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The physical basis and application of the fundamental relationship governing the balance and utilization of available energy for a chemical rocket operating within the atmosphere are described. The...

Small, low-thrust, long-burn-time solid propellant rocket motors could provide propulsion for a new class of kilogram-scale, transonic, uncrewed aerial vehicles (UAVs). This paper investigates technological challenges of small, low-thrust solid rocket motors: slow-burn solid propellants, motors that have low thrust relative to their size (and thus have low chamber pressure), thermal protection for the motor case, and small nozzles that can withstand long burn times. Slow-burn propellants were developed using ammonium perchlorate and 0–20% oxamide (burn-rate suppressant), with burn rates of 1–4 mm⋅s−1 at 1 MPa. Using these propellants, a low-thrust motor successfully operated at a thrust/burn area ratio 10 times less than that of typical solid rocket motors. This kilogram-scale motor can provide 5–10 N of thrust for 1–3 min. An ablative thermal protection liner was tested in these firings, and a new ceramic-insulated nozzle was demonstrated. This paper shows that small, low-thrust solid motors are feasible and presents a baseline design for the integration of such a motor into a small UAV., Department of Defense (DoD), MIT Lincoln Laboratories, BAE Systems, Inc.

This paper evaluates various sources of oxidizer to fuel mass ratio (O/F) shifts in hybrid rockets and paths (physical phenomena) through which these O/F shifts affect flight performance. Moreover, the performance increase of O/F control in hybrid rockets is evaluated. Vertical launches of O/F uncontrolled and O/F controlled of hybrid sounding rockets were simulated under two uncertainty models of fuel regression behavior based on experimental data: a) systematic errors with a constant deviation within ±3σ±3σ and 2) random errors subject to a probability distribution. These simulations included all sources of O/F shifts that originated in the fuel regression behavior and all paths through which the O/F shifts affect flight performance. Residual propellant mass and decreases in specific impulse are found to be the dominant causes of performance loss under both uncertainty models. For both cases 1 and 2, the O/F-controlled hybrid rockets maintained the performance expected under nominal fuel regression behavior, whereas the O/F-uncontrolled hybrid rockets had a lower performance by upwards of 6.69 and 4.06% in ΔVΔV for cases 1 and 2, respectively. For case 2, 3008 flight simulations revealed that the worst case of the O/F-controlled hybrid rocket had a 4.06 to 4.49% larger ΔVΔV and 10.5 to 13.3% higher apogee than that of the O/F-uncontrolled hybrid rocket, and that the O/F-uncontrolled hybrid rocket had a 6.61 times larger standard deviation in ΔVΔV. These results mean that the elimination O/F shift in hybrid rockets significantly improves performance, as well as the accuracy and reliability of performance predictions.

Although several investigations have been performed regarding the fundamental physics of axial-injection, end-burning hybrid rocket motors, investigations into the mission performance of a propulsi...

Small satellites may provide a low-cost platform for targeted science investigations in the Mars system. With current technology, small satellites require ride shares with larger orbiters to captur...

Properly determining the size and predicting the performance of a Hall-effect thruster (HET) is an essential process in the early phase of thruster design. In this paper, the scaling relations for ...

A general perturbation solution to a restricted low-thrust Lambert rendezvous problem, considering circular-to-circular in-plane maneuvers using tangential thrust and including a coast arc, is developed. This provides a fully analytical solution to the satellite reconnaissance problem. The solution requires no iteration. Its speed and simplicity allow problems involving numerous spacecraft and maneuvers to be studied; this is demonstrated through two case studies. In the first, a range of maneuvers providing a rapid flyover of Los Angeles is generated, giving an insight to the trade space and allowing the maneuver that best fulfills the mission to be selected. A reduction in flyover time from 13.8 to 1.6 days is possible using a less than 17 m/s velocity change. A comparison with a numerical propagator including atmospheric friction and an 18th-order tesseral model shows 4 s of difference in the time of flyover. A second study considers a constellation of 24 satellites that can maneuver into repeating ground track orbits to provide persistent coverage of a region. A set of maneuvers for all satellites is generated for four sequential targets, allowing the most suitable maneuver strategy to be selected. Improvements in coverage of greater than 10 times are possible as compared to a static constellation using 35% of the propellant available across the constellation.

A simplified theoretical treatment of rotating detonation engine (RDE) performance is developed for rocket applications with traceability to current combustion technology. In particular, the influe...

The dawn of "New Space" in recent years is changing the landscape of the space industry. In particular, the shift to smaller satellites, requiring shorter development time s and using off-the-shelf-components and standardized buses, has led to a continuing reduction in spacecraft cost. However, launch costs remain extremely high and frequently dominate the total mission cost. Additionally, many small satellites are designed to operate as part of a larger constellation, but traditional launch methods require a difference dedicated launch for each orbit plane to be populated. This need for multiple costly launches can stifle, and even prohibit, some missions requiring numerous orbit planes as the launch cost increases beyond what can be justified for the mission. As of 2014, most smallsats, including CubeSats, have been launched on opportunistic ‘rideshare’ or ‘piggy-back’ launches, in which the spacecraft shares its launch with other craft, often as a secondary payload. This has the advantage of providing a cheaper launch but restricts the operator’s choice of orbit, which will affect the system performance.

In-space propellant supply chain can be effectively established for manned missions if high-thrust crew vehicles and cargo tugs can be used in conjunction with low-thrust cargo tugs. Optimizing the...

Analysis was performed to assess the impact of atmospheric-breathing supersonic retropropulsion as a technology solution for Mars descent. Vehicle models were developed for three architectures, emp...

Multimode spacecraft micropropulsion systems that include a high-thrust chemical mode and high-specific impulse electric mode are assessed with specific reference to CubeSat-sized satellite applica...

Onboard, fuel-optimal, constrained powered-descent guidance based on the theory of lossless convexification has been implemented as the Guidance for Fuel-Optimal Large Diverts (G-FOLD) algorithm. Here, “guidance” means generating feedforward reference trajectories for control systems. This paper presents terrestrial flight-test demonstrations of large diverts planned by G-FOLD onboard a vertical-takeoff/vertical-landing rocket. The G-FOLD parser, which transforms the guidance problem into a second-order cone program and so encodes the divert constraints, is described at an engineering level, including new and modified constraints incorporated for these flight tests. Several practical issues, such as discretization effects, are addressed, and the flight-test architecture is presented. A total of eight flight tests were performed. In the first three, the rocket executed diverts of increasing size preplanned by G-FOLD on the ground. Then G-FOLD was demonstrated running onboard five times. Three qualitatively...

The orbital maneuver capabilities of several CubeSat propulsion systems are analyzed using trajectory simulations. Properties of several types of developmental micropropulsion systems are reviewed, and ΔV capabilities are compared. Mission simulations are used to analyze the relationship between thrust arc length and orbit change capability in a low-thrust spiral trajectory. Constraints on power, fuel mass, and mission duration, as well as system-level constraints, are considered. Feasible CubeSat architectures and mission designs are developed for three electric propulsion systems. The most effective combinations of thruster operational modes and trajectory control strategies are discussed.

Landing heavier payloads with pinpoint precision on the surface of Mars is a known challenge for future missions to Mars. Supersonic retropropulsion has been proposed as a means to deliver higher-mass payloads to the surface, traditionally with the sole intent of deceleration. By adding range control capability during the supersonic retropropulsion maneuver by means of thrust vectoring, it has been found that a substantial amount of propellant can be saved in a precision landing scenario when compared with traditional architectures. Decreasing the propellant mass necessary for the mission increases the amount of payload mass that can be brought to the surface. Propellant mass savings greater than 30% are possible if thrust vectoring is unconstrained during the supersonic phase of flight. Propellant mass fraction is found to be sensitive to the divert direction and also the altitude and flight-path angle at ignition, favoring low altitudes and shallow flight-path angles. Decreased nozzle cant angles and aerodynamic drag preservation have also been found to reduce propellant usage.

Aerocapture is the maneuver by an interplanetary spacecraft to fly through the atmosphere of a planet with the aim of attaining a specified orbit around the planet. By appropriately controlling the...

This paper presents an innovative space mission devoted to the survey of the small Earth companion asteroid by means of nano platforms. Also known as the second Earth moon, Cruithne, is the target identified for the mission. Both the trajectory to reach the target and a preliminary spacecraft budget are here detailed. The idea is to exploit high efficient ion thrusters to reduce the propellant mass fraction in such a high total impulse mission (of the order of 1e6 Ns). This approach allows for a 100 kg class spacecraft with a very small Earth escape energy (5 km2/s2) to reach the destination in about 320 days. The 31% propellant mass fraction allows for a payload mass fraction of the order of 8% and this is sufficient to embark on such a small spacecraft a couple of nano-satellites deployed once at the target to carry out a complete survey of the asteroid. Two 2U Cubesats are here considered as representative payload, but also other scientific payloads or different platforms might be considered according with the specific mission needs. The small spacecraft used to transfer these to the target guarantees the manoeuvre capabilities during the interplanetary journey, the protection against radiations along the path and the telecommunication relay functions for the data transmission with Earth stations. The approach outlined in the paper offers reliable solutions to the main issues associated with a deep space nano-satellite mission thus allowing the exploitation of distant targets by means of these tiny spacecraft. The study presents an innovative general strategy for the NEO observation and Cruithne is chosen as test bench. This target, however, mainly for its relevant inclination, requires a relatively large propellant mass fraction that can be reduced if low inclination asteroids are of interest. This might increase the payload mass fraction (e.g. additional Cubesats and/or additional scientific payloads on the main bus) for the same 100 kg class mission.

Reducing the future threat of orbital debris will require active removal of the larger objects, such as old satellites and upper stages. Material ablation by intense laser radiation, which can be applied to most materials, provides a unique means of solving this problem. Such ablation will allow the tug to use the inert mass of the debris itself as propellant for the deorbit process, greatly reducing the mass launched to orbit. It will further allow the tug to zero the rotation of the debris from a distance of several meters, reducing the risk associated with docking. This paper looks at some basic performance parameters for a laser ablation tug to show that a tug could use laser ablation propulsion to circumvent the high launch mass and some of the docking risks that are problematic for more traditional tugs.

Supersonic retropropulsion is an entry, descent, and landing technology applicable to and potentially enabling high-mass missions required for advanced robotic and human exploration on the surface of Mars. For conceptual design, it is necessary to understand the significance of retropropulsion configuration on an entry vehicle’s static aerodynamic characteristics and the relation of this configuration to other vehicle performance metrics. This investigation developed an approximate model for the supersonic retropropulsion flowfield to assist in evaluating the impact of design choices on the vehicle’s drag characteristics for flight-relevant conditions and scales. This model was used to explore the impact of operating conditions, required propulsion system performance, propulsion system composition, and vehicle configuration on the integrated aerodynamic drag characteristics of full-scale vehicles for Mars entry, descent, and landing. The forebody aerodynamic drag and axial force characteristics of vehicle...

A systematic methodology is presented for defining a minimum functionality baseline configuration of a human spacecraft. To estimate a lower bound for the spacecraft mass, a set of essential functions is coupled to single-string subsystems with zero fault tolerance. This minimum functionality baseline is defined to meet the physical requirements needed to transport the crew to the target destination and to ensure that their physiological needs are met, but without margin, dispersions, redundancy, or factor of safety. This constitutes a set of nonnegotiable requirements based on fundamental parameters derived from physics and physiology. By definition, this represents a technically feasible solution, but it results in the highest-risk design. Mass additions beyond the minimum functional configuration are allocated to increased safety or operability through the addition of component redundancy, fault tolerance, factor of safety, additional mission functionality, and improved human–system interfaces. This pr...

The effects of anode temperature on the performance of a 4.5 kW Hall-effect thruster are investigated. The approach separates the location of gas injection from discharge current collection using an anode band, which removes the main mechanism that heats the gas distributor. A thermal model predicts a 270 K reduction of the gas distributor temperature, which corresponds to a 28% increase in the propellant residence time. Collection of the discharge current on the anode band, which is upstream of the bulk Hall current region, generates a 10% increase in ion current density at the thruster centerline for discharge voltages of 100, 125, and 150 V at a xenon mass flow rate of 5 mg/s. The initial reduction in neutral velocity with the anode band is counteracted by the influence of the channel wall temperature, which increases the neutral velocity of the particles by up to 25% greater than the velocity at the gas distributor exit plane. This reduces the potential thruster efficiency improvement from 5.5 to 2.5...

A new electric propulsion device concept takes advantage of residual helium gas that is trapped in the chemical propellant feed system and currently unused at end of life. The helium ion thruster provides additional propellant resources to extend satellite lifetimes and transfer geostationary orbit space assets to ultrasafe disposal orbits. The predicted capability, if fully allotted to the disposal, allows for perigee heights above the geostationary altitude that are one order of magnitude greater than existing international guidelines of 250 km. Furthermore, the proposed helium ion thruster concept makes the classic propellant gauge uncertainty problem moot, as satellite operators could use all of their conventional propellant for nominal station-keeping operations. The helium ion thruster concept therefore mitigates future space debris arising from depleted assets in the geostationary orbit belt through both aggressive orbit raising and depressurization of satellites at end of life. An analysis of the helium ion thruster theoretical performance shows that the device could raise the altitude of an end-of-life 2500 kg, 5 kW spacecraft by 2200 km in two months using 2 kg of residual helium.

Human space exploration vehiclesmust be designedwith a reliable and safe launch escape system to be activated in the event of a failure on the pador at altitude.A safe, reliable, highlyflexible, and adaptable launch escape system that can be used with most proposed commercial space capsules was conceived after conducting propulsion and configuration trade studies. The outcome of these studies is a novel gas–gas propulsion system with a tractorconfiguration-based launch escape system. Preliminary design of this joint universal launch escape and assist system is detailed. It is designed to accelerate a crew capsule away froma launch vehicle in case of an emergency on the pador during the early portions of the trajectory at up to approximately 300,000 ft of altitude. It is designed to tractor the spacecraft to a sufficient altitude andwith enough downrange translation so that the parachute landing system of the capsule can safely function. During midand high-altitude aborts, a reentry gravity-reduction mode of this novel escape system can start and stop its high-pressure gaseous oxygen and gaseousmethane engines and can direct thrust in such a way as to shape the reentry trajectory to reduce the abort reentry deceleration. Finally, it has an ascentassistmode that can be used to offset the normal payload penalty in the event that the emergency abort function is not used.

A hybrid rocket is considered as the third stage of a three-stage launcher. The propulsion system design and the trajectory are simultaneously optimized by means of a nested direct/indirect procedure. Direct optimization of the parameters that affect the motor design is coupled with indirect trajectory optimization to maximize the launcher payload for assigned conditions at the stage ignition and final orbit. A mission profile based on the Vega launcher is considered. The feed system exploits a pressurizing gas, namely helium, with hydrogen peroxide as the oxidizer and polyethylene as the fuel. The simplest blowdown design is compared with a more complex pressurizing system, which has an additional gas tank that allows for a phase with constant oxidizer tank pressure. The optimization provides the optimal values of the main engine design parameters (pressurizing gas mass, nozzle expansion ratio, and initial values of tank pressure, mixture ratio and thrust), the corresponding grain and engine geometry, and the control law (thrust direction during the ascent trajectory and engine switching times). Results show that a hybrid rocket may be a viable option for small launchers.

This study investigates the performance and feasibility of a new entry, descent, and landing architecture onMars, termed Smart Divert, for landing in one of a number of small safe zones surrounded by hazardous terrain. Smart Divert consists of a ballistic entry followed by supersonic parachute deployment. After parachute release, the vehicle diverts to one ofmany predefined, fuel-optimal safe zone sites. The Smart Divert concept does not require hypersonic guidance or real-time terrain recognition. Instead, it relies on a priori orbital observations to identify small, multiple safe zones within a larger hazardous region and additional terminal descent propellant to land at the fuel-optimal safe zone.Before launch,mission designers could trade thenumber and size of the safe zones as part of the landing site selection process.Reasonable propellantmass fractions of 0.3 canbe achievedby initiating the divert at 5 kmaltitude, providing a 10 km horizontal divert capability. The number of safe zones is shown to be a function of landing ellipse size. Assuming Mars Science Laboratory state-of-the-art interplanetary navigation, four safe zone sites, randomly placed throughout the landing ellipse to simulate unknowndestinations of futuremissions, require a propellantmass fraction less than 0.3 for 97% of the cases analyzed. The unconstrained optimal arrangement of four safe zone sites within the same landing ellipse reduced the required propellant mass fraction from 0.3 to 0.22. The propellant mass fraction may be further reduced as the number of safe zone sites is increased. An example scenario using rock count data for the Phoenix landing site region demonstrates that Smart Divert can be implemented to land in previously unreachable terrain for a propellant mass fraction of 0.2.

A wide range of advanced launch concepts have been proposed in an effort to revolutionize space access through either a significant reduction in launch costs or significant improvements in launch performance. This paper briefly summarizes commonly proposed advanced launch concepts, including both concepts that employ propellant and propellantless concepts. Each concept is briefly described along with its potential in two generic mission classes: small satellite launch to LEO and large satellite launch to GEO. It is shown theoretically that there is significant room for improvement in the cost and performance of current launch systems. It is also shown, however, that historical predictions of launch costs reductions and/or performance improvements for new technologies have been highly optimistic with realized costs and performance leading to only incremental improvements instead of revolutionary advancements. All of the reviewed technologies still have significant technical challenges to overcome before yielding fully operational systems. The associated risk makes it difficult to justify the large investments required to develop such systems, indicating that a development path with useful products at points in between the current state-of-the-art and final goal is necessary.

Landing site selection is a compromise between safety concerns associated with the site’s terrain and scientific interest. Therefore, technologies enabling pinpoint landing performance (sub-100-m accuracies) on the surface of Mars are of interest to increase the number of accessible sites for in situ research, as well as allow placement of vehicles nearby prepositioned assets. A survey of the performance of guidance, navigation, and control technologies that could allow pinpoint landing to occur at Mars was performed. This assessment has shown that negligible propellant mass fraction benefits are seen for reducing the three-sigma position dispersion at the end of the hypersonic guidance phase (parachute deployment) below approximately 3 km. Four different propulsive terminal descent guidancealgorithms were examined. Of these four, a near propellant-optimal analytic guidance law showed promisefortheconceptualdesignofpinpointlandingvehicles.Theexistenceofapropellantoptimumwithregardto theinitiationtimeofthepropulsiveterminaldescentwasshowntoexistforvarious flightconditions.Subsonicguided parachutes were shown to provide marginal performance benefits, due to the timeline associated with descent through the thin Mars atmosphere. This investigation also demonstrates that navigation is a limiting technology for Mars pinpoint landing, with landed performance being largely driven by navigation sensor and map tie accuracy.

A study is presented detailing the simulation of a drag-free follow-up mission to NASA’s Gravity Recovery and Climate Experiment. This work evaluates controller performance, as well as thrust, power, and propellant mass requirements for drag-free spacecraft operation at orbital altitudes of 160–225 km. In addition, sensitivities to thermospheric wind, Global Positioning System signal accuracy, and availability of ephemeris data are studied. Thruster (control actuator) models are based on two different Hall thrusters for providing the orbital along-track acceleration, colloid thrusters for the normal acceleration, and a miniature xenon ion thruster for the cross-track acceleration. At an altitude of 160 km, the maximum along-track thrust component is calculated to be 98 mNwith a required dynamic (throttling) response of 42 mN=s. Themaximum position error at this altitude was shown to be in the along-track direction with a magnitude of 3314.9 nm. At 225 km, the maximum along-track thrust component reduces to 10.3 mN. The maximum dynamic response at this altitude is 4:0 mN=s. For the spacecraft point design considered with a propellant mass fraction of 0.18, the mission lifetime for the 160 km case was calculated to be 0.76 years. This increases 2.27 years at an altitude of 225 km.

The human exploration of Mars presents many challenges, not least of which is the task of entry, descent, and landing. Because human-class missions are expected to have landed payload masses on the order of 40 to 80 t, significant challenges arise beyond those of current robotic missions. This study uses parametric trade studies to provide insight into the feasibility of using Viking and Apollo heritage technologies to enable a human-class mission to Mars. The challenges encountered with human-class missions, as well as potential solutions, are highlighted through the results of parametric studies on vehicle size and mass. To populate the trade space, aerocapture, and entry-from-orbit analyses of 10 and 15-m diam aeroshells with a lift-to-drag ratio of 0.3 and 0.5 were investigated. The methodology developed to perform these trade studies represents a significant advancement in human Mars entry, descent, and landing system sizing. Numerous comparisons are made with past missions, both real and conceptual, and sources of discrepancies are discussed. Results indicate that in the limit, a crew capsule used only for entry from orbit could have an arrival mass as low as 20 t. For larger landed payloads, such as a 20-t surface power system, a vehicle with an arrival mass on the order of 80 t may be required. Finally, no feasible entry, descent, and landing systems were obtained with the capability to deliver more than approximately 25 t of landed payload to the Mars surface for arrivalmasses less than 100 t. This suggests that extension of traditional entry, descent, and landing technologies used for robotic exploration may be insufficient for human Mars exploration.

Earth-orbit-raising mission designs. However, although highly promising as a potential means of dispatching interplanetary payloads, a number of key factors are identified that can severely limit such potential. In particular, propellant boil-off in the chemical upper stage during orbit raising can lead to payload-optimal specific impulse values in many cases. Additional considerations such as the desirability of solar electric tug reuse can further contribute to marginalize the benefits of solar electric Earth orbit raising with respect to competing propulsion options. This study investigates key issues associated with solar electric Earth-orbit-raising missions, including circumstancesleadingtopayload-optimalspecificimpulsevalues,payload-optimalinclinationsfororbitraising,and achievable relative payload gain rates for solar electric Earth-orbit-raising missions compared with all-chemical injectionscenarios. Anewvariationofthelow-thrust rocketequationisderivedforinterplanetarymissions. Finally, a novel form of orbital transfer, referred to as an Earth-eclipse-evading transfer, is also introduced, which can potentially increase achievable solar electric Earth-orbit-raising payload ratios and rates when higher-inclination orbitsareaccessible.ItisconcludedthatsolarelectricEarthorbitraisingrepresentsapromisingwayofundertaking interplanetary missions, but that constraints nevertheless exist that can powerfully reduce the effectiveness of such mission designs.

In this paper a solar sail magnetotail mission concept was examined. The 43-m square solar sail is used to providethe required propulsion for continuous sun-synchronous apse-line precession. The main driver in this mission was found to be the reduction of launch mass and mission cost while enabling a nominal duration of 2 years within the framework of a demonstration mission. It was found that the mission concept provided an excellent solar sail technology demonstration option. The baseline science objectives and engineering goals were addressed, and mission analysis for solar sail, electric, and chemical propulsion performed. Detailed subsystems were defined for each propulsion system and it was found that the optimum propulsion system is solar sailing. A detailed tradeoff as to the effect of spacecraft and sail technology levels, and requirements, on sail size is presented for the first time. The effect of, for example, data acquisition rate and RF output power on sail size is presented, in which it is found that neither have a significant effect. The key sail technology requirements have been identified through a parametric analysis.

Future robotic and human missions to Mars require improved landed precision and increased payload mass. Low ballistic coefficient entry vehicles decelerate high in the thin Mars atmosphere and may be used to deliver higher-mass payloads to the surface. A high-altitude supersonic propulsive divert maneuver is proposed as a means of precision landing for low ballistic coefficient entry vehicles that decelerate to supersonic speeds at altitudes of 20–60 km. This divert maneuver compares favorably to traditional precision landing architectures with up to 100% improvement in range capability while saving over 30% in propellant mass. Through Monte Carlo simulations, it was found that architectures that use hypersonic vehicles with ballistic coefficients of 10 kg/m² can potentially land within 500 m of a target with this maneuver alone. This high-altitude divert range capability is sensitive to altitude and flight-path angle variations at maneuver initiation, and it is relatively insensitive to velocity at initiation. The propellant mass fraction is relatively invariant to the initial conditions and correlates directly with the divert distance.

An extension of the classical method by Alfano, for the analysis of optimal circle-to-circle two-dimensional orbit transfer, is presented for a deep space probe equipped with a solar electric primary propulsion system. The problem is formulated as a function of suitable design parameters, which allow the optimal transfer to be conveniently characterized in a parametric way, and an indirect approach is used to find the optimal steering law that minimizes the required propellant mass. The numerical results, obtained by solving a number of optimal control problems, are arranged into contour plots, characterized by different and well-defined behaviors depending on the value of the initial spacecraft propulsive acceleration, the final orbit radius, and the thruster’s specific impulse. The paper presents also a semi-analytical mathematical model for preliminary mission analysis purposes, which is shown to give excellent approximations of the (exact) numerical solutions when the number of revolutions of the spacecraft around the Sun is greater than five. An Earth–Mars cargo mission has been thoroughly investigated to validate the proposed approach. In this case, assuming a propulsion system with a specific impulse of 3000 s (comparable to that installed on the Deep Space 1 spacecraft), the results obtained with the semi-analytical model coincide, from an engineering point of view, with the numerical solutions both in terms of total mission time (about 8.3 years) and propellant mass fraction required (about 17.5%). By decreasing the value of the specific impulse, the differences between the results from the semi-analytical model and the numerical simulations tend to increase. However, good results are still possible if the number of revolutions of the spacecraft around the Sun is close to an integer number.

We compute optimal low-thrust transfers (with constant thrust and constant specific impulse) between Earth and Mars over a range of flight times (from 120 to 270 days) and launch years (between 2009 and 2022). Unlike impulsive transfers, the mass-optimal trajectory depends strongly on the thrust and specific impulse of the propulsion system. A low-thrust version of the rocket equation is provided where the initial mass or thrust may be minimized by varying the initial acceleration and specific impulse for a given power-system specific mass and for a trajectory time of flight. With fixed time-of-flight transfers there is a minimum thrust and a maximum allowable specific mass; that is, if the available thrust is too low or the specific mass is too large then the desired transfer does not exist. We find the minimum allowable thrust for constrained time-of-flight missions is on the order of a few Newtons per metric ton of payload for power systems with tens of kg/kW. As expected, the thrust and AV requirements of the trajectories decrease with increasing flight times. By extending the flight time from 180 to 270 days the A V is reduced by 40% for powered captures and up to 35% for aeroassisted capture trajectories.

Optimization of a spacecraft’s interplanetary trajectory and electric propulsion system remains a complex and difficult problem. Simultaneously solving for the optimal trajectory, power level, and exhaust velocity can be difficult and time consuming. If the power system’s technology level is unknown, multiple optimizations must be conducted to map out the trade space. Trajectories with constant-power, solar-power, variable-specific-impulse, and constant-specific-impulse low-thrust propulsion systems are analyzed and optimized. The technological variables, power system specific mass, propellant tank coefficient, structural coefficient, and the launch vehicle are integrated into the cost function allowing for maximization of the payload mass fraction. A classical solution is reviewed that allows trade studies to be conducted for constant-power, variable exhaust velocity systems. The analysis is expanded to include bounded-power constant specific impulse systems and solar electric propulsion spacecraft with constant and variable exhaust velocity engines. The cost function and mass fractions are dimensionless to allow for scaling of the spacecraft systems.

Mars ascent trajectories are calculated for small-scale vehicles that would improve the affordability of Mars sample return. Vehicle size, thrust levels, staging, and the importance of atmospheric drag are all taken into consideration. The high acceleration of conventional solid rockets requires a steep trajectory for drag avoidance, followed by a relatively large circularization burn, appropriate for a second stage. Lower thrust reduces total ∆v because reduced drag permits less steep trajectories that require small circularization burns. The results suggest the development of miniature liquid-propelled vehicles or advanced solid rockets having reduced thrust and multiple-burn capability.

Given the benefits of coupling low-thrust propulsion with gravity assists, techniques for easily identifying candidate trajectories would be extremely useful to mission designers. The computational implementation of an analytic, shape-based method for the design of low-thrust, gravity-assist trajectories is described. Two-body motion (central body and spacecraft) is assumed between the flybys, and the gravity-assists are modeled as discontinuities in velocity arising from an instantaneous turning of the spacecraft’s hyperbolic excess velocity vector with respect to the flyby body. The method is augmented by allowing coast arcs to be patched with thrust arcs on the transfers between bodies. The shape-based approach permits not only rapid, broad searches over the design space, but also provides initial estimates for use in trajectory optimization. Numerical examples computed with the shape-based method, using an exponential sinusoid shape, are presented for an Earth‐Mars‐Ceres rendezvous trajectory and an Earth‐Venus‐Earth‐Mars‐Jupiter flyby trajectory. Selected trajectories from the shape-based method are successfully used as initial estimates in an optimization program employing direct methods.

Missions such as Mariner 10, Voyager 1, Galileo, and Stardust all used gravity-assist flybys to achieve their mission goals efficiently. Methods to design such gravity-assist missions are fairly well developed and generally assume all major maneuvers are performed impulsively by chemical rockets. The recent success of the low-thrust Deep Space 1 mission demonstrates that low-thrust (high-efficiency) propulsion is ready to be used on future missions, potentially reducing the required propellant mass or the total time of flight. By combining both gravity-assist flybys and low-thrust propulsion, future missions could enjoy the benefits of both. To realize such missions, an effective design methodology is needed. A two-step approach to the design and optimization of low-thrust gravity-assist trajectories is described. The first step is a search through a broad range of potential trajectories. To speed up this search, a simplified shape-based trajectory model is used. The best trajectories are chosen using a heuristic cost function. The second step optimizes the most promising trajectories using an efficient parameter optimization. method. Examples of missions designed using this approach are presented, including voyages to Vesta, Tempel 1, Ceres, Jupiter, and Pluto.

A variety of reusable launch vehicle concepts has been designed and analyzed within the scope of the Future European Space Transportation Investigations Program systems study, which has been performed under a contract of the ESA by a joint European industry team. The designs under consideration included reusable two-stage-to-orbit configurations as well as several reusable single-stage-to-orbit vehicles with cryogenic rocket propulsion. All concepts were based on unified requirements, standards, technology assumptions, and design tools. To verify the consistency of the design process and to ensure the compatibility of the achieved results, as well as to identify the relative merits and inherent characteristics of the different basic vehicle concepts, a comparative technical assessment was performed among the various designs themselves as well as in relation to analogous configurations from other programs found in the open literature. Whereas the performance relations between the designs generated under common ground rules were found to be plausible, discrepancies in comparison to other programs could be traced to differences in requirements, technology assumptions, and design approaches. The results of associated considerations with respect to the main vehicle characteristics and related performance parameters are presented.

This paper investigates the use of passive electrodynamic tether drag as a method for quickly removing spent or dysfunctional spacecraft from low Earth orbits (LEO). The fundamental physical principles underlying the operation of an electrodynamic drag Terminator Tether TM are developed, some practical considerations are discussed, and calculationsof the area-time productare made forspacecraft orbits representative of those that will beused in the LEO satellite constellations of the next few decades. These calculations indicate that electrodynamic drag can remove a spacecraft from a typical 700 ‐2000-km LEO constellation orbit within a few months using a Terminator Tether system massing less than 3% of the spacecraft dry mass. Although the tether increases the cross-sectional area of the satellite system during the deorbit phase, the electrodynamic drag is so many times greater than atmospheric drag at these altitudes that the total area-time product can be reduced by several orders of magnitude, reducing the risks of collisions with other satellites. Concerns regarding tether survivability can be solved by using a multiline, fail-safe Hoytether TM construction. The Terminator Tether may thus provide a cost-effective method of mitigating the growth of debris in valuable constellation orbits.

The performance capabilities of an experimental Hall thruster were obtained experimentally for both variable and constant thrust modes at low power levels in order to enable orbit transfer under the ine uence of air drag with small satellites. For this purpose the measured thrusterperformance wasemployed in calculationsof transfer trajectories. As a result, by applying optimal thrust acceleration control the required propellant mass for a given low-Earth-orbit transfer was reduced by 15‐ 17% as compared to that required for spacecraft operation with a constant thrust to power ratio. For this purpose a Hall thruster with a movable anode is needed. In addition, the operation of an array of thrusters enables the increase of a factor of variations of the thrust acceleration during the e ight time.

I T IS a challenge to predict heat-shield response for space mission involving significant radiant heating. More data are needed to increase the knowledge in the high-temperature flight environment [1–3]. The discrepancies between the calculated and the measurement results led one to inquire into the details of the interaction of radiation, convection, and ablation.Most of the radiation is absorbed in the boundary layer, which contains the ablation product. The radiatively heated boundary layer raises the convective heat transfer rate partly at the point at which absorption occurs [4,5]. It is very difficult to study these phenomena in a ground-test facility because of the limitations on enthalpy and dimensions. A new flight test with an ablatingwall becomes highly desirable.We explore the possibility of a superorbital reentry flight experiment to test such interactions. A series of flight tests [6,7] are proving the usefulness of the Russian submarine-launched Volna [8] as an efficient means of studying the reentry problems. However, it produces only suborbital entry velocities. We propose that the reentry velocity is increased by fitting the payload of Volna with a small solid rocket booster (SRB). An inviscid Newtonian analysis is made to evaluate the aerodynamic coefficients. The mass and performance are estimated using data on existing small SRB through flight trajectory calculations. The shape of the reentry vehicle (RV) and entry velocity is taken to be that of the Fire-II [4] so that comparison can be made. The interaction between radiative, convective heating, and ablation is calculated at the stagnation point for a carbon-phenolic (CP) ablator. II. Analysis Method

Propellant inefficiency material in particulate form is characterized in a laboratory pulsed plasma thruster (PPT) operating at 1 Hz with a 204 discharge energy (20 W). Exhaust deposits are collected and analyzed using a combination of a scanning electron microscope with energy dispersive x-ray analysis and microscopic imaging. Teflon(trademark) particulates are observed with characteristic dianietens ranging from over 100 micrometers down to less than 1 micrometer.

˜A new simplie ed approach is developed to analyze the performance of alternative propulsion systems with an additional power source, such as augmented catalytic, nuclear, and electrical propulsion systems. The optimum performance characteristics are calculated for missions with a e xed time of e ight, constant thrust, or constant mass e ow rate, considering the effects of propellant density, tankage fraction, and chemical enthalpy of propellants. A sensitivity analysis is employed to determine the ine uence of different parameters on mass efe ciency.