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In this paper, the authors investigate whether airbreathing engines have useful application to vertical takeoff and vertical landing systems, which currently represent the mainstream for reusable l...

Purpose The capability to predict and evaluate various configurations’ performance during the conceptual design phase using multidisciplinary design analysis and optimization can significantly increase the preliminary design process’s efficiency and reduce design and development costs. This research paper aims to perform multidisciplinary design and optimization for an expendable microsatellite launch vehicle (MSLV) comprising three solid-propellant stages, capable of delivering micro-payloads in the low earth orbit. The methodology’s primary purpose is to increase the conceptual and preliminary design process’s efficiency by reducing both the design and development costs. Design/methodology/approach Multidiscipline feasible architecture is applied for the multidisciplinary design and optimization of an expendable MSLV at the conceptual level to accommodate interdisciplinary interactions during the optimization process. The multidisciplinary design and optimization framework developed and implemented in this research effort encompasses coupled analysis disciplines of vehicle geometry, mass calculations, aerodynamics, propulsion and trajectory. Nineteen design variables were selected to optimize expendable MSLV to launch a 100 kg satellite at an altitude of 600 km in the low earth orbit. Modern heuristic optimization methods such as genetic algorithm (GA), particle swarm optimization (PSO) and SA are applied and compared to obtain the optimal configurations. The initial population is created by passing the upper and lower bounds of design variables to the optimizer. The optimizer then searches for the best possible combination of design variables to obtain the objective function while satisfying the constraints. Findings All of the applied heuristic methods were able to optimize the design problem. Optimized design variables from these methods lie within the lower and upper bounds. This research successfully achieves the desired altitude and final injection velocity while satisfying all the constraints. In this research effort, multiple runs of heuristic algorithms reduce the fundamental stochastic error. Research limitations/implications The use of multiple heuristics optimization methods such as GA, PSO and SA in the conceptual design phase owing to the exclusivity of their search approach provides a unique opportunity for exploration of the feasible design space and helps in obtaining alternative configurations capable of meeting the mission objectives, which is not possible when using any of the single optimization algorithm. Practical implications The optimized configurations can be further used as baseline configurations in the microsatellite launch missions’ conceptual and preliminary design phases. Originality/value Satellite launch vehicle design and optimization is a complex multidisciplinary problem, and it is dealt with effectively in the multidisciplinary design and optimization domain. It integrates several interlinked disciplines and gives the optimum result that satisfies these disciplines’ requirements. This research effort provides the multidisciplinary design and optimization-based simulation framework to predict and evaluate various expendable satellite launch vehicle configurations’ performance. This framework significantly increases the conceptual and preliminary design process’s efficiency by reducing design and development costs.

A retardation effect, which is the decrease of the feed pressure and the thrust of an autophage rocket, is considered. The effect is caused by inertial forces acting on the rocket sliding engine an...

International audience; Reusability of the first stage of launch vehicles may offer new perspectives to lower the cost of payload injection into orbit if sufficient reliability and efficient refurbishment can be achieved. One possible option that may be explored is to design the vehicle first stage for both reusable and expendable uses, in order to increase the flexibility and adaptability to different target missions. This paper proposes a multilevel multidisciplinary design optimization (MDO) approach to design aerospace vehicles addressing multimission problems. The proposed approach is focused on the design of a family of launchers for different missions sharing commonalities using multi-objective MDO to account for the computational cost associated with the discipline simulations. The multimission problem addressed considers two missions: 1) a reusable configuration for a sun synchronous orbit with a medium payload range and recovery of the first stage using a gliding-back strategy; 2) an expendable configuration for a medium payload injected into a geostationary transfer orbit. A dedicated MDO formulation introducing couplings between the missions is proposed in order to efficiently solve such a coupled problem while limiting the number of calls to the exact multidisciplinary analysis thanks to the use of Gaussian processes and multi-objective efficient global optimization.

Uncertainty-based design optimization has been widely acknowledged as an advanced methodology to address competing objectives of aerospace vehicle design, such as reliability and robustness. Despite the usefulness of uncertainty-based design optimization, the computational burden associated with uncertainty propagation and analysis process still remains a major challenge of this field of study. The metamodeling is known as the most promising methodology for significantly reducing the computational cost of the uncertainty propagation process. On the other hand, the nonlinearity of the uncertainty-based design optimization problem's design space with multiple local optima reduces the accuracy and efficiency of the metamodels prediction. In this article, a novel metamodel management strategy, which controls the evolution during the optimization process, is proposed to alleviate these difficulties. For this purpose, a combination of improved Latin hypercube sampling and artificial neural networks are involved. The proposed strategy assesses the created metamodel accuracy and decides when a metamodel needs to be replaced with the real model. The metamodeling and metamodel management strategy are conspired to propose an augmented strategy for robust design optimization problems. The proposed strategy is applied to the multiobjective robust design optimization of an expendable launch vehicle. Finally, based on non-dominated sorting genetic algorithm-II, a compromise between optimality and robustness is illustrated through the Pareto frontier. Results illustrate that the proposed strategy could improve the computational efficiency, accuracy, and globality of optimizer convergence in uncertainty-based design optimization problems.

The Ina Irregular Mare Patch (IMP) in the central lunar nearside (18.7°N, 5.3°E) is a small, two-by-three-kilometer exposure of uncommon volcanic deposits. The unusually well-preserved deposits occurring at Ina suggest that they were recently emplaced and that the Moon experienced small-scale, ongoing volcanic eruptions during the last 100 million years. The existence of young volcanism on the Moon, if confirmed, would indeed challenge our current notions of how the Moon is structured and how it evolved, an interpretation still vigorously debated. Because of the possibility for unique volcanic materials, Ina is a high-priority destination for future study, and determination of relatively recent volcanism at Ina and the other IMPs requires ground-truth. We assessed the geologic setting, scientific rationale, and several potential landing sites at Ina. We also developed a SmallSat-class lander design and mission architecture, along with a corresponding science payload, to address the top science objectives at Ina. We named the mission concept IMPEL (Irregular Mare Patch Exploration Lander). The primary objective of the IMPEL mission is to determine the presence and abundance of any sub-meter-scale fractures or pits in Ina's deposits that would inform their origins and age. Secondary objectives include characterizing Ina's composition and physical properties. The spacecraft configuration (named the Dual ESPA Module Planetary Lander, DEMPL) that we developed in this study employs two ESPA (Evolved Expendable Launch Vehicle, EELV, Secondary Payload Adapter) modules integrated with flexible tethers and bands. One of the modules is used to decelerate the spacecraft from direct lunar insertion and the other functions both as additional propulsion and as a detached lander containing most of the spacecraft subsystems and the science payload. Our novel spacecraft configuration provides a new mechanism for surface exploration in deep space using small-scale planetary landers. The lander is an adaptable platform that can deliver up to 9 kg of science payload to the lunar surface. This approach is very capable for deep space exploration at the small spacecraft scale, and it has the potential to provide access to the lunar surface using a low-cost, low-mass, affordable, simplified configuration. The Ina IMP is one of many possible destinations.

The application of small satellites is a new focus on deep space exploration. Small satellites are easy to launch as secondary payloads, and more suitable for international cooperation. However, small satellites have several major weaknesses for deep space missions, such as the lack of orbit transfer capability and large antenna. This paper presents an innovative concept of small satellite bus based on Evolved Expendable Launch Vehicle (EELV) Secondary Payload Adapter (ESPA). The proposed satellite bus is equipped with orbit control subsystem and a large parabolic antenna for the exploration of Jovian and Saturnian systems. In this paper, the mission requirements, the launch feasibilities and the design parameters of the proposed satellite bus are also discussed.

Reduced scale and improved responsiveness will be the technical and economic drivers of future satellite systems. Based on decades of practical experience with rocket-only expendable launch vehicle...

This paper presents an analysis of the scenario of small satellites and its correspondent launch vehicles. The miniaturization of electronics, together with reliability and performance increase as well as reduction of cost, have allowed the use of commercials-off-the-shelf in the space industry, fostering the Smallsat use. An analysis of the launched Smallsats during the last 20 years is accomplished and the main factors for the Smallsat (r)evolution, outlined. Based on historic data, future scenarios for different mass categories of Smallsats are presented. An analysis of current and future launch vehicles reveals that we are currently in a phase of transition, where old launch vehicles get retired and new ones enter the market. However, the satellite launch vehicle business has been established to carry payloads of thousands of kilos into low Earth orbit and has not adjusted itself to the market of Smallsats. As a result, there is only 1 launch vehicle for dedicated Smallsat launches commercially available, but it carries a high price tag. Several small low-cost launch vehicles under development are identified and the challenges to overcome, discussed. Since these small launch vehicles have similar complexity as huge launch vehicles, high development costs are intrinsic, leading to a high specific price (USD/kg payload).

The practice of launching multiple payloads on a single space launch vehicle is becoming increasingly popular, with the mean number of payloads per launch increasing from 1.45 payloads per launch in 2000 to 1.84 payloads per launch in 2013. A best-fit descending algorithm was used to reassign existing payloads to launch vehicles with the goal of reducing launch vehicle usage and wastage of payload capacity. Assigning the existing set of geosynchronous payloads to a minimum number of a single existing launch vehicle can reduce wastage to as low as 2% in some cases for a single type of launch vehicle, compared with a current wastage of 15%. An extension of this technique to a scenario with multiple types of available launch vehicles with minimizing total cost as an objective shows that savings of as much as 45% in cost per payload mass delivered to geosynchronous orbit are possible by rearranging current payloads and changing usage of current launch vehicles.

Marman clamps are the most commonly used attachments for spacecraft to expendable launch vehicles. Accurate prediction of the load capacity of the marman clamps is required for spacecraft programs....

The Bi-sat Observations of the Lunar Atmosphere above Swirls (BOLAS) is a NASA planetary CubeSat mission concept in low lunar orbit. The BOLAS lower CubeSat is at a 90 km altitude above the lunar surface during spiraling down from the Evolved Expendable Launch Vehicle (EELV) Secondary Payload Adapter (ESPA) to the Moon. Without phase change material (PCM), the worst hot case temperature prediction for the Command and Data Handling (C&DH) exceeds the 61°C maximum operating limit, and those for the Iris solid state power amplifier (SSPA) and transponder exceed the 50°C maximum operating limit. Miniature n-Tricosane PCM packs on the Iris SSPA and transponder, and miniature n-Hexacosane PCM packs on the C&DH are used to store thermal energy in sunlight and release it in the eclipse. With paraffin PCM, all the temperatures are within the maximum operating limits.

Fractional calculus is an interdisciplinary area with multifarious applications especially in control systems due to more flexible parameters to adjust the dynamic behaviour of all the physical systems. This paper presents a fractional order controller structure for an expendable launch vehicle during ascent phase. The launch vehicle has highly flexible and unstable dynamic model, and in addition to that unpredictable control issues are caused by the sloshing of liquid propellants and the inertia problems of the engines. Under these circumstances controller with large robustness margins are required to meet deviations in model parameters which are unknown before the actual flight. The rigid body dynamics of the vehicle is considered for designing fractional order controller to overcome all these control issues. The robustness of fractional order controller is compared with existing classical controller for 20% variations in the aerodynamic coefficients. The stability margins of the open loop transfer function with all plant dynamics shows the supremacy of the proposed controller over the classical controller. The control inputs to launch vehicles, in the form of attitude commands, are affected by the wind. Hence the designed fractional controller simulated under the wind disturbance input and the result shows disturbance rejection capability of controller.

This paper proposes an optimal staging method for reusable launch vehicles considering velocity losses. The proposed method expands a traditional stage optimization approach for expendable launch vehicles by introducing the descent phase of the separated stage and combining it with the trajectory optimization procedure. For a reusable stage, the structural ratios responsible for its ascent and landing phases are defined to formulate the optimal sizing problem and optimized along with variables for traditional staging problems. An optimal staging case study on a reusable launch vehicle demonstrates the effectiveness of the proposed method.

Data for space launches and launch vehicles from 2000 to 2013 are analyzed to rate the relative capability and cost-effectiveness of launch vehicles for missions. The masses of mission payloads are used to determine the cost of payload mass delivered to low Earth and geosynchronous transfer orbits. These costs are higher due to wastage of launch vehicle capacity totaling 20.4% for the studied missions. A minimum cost scenario is then studied, where the lowest cost launch vehicle capable of performing a mission is used. In this scenario, the total price over the studied period for launches with complete data would be reduced from $44.2B to $33.4B, a savings of 24.5%. Usage is then compared with this scenario, and it is found that the cheapest launch vehicle capable of performing a mission is only used for 20.4% of studied missions.

The Indian space program is a spacefaring success story with demonstrated capability in the design and building of application and scientific satellites, and the means to launch them into desired orbits. The end-to-end mission planning and execution capability comes with a high emphasis on self-reliance. Sounding rockets and small satellite launch vehicles provided the initial experience base for India. This experience was consolidated and applied to realize larger satellite launch vehicles. While many of the launch vehicle technologies were indigenously developed, the foreign acquisition of liquid propulsion technologies did help in catalyzing the development efforts. In this case, launch vehicle concept studies showed the inevitability of using a cryogenic upper stage for geosynchronous Earth orbit missions, which proved to be difficult technically and encountered substantial delays, given the geopolitical situation. However, launch capability matured from development to operational phases, and ...

As the Reusable Launch Vehicle can reduce the launching cost, and improve the ability of Operationally Responsive Space (ORS), many aerospace powers in the world consider the Reusable Launch Vehicle as a main development tendency of space transportation system. Recently, Space Exploration Technologies Corporation (SpaceX), as the spokesman of private aerospace corporations, has attracted all the world’s attention to the Reusable Launch Vehicle. SpaceX has been developing technologies for rockets’ fully and rapid reusability, and several recovery tests have been conducted on technology-demonstrators and post-mission controlled-descent tests on Falcon 9 rockets’ first stages, both touchdown on the ocean platform and the land. On 22 December 2015, a Falcon 9 FT rocket of SpaceX, carrying 11 Orbcomm communications satellites, lifted off from Cape Canaveral. After cutoff and stage-separation, the Falcon 9s first stage flight back into the atmosphere and pulled off a powered landing on on Landing Zone 1, which is about 10 km far away from the Launch Site of SLC-40, settling to a smooth tail-first touchdown, making it a significant space “first”. After already celebrated a successful booster landing, SpaceX had decided to attempt a landing at sea on this flight despite available margins for a return to land. This decision was prompted by the need to master the landing sequence for sea-based recoveries which will be needed for about half of Falcon 9’s flights when lifting heavy satellites to high-energy orbits. On 8 April 2016, SpaceX’s Falcon 9 FT rocket, carrying CRS-8 Dragon cargo, lifted off from Cape Canaveral. The Falcon 9’s first stage has accomplished the first ever successful returning rocket on the Autonomous Spaceport Drone Ship (ASDS). This is considered a landmark accomplishment on the road to economical interplanetary and asteroid-to- planet space travel because it enables expensive launch vehicles to be reused. In this paper, based on the experiments of Falcon 9, and analyzing the main projects of the Reusable Launch Vehicle abroad in the past 60 years, the critical technologies will be researched and the results will provide references for the research of new type of space transportation systems.

The possible directions of development of space launch technology, including space launch to suborbital trajectory, in order to reduce the specific cost of launch at the expense of the majority of reusable carrier components, are analyzed. Opportunities of providing reusability for horizontal and vertical launch are compared. The experience of soft landing of the first stage of Falcon 9-R rocket of US firm SpaceX is taken into account. The requirements to the air breathing engine, which could provide an economical horizontal launch, are considered. The requirements for the engine could be reduced for suborbital launch, and in this simplified case they could be fulfilled well already at the present stage of development of aerospace technologies. The proposed principles of reusable HTHL system "WIG-craft +Aerospace Plane" are described and its advantages over the vertical launch systems are considered.

The article gives a retrospective review and comparison of the implemented and non-implemented projects of super-heavy launch vehicles in our country and in the United States. The basic features of the design-layouts are defined, and efficient ways of further development of super-heavy launch vehicles in Russia are offered.

As 1987 began, with the Reagan administration starting its final two years in the White House, much of the optimism about the ability to put the United States on a new course in space had dissipated. The Challenger accident and the long and contentious process of developing a launch recovery strategy had diverted administration attention from the possibility of setting long-range space objectives. Reagan’s signature space initiative, the space station, was in budget trouble, and, with the grounding of the shuttles, there was little progress in space commercialization. During 1987–1988, the interagency rivalries that had emerged over the roles of the shuttle and expendable launch vehicles shifted their focus to the NASA space station and a potential commercial alternative, the Industrial Space Facility. Having survived its budget problems and after a redesign, the space station was given a name: Freedom.

In addition to applying to the space sector Ronald Reagan’s conviction that the free market was the preferred path to future progress, there were three other influences supporting Reagan administration initiatives toward space commercialization during the 1983–1984 period. One was the emergence, just as the Reagan administration arrived in Washington, of a variety of private sector actors interested in space commercialization. Second, there was also a sudden flowering of new, entrepreneurial commercial space ventures. A third influential factor leading to the increased emphasis on commercialization was high enthusiasm regarding the potential economic payoffs from various commercial space activities. While the revenue from privatizing remote sensing and commercializing expendable launch vehicles was expected to be relatively modest, payoffs from new commercial activities carried out in Earth orbit were forecast to be in the multiple billions of dollars. These influences combined as the foundation for crafting a Reagan administration policy toward encouraging and facilitating commercial activities in orbit.

The NASA Space Network (SN) Operations began with the launch of the first Tracking and Data Relay Satellite (TDRS-1) on April 4, 1983 with on-orbit operations conducted from the White Sands Ground Terminal (WSGT) in Las Cruces, New Mexico. Over the past 35 years, the SN has evolved to its current configuration comprised of four strategically located satellite ground stations and a constellation of ten geosynchronous Tracking and Data Relay Satellites that provide customer support 24 hours a day, 365 days per year, at an unprecedented 99.9% proficiency. The Space Network provides total global Communications, Data Relay, and Tracking services for Low Earth Orbiting (LEO) satellites, Human Space Flight, Expendable Launch Vehicles (ELV) and Scientific missions. This paper will address the Operations Management of the NASA Space Network and the complexities associated with this responsibility, as well as significant mission support highlights.

Launch vehicle dynamics modeling, simulation, and trajectory optimization within a single modeling tool is a challenging task due to the highly interconnected disciplines involved such as propulsion, aerodynamics, structures, mechanisms, and GNC, amongst others. In particular, changing environmental conditions and perturbations have to be considered throughout the ascent of expendable launch vehicles (ELV) as well as in the more complex scenario of the ascent and descent of reusable launch vehicles (RLV). Both the multidisciplinary design approach and the vehicle's mission definition can have considerable consequences for the overall modeling and optimization strategy. Therefore, a standardized modeling tool able to meet design requirements for a broad range of mission scenarios from vertical takeoff vertical landing (VTVL) to winged horizontal takeoff horizontal landing (HTHL) configurations is needed. Dedicated developments of multidisciplinary frameworks for launch vehicle modeling and preliminary design optimization have been presented in the relevant literature and also developed in the industry. It is common that the modeling of launch vehicles is performed by several independent, discipline-specific tools. With such an approach, only a limited amount of interactions of the involved disciplines with the overall system dynamics can be accounted for. Therefore, we propose a new multidisciplinary modeling framework considering all relevant effects on the system dynamics of launch vehicles using the object-oriented, equation-based, multi-physical, and acausal modeling language MODELICA. By capitalizing MODELICA's modeling capabilities, the framework enables the object-oriented and mathematically efficient modeling of subsystems and components related to most of the key disciplines of a launcher system. Another objective of this paper is to present a subset of the modeling framework for expendable and reusable launch vehicles regarding Functional Mock-up Units (FMU) and to demonstrate the advantages and capabilities of such a modeling approach within a combined trajectory optimization of the ascent and descent phases of launch vehicles. The modeling framework is shown for a standardized three degrees of freedom (3-DOF) model, covering the kinematics and dynamics formulation, environmental effects, aerodynamics, and propulsion models for system dynamics and subsequent trajectory simulations. The 3-DOF launch vehicle model is integrated as an FMU into the Trajectory Optimization Package trajOpt of DLR-SR's multi-objective optimization tool MOPS. The benefits of our modeling framework are discussed in terms of future rigid and flexible multibody modeling capabilities as well as GNC design and trade-off studies.

Recent technology advancements in microelectric propulsion will enable the next generation of small spacecraft to perform trajectory and attitude maneuvers with significant ΔV requirements, provide thrust over long mission durations, and replace reaction wheels for attitude control. These advancements will open up the class of mission architectures achievable by small spacecraft to include formation flying, proximity operations, and precision pointing missions in both low Earth orbit and interplanetary destinations. The goal of this study is to establish the optimal performance parameters for future microelectric propulsion technology that are applicable to a broad range of flight demonstration platforms (for example, dedicated 3- to 12-unit CubeSats to evolved expendable launch vehicle secondary payload adaptor-class spacecraft) for a variety of applications, including low Earth orbit and Earth escape orbit transfers, travel to interplanetary destinations, hover and drag makeup missions, and performing r...

Reusable launch vehicle (RLV) is a goal pursued by different aerospace agencies, world-wide. Indian space research organization is also one among them. This paper covers the challenges faced and the solutions developed for control law design of a typical re-entry vehicle. The first stage of the test vehicle, used for technology demonstration mission, is a conventional solid rocket motor. The second stage of the test vehicle is a winged body vehicle termed as technology demonstrator vehicle (TDV). This TDV forms the orbiter stage of a conceptual two-stage-to-orbit vehicle. While developing control law for TDV, the experience gained in developing control law for expendable launch vehicles is fully utilized. Due to the presence of winged-second-stage, the solution for attitude control problem naturally inherits the procedure and technology used for designing control law of an aircraft. There are significant differences a re-entry vehicle has in comparison to both aircraft and launch vehicles. Due to these differences, the final control law depends on both aircraft and launch vehicle control law design methodologies. In this paper the authors share the insights gained in control law design of a RLV. The paper gives the reasons for the selection of a combination of co-ordinate systems, for the development of linear plant model of RLV. It emphasizes the difference in the plant modelling with respect to aircraft dynamics in accounting gravitational force in the linear plant model. Requirement of pseudo trim condition for force balance is described in the paper. The paper highlights the commands to be tracked and variables to be fed back during different regimes of re-entry flight. Issues in control law gain design with respect to selected sensors and their placement are elaborated in the paper. Insights on the aspect of trim scheduling, gain scheduling, and finalization of control law structure are also given in the paper.

Human missions to Mars, particularly to the Martian surface, are grand endeavors that place extensive demands on ground infrastructure, launch capabilities, and mission systems. The interplay of capabilities and limitations among these areas can have significant impacts on the costs and ability to conduct Mars missions and campaigns. From a mission and campaign perspective, decisions that affect element designs, including those based on launch vehicle and ground considerations, can create effects that ripple through all phases of the mission and have significant impact on the overall campaign. These effects result in impacts to element designs and performance, launch and surface manifesting, and mission operations. In current Evolvable Mars Campaign concepts, the NASA Space Launch System (SLS) is the primary launch vehicle for delivering crew and payloads to cis-lunar space. SLS is currently developing an 8.4m diameter cargo fairing, with a planned upgrade to a 10m diameter fairing in the future. Fairing diameter is a driving factor that impacts many aspects of system design, vehicle performance, and operational concepts. It creates a ripple effect that influences all aspects of a Mars mission, including: element designs, grounds operations, launch vehicle design, payload packaging on the lander, launch vehicle adapter design to meet structural launch requirements, control and thermal protection during entry and descent at Mars, landing stability, and surface operations. Analyses have been performed in each of these areas to assess and, where possible, quantify the impacts of fairing diameter selection on all aspects of a Mars mission. Several potential impacts of launch fairing diameter selection are identified in each of these areas, along with changes to system designs that result. Solutions for addressing these impacts generally result in increased systems mass and propellant needs, which can further exacerbate packaging and flight challenges. This paper presents the results of the analyses performed, the potential changes to mission architectures and campaigns that result, and the general trends that are more broadly applicable to any element design or mission planning for human exploration.

The National Aeronautics and Space Administration (NASA) Goddard's Rideshare Office estimates that between 2013 and 2022, NASA launches of primary satellites will have left unused more than 20,371 kilograms of excess capacity. This equates to hundreds of millions of dollars in launch-vehicle costs going unutilized. To fill this void with a standard CubeSat or SmallSat spacecraft platform, which when required to be more reliable, still will cost in the neighborhood of $1M a kilogram, making it prohibitively expensive. A newly proposed solution, which NASA is pursuing, is called the Capsulation Satellite or CapSat. CapSat is a modularized, pressurized, thermally controlled spacecraft designed to host ruggedized commercially available instrumentation in a terrestriallike environment on orbit. Using a technique that is under review for a patent, CapSat actively manages internal air temperatures in a manner similar to a household thermostat. This gives CapSat high-thermal stability, which, in turn, provides component longevity. CapSat was specifically designed to take advantage of the United States Air Force (USAF) Rideshare Program and the Evolved Expendable Launch Vehicle Secondary Payload Adaptor, or ESPA ring. The ESPA ring comes in two sizes: standard and Grande. CapSat primarily will take advantage of the ESPA Grande to provide a 300-kilogram payload capability per attachment point, with up to four attachment points per ring. This approach combines a high-mass capability with a proven Rideshare mechanical interface and secondary payload management infrastructure. Opportunities for ESPA-based co-manifests are continuing to expand. The CapSat program is currently funded to design and build a limited prototype and perform thermal-vacuum testing. CapSat is currently in the concept/study phase for both single missions and constellation of earth- and space-observing missions. One of these studies includes land imaging using state-of-the-art advanced infrared detector technology. This paper will report on the current status of the CapSat hardware design, testing, and results as well as any openly available advanced concept study results. The CapSat solution is intended to be a game-changing paradigm shift. CapSat will repurpose currently available, already-proven technology to reduce spaceflight hardware costs to less than $50,000 per kilogram.

Lapan, Indonesian National Institute of Aeronautics and Space is responsible for the development of aeronautics and space technology. One of the milestones is the currently orbiting LAPAN-TUBsat, put into orbit with the help of other nation launch systems. So, LAPAN still depends on other countries to launch its satellite. Currently, a program on a conventional land base rocket delivery system is under development. As an alternative to the ongoing program a simpler and cost effective launch system will be studied. One of the alternatives is an aerial launch system. This system is seen as a viable alternative due to its operational and support simplicity. A study of this system will be carried out by comparing it to the conventional launch system. The comparisons are limited to the vehicle of the system and its ability to launch a micro satellite size object. As an alternative to the conventional system, the aerial launch is within reach of the current needs and capabilities of LAPAN. As the study suggests, this alternative may be one of the solutions LAPAN needs for a launch system.

In May of 2011, NASA selected the Origins, Spectral Interpretation, Resource Identification, and Security-Regolith Explorer (OSIRIS-REx) asteroid sample return mission as the third mission in the New Frontiers program. The other two New Frontiers missions are New Horizons, which explored Pluto during a flyby in July 2015 and is on its way for a flyby of Kuiper Belt object 2014 MU69 on Jan. 1, 2019, and Juno, an orbiting mission that is studying the origin, evolution, and internal structure of Jupiter. The spacecraft departed for near-Earth asteroid (101955) Bennu aboard an United Launch Alliance Atlas V 411 evolved expendable launch vehicle at 7:05 p.m. EDT on September 8, 2016, on a seven-year journey to return samples from Bennu. The spacecraft is on an outbound-cruise trajectory that will result in a rendezvous with Bennu in August 2018. The science instruments on the spacecraft will survey Bennu to measure its physical, geological, and chemical properties, and the team will use these data to select a site on the surface to collect at least 60 g of asteroid regolith. The team will also analyze the remote-sensing data to perform a detailed study of the sample site for context, assess Bennus resource potential, refine estimates of its impact probability with Earth, and provide ground-truth data for the extensive astronomical data set collected on this asteroid. The spacecraft will leave Bennu in 2021 and return the sample to the Utah Test and Training Range (UTTR) on September 24, 2023., Comment: 89 pages, 39 figures, submitted to Space Science Reviews - OSIRIS-REx special issue

Evolution, current status, and key future trends in the development of space launchers and launch services, including space elevators and other new technologies; Analysis of the adequacy of the current governance system applicable to space launchers and launch services, and recommendations for actions to improve relevant global governance; Specific and appropriate recommendations for global governance actions that should be taken to ameliorate current and future conditions, along with the organizations where positive action might be taken to ensure the improvement of these services and to sustain the use of space for peaceful purposes and for the benefit of all humankind.

The term ‘spaceport’ encompasses a broad variety of facilities (Fig. 5.1). Some spaceports are used to launch large rockets into Earth orbit, some of which carry their payloads into deep space while others carry probes and yet others ferry cargo and/or astronauts. Then there are spaceports that have been developed for launching jet-like spacecraft from runways for trips into suborbital space. And at the most basic level there are the spaceports that simply support short flights with small rockets. But one factor common to all spaceports is that each is the location where space launch vehicles and their payloads are prepared and subsequently launched. As of October 2016 there are eight commercially licensed spaceports in the United States, but the regulations that governed the application process (14 CFR Part 420) were developed at a time when most launch operators were in the business of launching orbital launch vehicles that operated from federal launch ranges. But since those regulations came into effect the launch vehicle arena has changed significantly. Perhaps the most notable change has been with the arrival of suborbital reusable launch vehicles (sRLV) and the development of private launch sites such as Blue Origin’s remote location in Van Horn, Texas. Since spaceport design is driven largely by the types of launch vehicles being launched, it is useful to know what capabilities these vehicles have and the facilities required to support their launch and recovery. To that end, what follows is a synopsis of some of the current crop of orbital and suborbital vehicles around the world.

Purpose: This paper explains how to obtain the Bayes estimates of the whole launch vehicle and of a vehicle stage, respectively, for a newly developed expendable launch vehicle. Methods: We determine the parameters of the beta prior distribution using the upper bound of the 60% Clopper-Pearson confidence interval of failure probability which is calculated from previous launch data con-sidering the experience of the developer. Results: Probability that a launch vehicle developed from an inexperienced developer succeeds in the first launch is obtained by about one third, which is much smaller than that estimated from the previous research. Conclusion: The proposed approach provides a more conservative estimate than the previous noninformative prior, which is more reasonable especially for the initial reliability of a new vehicle which is developed by an inexperienced developer.Key Words: Clopper-Pearson Confidence Interval, Beta Posterior Distribution, Stage-Based Reliability

International audience; A 3-D A* mission planning/replanning algorithm satisfying both environmental constraints as static and mobile obstacles and kinematic and dynamic constraints coming from the aircraft limitations is proposed. This algorithm carries out the planning of the shortest path selected from the interconnections of several waypoints generated in the mission region. The shortest path is identified according to the presence of various obstacles during the path search, and its objective is to reach different goal points. Analysis of simulation results of this improved 3-D A *algorithm shows that it is adapted to different environments where more flight constraints can be considered, and it proposes an optimal solution to generate a free obstacle path. New trajectory plans can be generated from updated environmental data. In this specific launch mission for an autonomous airplane, the environmental uncertainties are important parameters.

NASA has created a plan to implement the Flexible Path strategy, which utilizes a heavy lift launch vehicle to deliver crew and cargo to orbit. In this plan, NASA would develop much of the transportation architecture (launch vehicle, crew capsule, and in-space propulsion), leaving the other in-space elements open to commercial and international partnerships. This paper presents a space exploration strategy that reverses that philosophy, where commercial and international launch vehicles provide launch services. Utilizing a propellant depot to aggregate propellant on orbit, smaller launch vehicles are capable of delivering all of the mass necessary for space exploration. This strategy has benefits to the architecture in terms of cost, schedule, and reliability.

In this paper, Reliability-Based Multidisciplinary Design Optimization (RBMDO) of a two-stage solid propellant expendable launch vehicle (LV) is investigated. Propulsion, weight, aerodynamics (geometry) and trajectory (performance) disciplines are used in an appropriate combination. Throw weight minimization is chosen as objective function. Design variables for system level optimization are selected from propulsion, geometry and trajectory disciplines. Mission constraints contain the final velocity, the height above ground, and flight path angle. The constraints that appear during the flight are also considered. Assuming a normal distribution for the uncertain variables, Latin Hypercube Sampling (LHS) method selects the sample values for simulation runs which are eventually utilized for calculating probability density function of constraints and their reliability at each design point. Sequential Quadratic Programming (SQP) technique is used to achieve the optimal solution. Although the launch vehicle throw weight is increased negligibly in comparison with deterministic optimization, results show that the reliability-based method satisfied desired reliability of the constraints.

Requirements of a propulsion system for lunar-exploration missions using a launch vehicle, the development of which has recently started in Korea, were conceptualized. A new three-stage-to-orbit launch system, the Korea Space Launch Vehicle 2 will be a middle-class launch vehicle in that its expected payload capacity to low Earth orbit will be about 2.6 tons. The Korea Space Launch Vehicle 2 is under development and may possibly launch a lunar orbiter and a lander within 10–15 years. Considering some of the limitations and requirements a new space-propulsion system must be developed, and it has been found that a H2O2/kerosene-bipropellant rocket is the optimum propulsion system among several types of chemical rockets. The optimum thrust and burn-time requirements for orbital transfers and landing using H2O2/kerosene-bipropellant rockets were also derived by mass budget design and three-degree-of-freedom orbit-trajectory calculations. It has been found that, using the Korea Space Launch Vehicle 2, 1800 N-c...

Expendable Launch Vehicles are designed with a single mission perspective. Reusable Launch Vehicle design aims for reusability and maintainability for multiple missions. Power system of such missions has features that are distinctive from operational launch vehicles like PSLV & GSLV. The instrumentation, Navigation Guidance & Control, pyro and actuation system power are derived from Li ion batteries, which is a notable feature of Reusable Launch Vehicle- Technology Demonstrator. RLV-TD power system consists of a unified redundant power bus running throughout the vehicle unlike conventional launch vehicles. This paper discusses the configuration of electrical power distribution system of RLV-TD highlighting the constraints and design considerations. Load requirements and power characteristics are defined. Techniques adopted for fault protection and redundancy to achieve reliability are discussed. Battery capacity estimation method followed for continuous and pulse current requirements are described.

Significant and substantial progress continues to be accomplished in the design, development, and testing of the Space Launch System (SLS), the most powerful human-rated launch vehicle the United States has ever undertaken. Designed to support human missions into deep space, SLS is one of three programs being managed by the National Aeronautics and Space Administration's (NASA's) Exploration Systems Development directorate. The Orion spacecraft program is developing a new crew vehicle that will support human missions beyond low Earth orbit, and the Ground Systems Development and Operations program is transforming Kennedy Space Center into next-generation spaceport capable of supporting not only SLS but also multiple commercial users. Together, these systems will support human exploration missions into the proving ground of cislunar space and ultimately to Mars. SLS will deliver a near-term heavy-lift capability for the nation with its 70 metric ton (t) Block 1 configuration, and will then evolve to an ultimate capability of 130 t. The SLS program marked a major milestone with the successful completion of the Critical Design Review in which detailed designs were reviewed and subsequently approved for proceeding with full-scale production. This marks the first time an exploration class vehicle has passed that major milestone since the Saturn V vehicle launched astronauts in the 1960s during the Apollo program. Each element of the vehicle now has flight hardware in production in support of the initial flight of the SLS -- Exploration Mission-1 (EM-1), an un-crewed mission to orbit the moon and return. Encompassing hardware qualification, structural testing to validate hardware compliance and analytical modeling, progress in on track to meet the initial targeted launch date in 2018. In Utah and Mississippi, booster and engine testing are verifying upgrades made to proven shuttle hardware. At Michoud Assembly Facility in Louisiana, the world's largest spacecraft welding tool is producing tanks for the SLS core stage. This paper will particularly focus on work taking place at Marshall Space Flight Center (MSFC) and United Launch Alliance in Alabama, where upper stage and adapter elements of the vehicle are being constructed and tested. Providing the Orion crew capsule/launch vehicle interface and in-space propulsion via a cryogenic upper stage, the Spacecraft/Payload Integration and Evolution (SPIE) Element serves a key role in achieving SLS goals and objectives. The SPIE element marked a major milestone in 2014 with the first flight of original SLS hardware, the Orion Stage Adapter (OSA) which was used on Exploration Flight Test-1 with a design that will be used again on EM-1. Construction is already underway on the EM-1 Interim Cryogenic Propulsion Stage (ICPS), an in-space stage derived from the Delta Cryogenic Second Stage. Manufacture of the Orion Stage Adapter and the Launch Vehicle Stage Adapter is set to begin at the Friction Stir Facility located at MSFC while structural test articles are either completed (OSA) or nearing completion (Launch Vehicle Stage Adapter). An overview is provided of the launch vehicle capabilities, with a specific focus on SPIE Element qualification/testing progress, as well as efforts to provide access to deep space regions currently not available to the science community through a secondary payload capability utilizing CubeSat-class satellites.

For decades, economic reuse of launch vehicles has been an elusive goal. Recent attempts at demonstrating elements of launch vehicle recovery for reuse have invigorated a debate over the merits of different approaches. The parameter most often used to assess the cost of access to space is dollars-per-kilogram to orbit. When comparing reusable vs. expendable launch vehicles, that ratio has been shown to be most sensitive to the performance lost as a result of enabling the reusability. This paper will briefly review the historical background and results of recent attempts to recover launch vehicle assets for reuse. The business case for reuse will be reviewed, with emphasis on the performance expended to recover those assets, and the practicality of the most ambitious reuse concept, namely propulsive return to the launch site. In 2015, United Launch Alliance (ULA) announced its Sensible, Modular, Autonomous Return Technology (SMART) reuse plan for recovery of the booster module for its new Vulcan launch vehicle. That plan employs a non-propulsive approach where atmospheric entry, descent and landing (EDL) technologies are utilized. Elements of such a system have a wide variety of applications, from recovery of launch vehicle elements in suborbital trajectories all the way to human space exploration. This paper will include an update on ULA's booster module recovery approach, which relies on Hypersonic Inflatable Aerodynamic Decelerator (HIAD) and Mid-Air Retrieval (MAR) technologies, including its concept of operations (ConOps). The HIAD design, as well as parafoil staging and MAR concepts, will be discussed. Recent HIAD development activities and near term plans including scalability, next generation materials for the inflatable structure and heat shield, and gas generator inflation systems will be provided. MAR topics will include the ConOps for recovery, helicopter selection and staging, and the state of the art of parachute recovery systems using large parafoils for space asset recovery and high altitude deployment. The next proposed HIAD flight demonstration is called HULA (for HIAD on ULA), and will feature a 6m diameter HIAD. An update for the HULA concept will be provided in this paper. As proposed, this demonstration will fly as a secondary payload on an Atlas mission. The Centaur upper stage provides the reentry pointing, deorbit burn, and entry vehicle spin up. The flight test will culminate with a recovery of the HIAD using MAR. HULA will provide data from a Low Earth Orbit (LEO) return aeroheating environment that enables predictive model correlation and refinement. The resultant reduction in performance uncertainties should lead to design efficiencies that are increasingly significant at larger scales. Relevance to human scale Mars EDL using a HIAD will also be presented, and the applicability of the data generated from both HULA and SMART Vulcan flights, and its value for NASA's human exploration efforts will be discussed. A summary and conclusion will follow.

For the first time in almost 40 years, a NASA human-rated launch vehicle has completed its Critical Design Review (CDR). By reaching this milestone, NASA's Space Launch System (SLS) and Orion spacecraft are on the path to launch a new era of deep space exploration. NASA is making investments to expand science and exploration capability of the SLS by developing the capability to deploy small satellites during the trans-lunar phase of the mission trajectory. Exploration Mission 1 (EM-1), currently planned for launch no earlier than July 2018, will be the first mission to carry such payloads on the SLS. The EM-1 launch will include thirteen 6U Cubesat small satellites that will be deployed beyond low earth orbit. By providing an earth-escape trajectory, opportunities are created for advancement of small satellite subsystems, including deep space communications and in-space propulsion. This SLS capability also creates low-cost options for addressing existing Agency strategic knowledge gaps and affordable science missions. A new approach to payload integration and mission assurance is needed to ensure safety of the vehicle, while also maintaining reasonable costs for the small payload developer teams. SLS EM-1 will provide the framework and serve as a test flight, not only for vehicle systems, but also payload accommodations, ground processing, and on-orbit operations. Through developing the requirements and integration processes for EM-1, NASA is outlining the framework for the evolved configuration of secondary payloads on SLS Block upgrades. The lessons learned from the EM-1 mission will be applied to processes and products developed for future block upgrades. In the heavy-lift configuration of SLS, payload accommodations will increase for secondary opportunities including small satellites larger than the traditional Cubesat class payload. The payload mission concept of operations, proposed payload capacity of SLS, and the payload requirements for launch and deployment will be described to provide potential payload users an understanding of this unique exploration capability.