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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.

Aerosol sprays are used in many fields because they are easy to use and low in cost. Because flash-boiling has a great influence on atomization performance and spray characteristics such as droplet diameter and spray cone angle of aerosol spray, it is necessary to adjust the flash-boiling by changing the internal design of the aerosol vessel to obtain the desired atomization performance. Therefore, spray structure formed on the outside of nozzle during both transient and steady state. with various nozzle design and internal liquid/propellant composition ratio were observed in this study. The spray development process was visualized using a high-speed camera and Nd:Yag laser was used to form a laser sheet and it illuminates the internal structure of spray at steady state. In addition, the diameter and velocity of spray droplets were measured by PDPA equipment, which uses a noninvasive method. Selected design parameters were the number of stem orifice and the diameter of under-tap of the valve and vapor phase tap, and the mass fraction of propellant inside the aerosol vessel. As the diameter of the under tap increased, the ratio of the high-density area increased. However, as the difference between the area of the under-tap orifice and the total area of the stem orifice became smaller, the spray cross section area became larger and the shape of low-density region became closer to the concentric circle. In addition, finer droplets were generated in case with large stem orifice & smaller under-tap. The Vapor Phase Tap and propellant mass fraction have the greatest influence on the atomization characteristics of all design parameters because they change the gas phase ratio inside the jet. However, higher propellant mass fraction makes more uniform droplet diameter distribution than increased vapor phase tap diameter.

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...

Speculation on near-term scientific reasons for the exploration of lunar pits is offered alongside comments on possible longer-term human exploitation. It is proposed that in order to determine whether or not one or more of the pits offer access the large subsurface voids e.g. a non-collapsed lava tube, a preliminary reconnaissance mission solely focused on obtaining lateral images (and/or LiDAR maps) is needed. Possible concept options for such a preliminary reconnaissance mission are discussed. It is suggested that one of the best possible strategies is to employ a micro-sized probe (~0.3 m) that would hop from a nearby main landing spacecraft to the selected pit. After the surface position of the main lander is determined accurately, the probe would perform a ballistic hop, or hover-traverse, a distance of ~3 km over the lunar surface using existing propulsive and guidance technology capability. Once hovering above the pit, the probe or a separate tethered imaging unit would then be lowered into the pit to acquire the necessary subsurface void topology data. This data would then be transmitted back to Earth, directly, via the lander, or via a store-and-forward orbiting relay. Preliminary estimates indicate that a probe of ~14 kg (dry mass) is viable using a conventional hydrazine monopropellant system with a propellant mass fraction of less than ~0.2 (20%) including margins, suggesting a piggyback architecture would be feasible.

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.

Purpose The purpose of this paper is to deal with the study of an innovative unmanned mission to Mars, which is aimed at acquiring a great amount of detailed data related to both Mars’ atmosphere and surface. Design/methodology/approach The Mars surface exploration is conceived by means of a fleet of drones flying among a set of reference points (acting also as entry capsules and charging stations) on the surface. The three key enabling technologies of the proposed mission are the use of small satellites (used in constellation with a minimum of three), the use of electric propulsion systems for the interplanetary transfer (to reduce the propellant mass fraction) and lightweight, efficient, drones designed to operate in the harsh Mars environment and with its tiny atmosphere. Findings The low-thrust Earth-Mars transfer is designed by means of an optimization approach resulting in a duration of slightly more than 27 months with a propellant amount of about 125 kg, which is compatible with the choice of considering a 500 kg-class spacecraft. Four candidate drone configurations have been selected as the result of a sensitivity analysis. Flight endurance, weight and drone size have been considered as the driving design parameters for the selection of the final configuration, which is characterized by six rotors, a total mass of about 6.5 kg and a flight endurance of 28 minutes. In the mission scenario proposed, the drone is assumed to be delivered on the Mars surface by means of a passive entry capsule, which acts also as a docking station and charging base. Such a capsule has been sized both in terms of mass (68 kg) and power (80 W), showing to be compatible with 500 kg-class spacecraft. Research limitations/implications As a general conclusion, the study shows the mission concept feasibility. Practical implications The concept would return incomparable scientific data and can be also be potentially implemented with a relatively low budget exploiting of the shelf components to the larger extent, small identical spacecraft buses and modular low-cost drones. Originality/value The innovative mission architecture proposed in this study aims at providing a complete coverage of the surface and lowest atmospheric layers. The main innovation factor of the proposed mission consists in the adoption of small multi-copter UAVs, also called “drones,” as remote-sensing platforms.

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 describes an architecture for human exploration of Mars that would be enabled by 10 metric tons (t) per day of ISRU water extracted at the lunar poles. The architecture involves a concept for a standardized “self-propelled tank” having an 85% propellant mass fraction that could hold up to 51 t of LH 2 /LO 2 propellant. A variant of this tank would be equipped with landing legs so that, when fully-loaded, it could lift another full tank into Low Lunar Orbit (LLO), release it, and have enough propellant remaining to mate with an empty tank and to land safely back at the ISRU plant in a lunar polar crater. Two such pairs would be launched each lunar sidereal month such that the two payload tanks could mate together in LLO and depart to a carefully-selected High Earth Orbit (HEO). This orbit is such that exactly one tank would be consumed propelling one full tank to HEO and returning one empty tank from HEO to LLO, with both tanks arriving empty. Also, a Low Earth Orbit (LEO) is selected where two full tanks in HEO could mate together such that one full tank arrives in LEO while the other tank has just enough propellant left to return to HEO, arriving empty. This combination of capabilities allows payloads to be launched from Earth into LEO, where they could be ferried to HEO using lunar propellant. Such payloads could be assembled together as stacks in HEO, where an array of full tanks accumulated over time from lunar ISRU could propel them on Hohmann transfers to Mars on every opportunity (every 26 months). The notional payload we envision launching to Mars every 26 months includes two 31 t Deep Space Habitats (DSHs, one for the outbound journey and one for the return journey) plus a 48 t Mars Surface Habitat with Ascent Vehicle (MSHAV). The stack departs HEO with a small burn from the array of 24 standardized propellant tanks — just enough to put it on an Oberth Maneuver trajectory, swinging by the top of the Earth's atmosphere where the main Trans-Mars Injection (TMI) burn occurs, consuming 7.8 of the 24 tanks. No tanks are jettisoned, so as to make this architecture sustainable and affordable. Upon arrival at Mars ∼7–9 months later, another 9.2 tanks are burned to inject the stack into Low Mars Orbit (LMO). Exactly one tank is burned to de-orbit the MSHAV, reducing its velocity to zero at some point deep in the Mars atmosphere. The one tank that de-orbits the Mars entry stack is lost. After ∼500 days on the Mars surface, the MAV launches and mates with the main stack in LMO. The returning DSH and the 23 tanks depart LMO with a burn of 4.3 tanks for Trans-Earth Injection. Upon arrival at Earth, 1.7 tanks are burned to inject into HEO, where the tanks could be shuttled back to the lunar surface for re-fueling. The DSH is mated with a single full tank at HEO that brings it to LEO to rendezvous with a crew capsule for re-entry.

In any manned mission architecture, upwards of seventy percent of all payload delivered to orbit is propellant, and propellant mass fraction dominates almost all transportation segments of any mission requiring a heavy lift launch system like the Saturn V. To mitigate this, the use of an orbital propellant depot has been extensively studied. In this paper, a thermal model of an orbital propellant depot is used to examine the effects of passive and active thermal management strategies. Results show that an all passive thermal management strategy results in significant boil-off for both hydrogen and oxygen. At current launch vehicle prices, these boil-offs equate to millions of dollars lost per month. Zero boil-off of propellant is achievable with the use of active cryocoolers; however, the cooling power required to produce zero-boil-off is an order of magnitude higher than current state-of-the-art cryocoolers. This study shows a zero-boil-off cryocooler minimum power requirement of 80–100 W at 80 K for liquid oxygen, and 100–120 W at 20 K for liquid hydrogen for a representative Near-Earth Object mission. Research and development effort is required to improve the state-of-the-arts in-space cryogenic thermal management.

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 experimental pressure-apparent burning rate model was established based on the recorded pressure-time data in the closed vessel to study the actual burning gas generate rate of a propellant whose form function is not available. The new burning model can replace the Vieille’s law and form function of propellants in the interior ballistic equations to predict the interior ballistic performance of gun. As distinct from the traditional geometric form function based burning rate model, the new model introduced the concept of relative pressure impulse related to the actual burnt web thickness of propellant statistically. The apparent burning rate was expressed by the propellant mass fraction burnt vs. relative pressure impulse curve. The interior ballistic simulation of selected gun propellants was conducted using the new burning rate model. The results predicted by the new model show a good agreement with the classic interior ballistic predication results though the agreement with the experimental results is still to be improved.

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.

In this work, we consider multiobjective space mission design problems. We will start from the need, from a practical point of view, to consider in addition to the (Pareto) optimal solutions also nearly optimal ones. In fact, extending the set of solutions for a given mission to those nearly optimal significantly increases the number of options for the decision maker and gives a measure of the size of the launch windows corresponding to each optimal solution, i.e., a measure of its robustness. Whereas the possible loss of such approximate solutions compared to optimal—and possibly even ‘better’—ones is dispensable. For this, we will examine several typical problems in space trajectory design—a biimpulsive transfer from the Earth to the asteroid Apophis and two low-thrust multigravity assist transfers—and demonstrate the possible benefit of the novel approach. Further, we will present a multiobjective evolutionary algorithm which is designed for this purpose.

In order to provide an improved understanding of the flow in pressurized-metered dose inhalers (pMDIs), especially monitoring the output temperature and mass flow rate to obtain maximum atomization efficiency from the available energy, a numerical model for a two phases, multi-component compressible flow in a pressurized-metered dose inhaler is presented and validated. It is suitable for testing with various formulations and different geometries for a range of pMDI devices. We validated the model against available data in the literature for a single component HFA 134a propellant, and then investigated the response of the model to other formulations containing non-volatile components. Further validation is obtained by an experiment using the dual beam method which acquired the actuation flow properties such as spray velocity and duration. The deviation of the numerical predictions for the peak exit velocity against the experimental results is 5.3% and that for effective spray duration 5.0%. From the numerical and experimental results, it is found that for the formulations with the mass fraction of HFA 134a > 80%, the effective spray duration of the pMDI is around 0.1 s. Furthermore the droplet peak exit velocity at the axial station x = 25 mm from the actuation nozzle decreases from 20 to 15 m/s with the reduction of the propellant (HFA 134a) from 95%. Formulations with the mass fraction of HFA 134a below 80% produce poor quality spray which is indicated from the unsteady peak exit velocity, changeable spray number density in each experimental test, and numerical simulations also confirmed the non-viability of this condition.

The fundamental relationship between entropy and aerospace vehicle and mission performance is analyzed in terms of the general availability rate balance between force-based vehicle performance, available energy associated with expended propellant, and the overall loss rate of availability, including the vehicle wake. The availability relationship for a vehicle is analytically combined with the vehicle equations of motion; this combination yields the balance between on-board energy rate usage and rates of changes in kinetic and potential energies of the vehicle and overall rate of entropy production. This result is then integrated over time for a general aerospace mission; as examples, simplified single-stage-to-orbit rocket-powered and air-breathing missions are analyzed. Examination of rate of availability loss for the general case of an accelerating, climbing aerospace vehicle provides a powerful loss superposition principle in terms of the separate evaluation and combination of loss rates for the same vehicle in cruise, acceleration, and climb. Rate of availability losses is also examined in terms of separable losses associated with the propulsion system and external aerodynamics. These loss terms are cast in terms of conventional parameters such as drag coefficient and engine specific impulse. Finally, rate losses in availability for classes of vehicles are described.

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.

Single-stage-to-orbit (SSTO) rocket technology offers the potential to substantially reduce launch costs, but is not considered practical for conventional launch vehicles. However, new research in composite propellant tank technology opens the field for reevaluation. One technology that increases the efficiency and feasibility of SSTO flight is an altitude compensating rocket engine nozzle, as opposed to a conventional constant-area, bell nozzle. By implementing an altitude compensation nozzle, such as an aerospike nozzle for inatmosphere flight, the propellant mass fraction (PMF) may be reduced by as much as two percent compared to a conventional rocket engine. In this paper, Optimal Trajectories by Implicit Simulation (OTIS) is used to model SSTO flight trajectories by comparing an aerospike nozzle to a conventional bell nozzle; this includes vacuum thrust, vacuum specific impulse, and nozzle combinations to show that nozzle variability increases the efficiency of SSTO flight through a reduction in PMF. Results suggest that limiting the altitude compensation to lower altitudes further decreases PMF by 0.1-0.2%. Uncertainty estimates suggest that this added benefit is within the range of uncertainty of OTIS, and performs at least as well as the theoretical, infinite nozzle expansion. It is apparent that altitude compensating nozzles outperform conventional bell nozzle designs.