algorithm will control the direction of the lift vector, via bank angle modulation, to keep the vehicle on the desired trajectory. While numerous autonomous guidance algorithms have been developed for use during hypersonic flight at Earth, this will be the first flight of an autonomously directed lifting entry vehicle at Mars. However, without sufficient control and knowledge of the atmospheric entry conditions, the guidance algorithm will not perform effectively. The goal of the interplanetary navigation strategy is to deliver the spacecraft to the desired entry condition with sufficient accuracy and knowledge to enable satisfactory guidance algorithm performance. Specifically, the entry flight path angle must not exceed 0.27 deg. to a 3 sigma confidence level. Entry errors will contribute directly to the size of the landed footprint and the most significant component is entry flight path angle. The size of the entry corridor is limited on the shallow side by integrated heating constraints, and on the steep side by deceleration (g-load) and terminal descent propellant. In order to meet this tight constraint it is necessary to place a targeting maneuver seven hours prior to the time of entry. At this time the trajectory knowledge will be quite accurate, and the effects of maneuver execution errors will be small. The drawback is that entry accuracy is dependent on the success of this final late maneuver. Because propulsive maneuvers are critical events, it is desirable to minimize their occurrence and provide the flight team with as much response time as possible in the event of a spacecraft fault. A mission critical maneuver at Entry - 7 hours does not provide much fault tolerance, and it is desirable to provide a strategy that minimizes reliance on this maneuver. This paper will focus on the Improvements in interplanetary navigation that will decrease entry errors and will reduce the landed footprint, even in the absence of aeromaneuvering. The easiest to take advantage of are Improvements In the knowledge of the Mars ephemeris and gravity field due to the MGS and MSP'98 missions. Improvements In data collection and reduction techniques such as "precislon ranging' and near-simultaneous tracking will also be utilized. In addition to precise trajectory control, a robust strategy for communications and flight operations must also be demonstrated. The result Is a navigation and communications strategy on approach that utilizes optimal maneuver placement to take advantage of trajectory knowledge, minimizes risk for the flight operations team, is responsive to spacecraft hardware limitations, and achieves the entry corridor. The MSP2001 mission Is managed at JPL under the auspices of the Mars Exploration Directorate. The spacecraft flight elements are built and managed by Lockheed-Martin Astronautics in Denver, Colorado.

Mase, Robert A.↗
NASA NTRS1994

Particle kinetic simulation of high altitude hypervelocity flight

Rarefied flows about hypersonic vehicles entering the upper atmosphere or through nozzles expanding into a near vacuum may only be simulated accurately with a direct simulation Monte Carlo (DSMC) method. Under this grant, researchers enhanced the models employed in the DSMC method and performed simulations in support of existing NASA projects or missions. DSMC models were developed and validated for simulating rotational, vibrational, and chemical relaxation in high-temperature flows, including effects of quantized anharmonic oscillators and temperature-dependent relaxation rates. State-of-the-art advancements were made in simulating coupled vibration-dissociation recombination for post-shock flows. Models were also developed to compute vehicle surface temperatures directly in the code rather than requiring isothermal estimates. These codes were instrumental in simulating aerobraking of NASA's Magellan spacecraft during orbital maneuvers to assess heat transfer and aerodynamic properties of the delicate satellite. NASA also depended upon simulations of entry of the Galileo probe into the atmosphere of Jupiter to provide drag and flow field information essential for accurate interpretation of an onboard experiment. Finally, the codes have been used extensively to simulate expanding nozzle flows in low-power thrusters in support of propulsion activities at NASA-Lewis. Detailed comparisons between continuum calculations and DSMC results helped to quantify the limitations of continuum CFD codes in rarefied applications.

Boyd, Iain↗
NASA NTRS2020

Parametric Cost Estimates of Four 20 Ton Payload Mars EDL Vehicle Concepts

Cost is one of the biggest obstacles to sending humans to Mars. Many studies have examined various vehicles and/or technologies which could be used to send humans to Mars. However, nearly all of these studies have either ignored cost or have used the mass of the vehicles as a surrogate for cost. This work presents parametric cost estimates of four Mars Entry, Descent, and Landing vehicle concepts capable of landing 20 tons of payload on the surface of Mars. The concepts are the Co-Optimization Blunt-body Re-entry Analysis-Mid lift to drag ratio Rigid Vehicle (Cobra–MRV), Adaptive Deployable Entry and Placement Technology (ADEPT), Hypersonic Inflatable Aerodynamic Decelerator (HIAD), and Capsule concepts. Two variants of the HIAD concept are studied, one with a single HIAD heatshield and one with two HIADs. The study presents the development and production cost of each vehicle and the total cost to build five flight units and fifteen flight units. Five flight units is representative of a single mission, or a "Flags and Footprints" style campaign, whereas fifteen flight units is representative of multiple missions going to multiple sites on Mars. All costs presented in this work are normalized to the average of the first unit production cost of the lander concepts. There is a large amount of uncertainty both in the design of the vehicles and the cost of the vehicles and it is impossible to claim with certainty which vehicle will cost the least. However, it is likely that the total cost to produce five flight units will be in the following ranking from least costly to most costly, Single HIAD, Capsule, Dual HIAD, Cobra, ADEPT. The ranking for fifteen units is similar but with the positions of the Dual HIAD and Cobra reversed. This study only considers the cost of the entry vehicle and does not take into account the effect of the entry vehicle’s mass on the cost of the in-space propulsion stage or launch vehicle.

Paul D Friz↗
NASA NTRS2016

HIAD Advancements and Extension of Mission Applications

The Hypersonic Inflatable Aerodynamic Decelerator (HIAD) technology has made significant advancements over the last decade with flight test demonstrations and ground development campaigns. The first generation (Gen-1) design and materials were flight tested with the successful third Inflatable Reentry Vehicle Experiment flight test of a 3-m HIAD (IRVE-3). Ground development efforts incorporated materials with higher thermal capabilities for the inflatable structure (IS) and flexible thermal protection system (F-TPS) as a second generation (Gen-2) system. Current efforts and plans are focused on extending capabilities to improve overall system performance and reduce areal weight, as well as expand mission applicability. F-TPS materials that offer greater thermal resistance, and ability to be packed to greater density, for a given thickness are being tested to demonstrated thermal performance benefits and manufacturability at flight-relevant scale. IS materials and construction methods are being investigated to reduce mass, increase load capacities, and improve durability for packing. Previous HIAD systems focused on symmetric geometries using stacked torus construction. Flight simulations and trajectory analysis show that symmetrical HIADs may provide L/D up to 0.25 via movable center of gravity (CG) offsets. HIAD capabilities can be greatly expanded to suit a broader range of mission applications with asymmetric shapes and/or modulating L/D. Various HIAD concepts are being developed to provide greater control to improve landing accuracy and reduce dependency upon propulsion systems during descent and landing. Concepts being studied include a canted stack torus design, control surfaces, and morphing configurations that allow the shape to be actively manipulated for flight control. This paper provides a summary of recent HIAD development activities, and plans for future HIAD developments including advanced materials, improved construction techniques, and alternate geometry concepts that will greatly expand HIAD mission applications.

Johnson, R. Keith↗
DOE OSTI2026

Topographic effects on reflected acoustic waves from the OSIRIS-REx reentry observed from stratospheric balloons

During long-distance sound propagation in planetary atmospheres, acoustic waves may reflect off the air/surface interface one or more times. For low sound frequencies and flat interfaces, the incident and reflected wave tend to be nearly identical. However, this may not be the case when the downgoing acoustic wave encounters topography. Here, we describe a set of direct and reflected acoustic signals recorded on free-flying balloons during the hypersonic entry of the OSIRIS-REx sample return capsule (SRC). In two of the three cases presented here, an impulsive reflected arrival similar in form to the direct sonic boom of the SRC is observed, followed by a diffuse coda. In contrast, one of the floating stations lacked an impulsive reflection entirely, with only the coda present. We use full waveform modeling to show how reflection in the presence of complex topography can explain coda seen in all three examples as well as the lack of impulsive arrival on the third. Our results indicate that the complex signals often observed in long range acoustic propagation could be due, in part, to interactions with topography during transmission.

Lees, Jonathan M. [University of North Carolina, C↗
DOE OSTI2023

Hypersonic Fluid–Structure Interaction on a Cone–Slice–Ramp Geometry

In this study, fluid–structure interactions were measured between a representative control surface and the hypersonic flow deflected by it. The control surface is simplified as a spanwise finite ramp placed on a longitudinal slice of a cone. The front surface of the ramp contains a thin panel designed to respond to the unsteady fluid loading arising from the shock-wave/boundary-layer interactions. Experiments were conducted at Mach 5 and Mach 8 with ramps of different angles. High-speed schlieren captured the unsteady flow dynamics and accelerometers behind the thin panel measured its structural response. Panel vibrations were dominated by natural modes that were excited by the broadband aerodynamic fluctuations arising in the flowfield. However, increased structural response was observed in two distinct flow regimes: 1) attached or small separation interactions, where the transitional regime induced the strongest panel fluctuations. This was in agreement with the observation of increased convective undulations or bulges in the separation shock generated by the passage of turbulent spots, and 2) large separated interactions, where shear layer flapping in the laminar regime produced strong panel response at the flapping frequency. In addition, panel heating during the experiment caused a downward shift in its natural mode frequencies.

33 ADVANCED PROPULSION SYSTEMS↗
NASA NTRS1993

Towards a generalized computational fluid dynamics technique for all Mach numbers

Currently there exists no single unified approach for efficiently and accurately solving computational fluid dynamics (CFD) problems across the Mach number regime, from truly low speed incompressible flows to hypersonic speeds. There are several CFD codes that have evolved into sophisticated prediction tools with a wide variety of features including multiblock capabilities, generalized chemistry and thermodynamics models among other features. However, as these codes evolve, the demand placed on the end user also increases simply because of the myriad of features that are incorporated into these codes. In order for a user to be able to solve a wide range of problems, several codes may be needed requiring the user to be familiar with the intricacies of each code and their rather complicated input files. Moreover, the cost of training users and maintaining several codes becomes prohibitive. The objective of the current work is to extend the compressible, characteristic-based, thermochemical nonequilibrium Navier-Stokes code GASP to very low speed flows and simultaneously improve convergence at all speeds. Before this work began, the practical speed range of GASP was Mach numbers on the order of 0.1 and higher. In addition, a number of new techniques have been developed for more accurate physical and numerical modeling. The primary focus has been on the development of optimal preconditioning techniques for the Euler and the Navier-Stokes equations with general finite-rate chemistry models and both equilibrium and nonequilibrium thermodynamics models. We began with the work of Van Leer, Lee, and Roe for inviscid, one-dimensional perfect gases and extended their approach to include three-dimensional reacting flows. The basic steps required to accomplish this task were a transformation to stream-aligned coordinates, the formulation of the preconditioning matrix, incorporation into both explicit and implicit temporal integration schemes, and modification of the numerical flux formulae. In addition, we improved the convergence rate of the implicit time integration schemes in GASP through the use of inner iteration strategies and the use of the GMRES (General Minimized Resisual) which belongs to the class of algorithms referred to as Krylov subspace iteration. Finally, we significantly improved the practical utility of GASP through the addition of mesh sequencing, a technique in which computations begin on a coarse grid and get interpolated onto successively finer grids. The fluid dynamic problems of interest to the propulsion community involve complex flow physics spanning different velocity regimes and possibly involving chemical reactions. This class of problems results in widely disparate time scales causing numerical stiffness. Even in the absence of chemical reactions, eigenvalue stiffness manifests itself at transonic and very low speed flows which can be quantified by the large condition number of the system and evidenced by slow convergence rates. This results in the need for thorough numerical analysis and subsequent implementation of sophisticated numerical techniques for these difficult yet practical problems. As a result of this work, we have been able to extend the range of applicability of compressible codes to very low speed inviscid flows (M = .001) and reacting flows.

Walters, R. W.↗
NASA NTRS2000

Mission Design Overview for Mars 2003/2005 Sample Return Mission

In May 2003, a new and exciting chapter in Mars exploration will begin with the launch of the first of three spacecraft that will collectively contribute toward the goal of delivering samples from the Red Planet to Earth. This mission is called Mars Sample Return (MSR) and will utilize both the 2003 and 2005 launch opportunities with an expected sample return in October 2008. NASA and CNES are major partners in this mission. The baseline mission mode selected for MSR is Mars orbit rendezvous (MOR), analogous in concept to the lunar orbit rendezvous (LOR) mode used for Apollo in the 1960s. Specifically, MSR will employ two NASA-provided landers of nearly identical design and one CNES-provided orbiter carrying a NASA payload of rendezvous sensors, orbital capture mechanisms, and an Earth entry vehicle (EEV). The high-level concept is that the landers will launch surface samples into Mars orbit, and the orbiter will retrieve the samples in orbit and then carry them back to Earth. The first element to depart for Mars will be one of the two landers. Currently, it is proposed that an intermediate class launch vehicle, such as the Boeing Delta 3 or Lockheed Martin Atlas 3A, will launch this 1800-kg lander from Cape Canaveral during the May 2003 opportunity. The lander will utilize a Type-1 transfer trajectory with an arrival at Mars in mid-December 2003. Landing will be aided by precision approach navigation and a guided hypersonic entry to achieve a touchdown accuracy of 10 km or better. Although the exact landing site has not yet been determined, it is estimated that lander resource constraints will limit the site to between 15 degrees north and south latitudes. Following touchdown, the lander will deploy a six-wheeled, 60-kg rover carrying an extensive suite of instruments designed to aid in the analysis of the local terrain and collection of core samples from selected rocks. The surface mission is currently designed around a concept called the surface traverse. Each traverse will involve the rover exploring a selected area of terrain up to 100 meters from the lander, the collection of rock core samples, and the delivery of the samples from the traverse back to a sample canister on the lander. Planning estimates indicate that up to three traverses may be possible during the expected 90-sol lifetime of the lander. The canister that will receive the samples from the rover will be attached to the top stage of a small solid-fueled rocket mounted to the deck of the lander. This rocket is called the Mars Ascent Vehicle (MAV) and consists of three stages weighing a total of about 140 kg. After the conclusion of the surface mission, the MAV will lift-off and insert the sample canister into a near-circular orbit with an altitude of about 600 km and inclination of 45 degrees. The sample canister will wait in this orbit until it is retrieved by the orbiter sometime in early 2007. In August 2005, the second lander and a CNES-provided orbiter weighing 2700 kg will depart for Mars. Currently, it is proposed that a single Ariane 5 provided by CNES will launch both of these two elements onto a Type-2 transfer trajectory. Although the orbiter and lander will be launched together, they will separate shortly after injection and will fly to Mars as two independent spacecraft. However, both spacecraft will perform a maneuver between 10 and 15 days after launch so that their arrival times at Mars differ by between 12 and 24 hours. This scheme will reduce the operational complexity at the encounter date. A set of four 60-kg surface probes will ride piggyback on the orbiter to Mars. These CNES-provided probes are called Netlanders and will serve as surface stations for scientific investigations independent of the Mars Sample Return goals. Starting approximately one month prior to arrival at Mars, the orbiter will begin to release the Netlanders one at a time. Each release cycle will take several days, and will include time for precision navigation to execute one or two maneuvers that will target the Netlanders to their proper landing site. All four deployment cycles will be completed prior to 10 days before arrival. Both the orbiter and lander will arrive in late-July 2006. Upon arrival, the lander will perform a precision landing and surface mission similar in concept to the one that was executed during the 2003 opportunity. Although the landing site for the 2005 opportunity has not been selected, it is expected to be different from the 2003 site to enhance the diversity of the collected samples. The orbiter's arrival at Mars will be highlighted by the first use of aerocapture to insert a spacecraft into a capture orbit around another planet. The choice of aerocapture, as opposed to a propulsive orbit insertion, was considered mission enabling due to a reduction of over 2000 m/s in mission AV. Aerocapture will be targeted to produce a 250 km x 1400 km capture orbit with an inclination of 45 degrees. Current analysis indicates that achieving this goal will require approximately six minutes of flight deep in the atmosphere with a targeted periapsis of approach of about 43 km. After factoring into account the penalty for carrying a heat shield to survive aerocapture, the net savings compared to a propulsive orbital insertion amounts to several hundred kilograms.

Lee, Wayne J.↗