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This study presents a novel forward dynamics model of racing kayak surge motion. The model, driven by adjustable oscillating propulsion forces, replicates true paddle stroke forces. It is based on Newton's second law of motion, with the net force comprising of propulsion force, skin friction drag, wave-making drag, form drag, and air drag. The total mass includes the kayak, paddler, and added mass. The fourth-order Runge–Kutta method is used to solve for speed and distance. The model incorporates a unique hull formulation for wetted surface area estimation, Michell's integral for wave-making drag, and a log wind profile with allometric scaling for air drag. It accounts for varying stature height, body mass, kayak dimensions, wind speed, and stroke propulsion profiles. The model is a significant advancement in its domain, with broad applicability for the future. It aligns with existing data and analyses, but validation data is limited. [ABSTRACT FROM AUTHOR]

Abstract Efficient and reliable space transportation systems are fundamental to the future success of routine scientific, commercial and strategic space missions. RBCC (rocket-based combined cycle) comprising rocket, ramjet and scramjet engines is a propulsion technology that offers promise for reusable space launch systems owing to advantages over traditional propulsion in various aspects including efficiency and flexibility. However, the viability of RBCC-powered access-to-space requires careful consideration and its assessment represents a challenge to conventional design approaches due to the highly complex and coupled characteristics of the system associated with multi-mode propulsion and multi-stage launch. This study has been undertaken to examine the performance of a conceptual RBCC-based TSTO (two-stage-to-orbit) system and identify the key requirements and design factors to achieve space launch via this system. Multi-objective design optimization has been conducted with respect to important design criteria by means of evolutionary algorithms assisted by surrogate modeling and trajectory optimization. The influence of RBCC engine characteristics particularly in terms of performance and operation as well as the scaling of the vehicle and propulsion on the overall performance of the TSTO system and its feasibility has been quantified and examined. Furthermore, a comparative study with a rocket-only-based TSTO system has also been conducted, verifying the advantages of the RBCC-based TSTO system. The results highlight complex aerodynamic characteristics for both concepts as well as highly nonlinear propulsion characteristics for the RBCC-based concept. Highlights • Characteristics of access-to-space using RBCC propulsion have been investigated. • Multi-objective design optimization has been conducted for RBCC-based TSTO. • Trajectory simulation is coupled into surrogate-assisted evolutionary algorithms. • Unique advantages and feasibility of RBCC-based TSTO have been discussed. • Key design factors and areas for technological improvements have been identified. [ABSTRACT FROM AUTHOR]

Purpose This paper deals with the estimation of the necessary masses of propulsion components for multirotor UAS. Originally, within the design process of multirotors, this is an iterative problem, as the propulsion masses contribute to the total takeoff mass. Hence, they influence themselves and cannot be directly calculated. The paper aims to estimate the needed propulsion masses with respect to the requested thrust because of payload, airframe weight and drag forces and with respect to the requested flight time.Design/methodology/approach Analogue to the well-established design synthesis of airplanes, statistical data of existing electrical motors, propellers and rechargeable batteries are evaluated and analyzed. Applying Rankine and Froude's momentum theory and a generic model for electro motor efficiency factors on the statistical performance data provides correlations between requested performance and, therefore, needed propulsion masses. These correlations are evaluated and analyzed in the scope of buoyant-vertical-thrusted hybrid UAS.Findings This paper provides a generic mathematical propulsion model. For given payloads, airframe structure weights and a requested flight time, appropriate motor, propeller and battery masses can be modelled that will provide appropriate thrust to lift payload, airframe and the propulsion unit itself over a requested flight time.Research limitations/implications The model takes into account a number of motors of four and is valid for the category of nano and small UAS.Practical implications The presented propulsion model enables a full numerical design process for vertical thrusted UAS. Hence, it is the precondition for design optimization and more efficient UAS.Originality/value The propulsion model is unique and it is valid for pure multirotor as well as for hybrid UAS too. [ABSTRACT FROM AUTHOR]

The article presents a discussion of the National Aeronautics and Space Administration's (NASA) proposal of a reusable space shuttle system. This space shuttle system has been actively promoted by NASA with an aim of reducing the space transportation costs for a variety of space programs projected for the 1980s and beyond. The author examines many of the issues that surround the proposed shuttle, some of which include its role in national priorities, a lack of a clear definition of the shuttle's use, uncertainties surrounding the cost and reuse of its solid-fueled boosters and the overall ongoing costs of the system.

Abstract The performance of the inlet is crucial to the cruise flight of a hypersonic air-breathing propulsion vehicle. The objective of this work is to investigate the uncertainty and sensitivity of pressure field and the performance parameters for a hypersonic inlet due to the uncertainty of five flow parameters, including freestream Mach number, Reynolds number, angle of attack, temperature and wall temperature. The steady Reynolds Averaged Navier-Stokes equations are solved to predict the inlet start and unstart flows within the hysteresis loop. Then, a point-collocation non-intrusive polynomial chaos method (NIPC) is utilized to quantify the uncertainty and sensitivity in the output quantities of interest. The uncertainty analysis in pressure field shows that Mach number and angle of attack of freestream make dominant contributions to the total uncertainty, and the Mach number has remarkable impacts in the isolator. In the start flow, the angle of attack exerts its prominent influence in the post-shock regions, while Mach number mainly dominates these regions ahead of and around the shocks. The reason may be interpreted as the much greater pressure derivatives with respect to angle of attack in the post-shock regions. Significant discrepancies are presented for the unstart flow. The reflected shock waves in the unstart flow are less sensitive to the variations of flow parameters. The external flow field, separation bubble and reflected shocks are significantly affected by angle of attack. Besides, the uncertainties of the performance parameters in the start flow are about twice those in the unstart flow. The sensitivity analysis further reveals that Mach number is the major contributor to the total uncertainty of performance parameters. The correlation coefficients via linear regression method clearly illustrate the relationships between the five input parameters and the performance parameters. Highlights • Effects of five flow parameters on the aerodynamics is investigated. • The uncertainty in pressure field and the performance parameters are quantified. • The critical input parameters are identified and the explanation is provided. • Differences between the start and unstart flow are compared and discussed. [ABSTRACT FROM AUTHOR]

Abstract The Space Propulsion Laboratory (SPL) of the Massachusetts Institute of Technology (MIT) is developing the Ion Electrospray Propulsion System (iEPS), designed to address a current need in CubeSat technology: miniaturized electric thrusters. These could be used for different applications, ranging from attitude control to interplanetary flights. In this work, performed together with the Space Systems Laboratory of the Costa Rica Institute of Technology (SETEC Lab), we explore a case study in which the iEPS is used for constellation management in Low Earth Orbit (LEO) when integrated in a 3U CubeSat. We analyze how a 180° separation in the Right Ascension of the Ascending Node (RAAN) between two CubeSats (SatA and SatB) starting in the same orbit can be achieved by modifying one of the spacecraft's orbital altitude, resulting in a difference in their rate of nodal precession (defined as the drift rate) due to the J 2 effect, and therefore a difference in their relative RAAN. The method consists of SatB increasing its semi-major axis, drifting in a higher orbit with a lower drift rate, and returning to the original semi-major axis once the desired difference in RAAN in achieved relative to the other spacecraft. SatA will stay in its original orbit, using its thruster to compensate for orbital energy loss due to atmospheric drag, therefore demonstrating another application of iEPS for constellation management. Three different simulations were studied, defined as the minimum time trajectory, minimum propellant trajectory and a hybrid trajectory, consisting of reaching a higher altitude orbit, but actively changing the RAAN using the propulsion system instead of drifting. It was observed that the difference in this orbital element could be achieved using 85 g of propellant in as little as 164 days for the minimum time trajectory. The same difference could also be achieved using only 44 g of propellant in 245 days for the minimum propellant trajectory. Furthermore, the results of the hybrid trajectory showed that the goal could be achieved in 161 days, but using 158 g of propellant mass, demonstrating the benefit of using a drift orbit. The results proved the feasibility of implementing iEPS for constellation management using 3U CubeSats in LEO. Highlights • A fuel-efficient method for CubeSat constellation management is presented. • The method generates a difference in the rate of nodal precession of spacecraft. • A description of the Ionic Liquid Electrospray Propulsion (iEPS) is presented. • The management of two CubeSats with iEPS, deployed in the same orbit, is analyzed. • A 180° difference in their ascending nodes is achievable when using the method. [ABSTRACT FROM AUTHOR]

The Lorentz force acting on an electrostatically charged spacecraft provides a new means of propulsion for orbital maneuvers such as Lorentz-augmented spacecraft formation flying. Modeling the Earth's magnetic field as a tilted dipole corotating with Earth, a nonlinear dynamical model that describes the orbital motion of a Lorentz spacecraft about an arbitrary elliptic orbit is developed. Then, the optimal open-loop control trajectories of formation establishment are solved by the pseudospectral method. To guarantee trajectory tracking in the presence of external perturbations and system uncertainties, a closed-loop neural network-based adaptive second order fast terminal sliding mode controller is designed using state feedback, which simultaneously solves the singularity and chattering problems that generally exist in conventional terminal sliding mode. Neural networks are employed to approximate the unknown nonlinearities in the system dynamics. Meanwhile, to ensure closed-loop tracking control without velocity measurements, an output feedback controller is also proposed with an observer introduced to capture the velocity signals. The overall stabilities for both control schemes are proved by a Lyapunov-based method. Numerical simulations are presented to verify the performance of the proposed controllers. Furthermore, the controllers could be applied to other Lorentz-augmented relative orbital control problems. [ABSTRACT FROM AUTHOR]

As a promising candidate for future air-breathing systems, the viability and efficiency of scramjet propulsion is challenged by a variety of factors including the combustion oscillation in scramjet combustor. A series of comparative experiments focusing on the combustion oscillation issue has been carried out in the present work. The obtained experimental results show that as the global equivalence ratio increases, the combustion oscillation becomes more regular and frequent which is the most intensive in the vicinity of the fuel jet and the periodic combustion oscillation is more possible when the injectors and flame-holding cavity are mounted on the expansion-side wall. In order to avoid the combustion oscillation in scramjet combustor, distributed injection scheme is an effective method which can induce two parts interacting stable flame. In addition, the results reveal that the varying fuel including hydrogen, ethylene and kerosene with different chemical kinetics has a significant effect on the reaction process in scramjet combustor, which can result in stable combustion, periodic oscillation and failed ignition respectively on the same operating condition of this paper. We believe that the present work is helpful to the designing of scramjet propulsion device. [ABSTRACT FROM AUTHOR]

Nuclear Thermal Rocket (NTR) propulsion is a viable and meritorious option for human exploration into deep-space because of its high thrust, improved specific impulse, well established technology, bimodal capability, and enhanced mission safety and reliability. The NTR technology has already been investigated and tested by the United States of America and Russia and the former Soviet Union. The representative Nuclear Engine for Rocket Vehicle Applications (NERVA) type reactors traditionally used Highly Enriched Uranium (HEU) fuels, shaped in hexagonal fuel element geometries because of the importance of making a high power reactor with a minimum size. Although the HEU-NTR designs are the best choice in terms of rocket performance and technical maturity, they inevitably provoke nuclear proliferation obstacles not only for all research and development activities by civilians and non-nuclear weapon states but also for potential commercialization. To overcome the security issues due to HEU, the non-proliferative, small-size NTR engine with low thrust levels of 41 kN–53 kN (9.2 klb f ∼ 11.9 klb f ), Korea Advanced NUclear Thermal Engine Rocket utilizing a Low-Enriched Uranium fuel (KANUTER-LEU), is being designed for future generations. Its design goals are to make use of an LEU fuel for its fairly compact core, but to minimize the rocket performance sacrifice relative to the traditional HEU-NTRs. To achieve these goals, a new space propulsion reactor is conceptually designed with the key concepts of a high uranium density fuel with resistance against high heating and H 2 corrosion, a thermal neutron spectrum core, and a compact and integrated fuel element core design with protective cooling capability. In addition, a preliminary design study of neutronics and thermal-hydraulics was performed to explore the design space of the new LEU-NTR reactor concept. The result indicates that the innovative reactor concept has great potential, both to implement the use of an LEU fuel and to create comparable rocket performance, compared to the existing HEU-NTR designs. [ABSTRACT FROM AUTHOR]

Space propulsion systems play an increasingly important role in planning of space missions. The traditional method for design of space propulsion systems includes numerous design loops, which does not guarantee to reach the best optimal solution. Multidisciplinary Design Optimization (MDO) is an approach for the design of complex systems that considers a design environment with multiple disciplines. The aims of this study are to implement and compare Multidisciplinary Feasible and Collaborative Optimization architectures for the multi-Objective optimization of a bi-propellant space propulsion system design. Several disciplines such as thrust chamber, cooling, and structure were exploited in a proper combination. The main optimization objectives in the MDO frameworks were to minimize the total wet mass and maximize the total impulse by considering several constraints. Furthermore, Genetic Algorithm and Sequential Quadratic Programming are employed as the system-level and local-level optimizers. The presented design methodology provides an interesting decision making approach to select the best system parameters of space propulsion systems under conflicting goals. [ABSTRACT FROM AUTHOR]

Recent studies have concluded that Viking-era entry system deceleration technologies are extremely difficult to scale for progressively larger payloads (tens of metric tons) required for human Mars exploration. Supersonic retropropulsion is one of a few developing technologies that may enable future human-scale Mars entry systems. However, in order to be considered as a viable technology for future missions, supersonic retropropulsion will require significant maturation beyond its current state. This paper proposes major milestones for advancing the component technologies of supersonic retropropulsion such that it can be reliably used on Mars technology demonstration missions to land larger payloads than are currently possible using Viking-based systems. The development roadmap includes technology gates that are achieved through ground-based testing and high-fidelity analysis, culminating with subscale flight testing in Earth's atmosphere that demonstrates stable and controlled flight. The component technologies requiring advancement include large engines (100s of kilonewtons of thrust) capable of throttling and gimbaling, entry vehicle aerodynamics and aerothermodynamics modeling, entry vehicle stability and control methods, reference vehicle systems engineering and analyses, and high-fidelity models for entry trajectory simulations. Finally, a notional schedule is proposed for advancing the technology from suborbital free-flight tests at Earth through larger and more complex system-level technology demonstrations and precursor missions at Mars. [ABSTRACT FROM AUTHOR]

The paper presents the 50th anniversary celebration of the Yuri Gagarin's pioneering space flight and highlights the technology that made it. It was Gagarin's mission that triggered the second phase of the space race-that of landing a man on the Moon-and the raft of technologies: from propulsion system to guidance computers-that would make it possible. [ABSTRACT FROM AUTHOR]

Abstract: This paper describes a development study of a precooled-cycle hypersonic turbojet engine for the first stage of TSTO space plane and hypersonic airplane. With reflecting the key technologies accumulated from ATREX (expander cycle ATR engine) ground tests, the next flyable subscale engine “S-engine” is now developed. S-engine has of rectangular cross-section, 2.2m of the overall length and about 100kg of the weight employing a variable-geometry rectangular inlet and nozzle. It produces 1.2kN of thrust at SLS, which corresponds to of the ATREX engine. Design of the hypersonic components such as the inlet, precooler and nozzle has been finished and their aerodynamic performances were verified by wind tunnel tests and CFD analyses. A prototype model of the diagonal-flow compressor whose pressure ratio is 6 was manufactured. Its rotating tests under the very-low pressure conditions are now in progress. The reverse-flow annular combustion chamber was successfully tested. The first flight test of the S-engine is to be conducted in 2008 by the balloon-based operation vehicle (BOV) which is about 5m in length, 0.55m in diameter and 500kg in weight. The vehicle is dropped from an altitude of 40km by a high altitude balloon. After 40-s free-fall, the vehicle pulls up and S-engine operates for 30s at about Mach 2. High altitude tests of the engine components corresponding to the BOV''s flight condition have been conducted. [Copyright &y& Elsevier]

Abstract: The research and development of the JAXA 150-mN ion engine are in progress. The objective of this activity is to establish the ion propulsion technology for future applications, which require high thrust levels, such as orbit raising and deep space exploration. Current efforts are directed to improvement in endurance of the thruster, power conditioner development, and test facility improvement because sufficiently high thruster performance has already been achieved. An endurance test of a main hollow cathode is being conducted. Back sputtering characteristics of a test facility were improved for a coming thruster endurance test. Wide range operation of the thruster, neutralizer characterization tests, and basic study on power processing units were also conducted. [Copyright &y& Elsevier]

Abstract: Development of MEMS-based (micro electro mechanical system) components and subsystems for space applications has been pursued by various research groups and organizations around the world for at least two decades. The main driver for developing MEMS-based components for space is the miniaturization that can be achieved. Miniaturization can not only save orders of magnitude in mass and volume of individual components, but it can also allow increased redundancy, and enable novel spacecraft designs and mission scenarios. However, the commercial breakthrough of MEMS has not occurred within the space business as it has within other branches such as the IT/telecom or automotive industries, or as it has in biotech or life science applications. A main explanation to this is the highly conservative attitude to new technology within the space community. This conservatism is in many senses motivated by a very low risk acceptance in the few and costly space projects that actually ends with a space flight. To overcome this threshold there is a strong need for flight opportunities where reasonable risks can be accepted. Currently there are a few flight opportunities allowing extensive use of new technology in space, but one of the exceptions is the PRISMA program. PRISMA is an international (Sweden, Germany, France, Denmark, Norway, Greece) technology demonstration program with focus on rendezvous and formation flying. It is a two satellite LEO mission with a launch scheduled for the first half of 2009. On PRISMA, a number of novel technologies e.g. RF metrology sensor for Darwin, autonomous formation flying based on GPS and vision-based sensors, ADN-based “green propulsion” will be demonstrated in space for the first time. One of the satellites will also have a miniaturized propulsion system onboard based on MEMS-technology. This novel propulsion system includes two microthruster modules, each including four thrusters with micro- to milli-Newton thrust capability. The novelty of this micropropulsion system is that all critical components such as thrust chamber/nozzle assembly including internal heaters, valves and filters are manufactured using MEMS technology. Moreover, miniaturized pressure sensors, relying on MEMS technology, is also part of the system as a self-standing component. The flight opportunity on PRISMA represents one of the few and thus important opportunities to demonstrate MEMS technology in space. The present paper aims at describing this development effort and highlights the benefits of miniaturized components and systems for space using MEMS technology. [Copyright &y& Elsevier]

This article reviews techniques applicable to enhancing the mixing of jets, with particular emphasis on infrared (IR) signature reduction of high-speed jets. Following a brief introduction to the IR signature of jet plumes and the fundamentals of jet mixing, this paper discusses rapid mixing technologies under the categories of: geometric modifications (to the nozzle); high shear stress mixing; normal stress mixing; self-acoustic excitation; external acoustic excitation; mechanically oscillated; self-oscillated. It is shown that mixing enhancements of the order of 100 per cent are possible with some techniques and that by combining techniques this can be increased by at least as much again. Simple geometric calculations are presented which demonstrate that with rectangular nozzles such high levels of mixing enhancement may be necessary in order to reduce IR signature. Some apparent rapid mixing technologies, however, have been shown to increase jet spreading without increasing entrainment, whereas other techniques can reduce entrainment as easily as they can increase it. [ABSTRACT FROM AUTHOR]

This article, continuation of previous works,5,3 presents the applications of geometric optimal control theory to the analysis of the Earth re-entry problem for a space shuttle where the control is the angle of bank, the cost is the total amount of thermal flux, and the system is subject to state constraints on the thermal flux, the normal acceleration and the dynamic pressure. Our analysis is based on the evaluation of the reachable set using the maximum principle and direct computations with the boundary conditions according to the CNES research project. The optimal solution is approximated by a concatenation of bang and boundary arcs, and is numerically computed with a multiple-shooting method. [ABSTRACT FROM AUTHOR]

The aim of this paper is to evaluate the feasibility of a mission to Mars using the Integrated Propulsion Systems (IPS) which means to couple Nuclear-MPD-ISPU propulsion systems. In particular both mission analysis and propulsion aspects are analyzed together with technological aspects.Identifying possible mission scenarios will lead to the study of possible strategies for Mars Exploration and also of methods for reducing cost.As regard to IPS, the coupling between Nuclear Propulsion (Rubbia''s engine) and Superconductive MPD propulsion is considered for the Earth–Mars trajectories: major emphasis is given to the advantages of such a system. The In Situ Resource Utilization (ISRU) concerns on-Mars operations; In Situ Propellant Utilization (ISPU) is foreseen particularly for LOX-CH4 engines for Mars Ascent Vehicles and this possibility is analyzed from a technological point of view. Tether Systems are also considered during interplanetary trajectories and as space elevators on Mars orbit.Finally strategic considerations associated to this mission are considered also. [Copyright &y& Elsevier]

This paper is concerned with the elastic buckling of vertical cylindrical shells under combined end pressure and body force. Such buckling problems are encountered when cylindrical shells are used in a high-g environment such as the launching of rockets and missiles under high-propulsive power. The vertical shells may have any combination of free, simply supported, and clamped ends. Based on the Goldenveizer-Novozhilov thin shell theory, the total potential energy functional is presented and the buckling problem is solved using the Ritz method. Highlight in the formulation is the importance of the correct potential energy functional which includes the shell shortening due to the circumferential displacement. The omission of this contributing term leads to erroneous buckling solutions when the cylindrical shell is not of moderate length (length-to-radius ratio smaller than 0.7 or larger than 3). New solutions for body-force buckling parameters are presented for stubby cylindrical shells to long tube-like shells that approach the behavior of columns. The effects of the shell thickness and length on buckling parameter are also investigated. [ABSTRACT FROM AUTHOR]

Small hybrid rocket motors using solid propellant and gaseous oxidizer are becoming increasingly popular as a propulsion device. This paper describes the development of a one-dimensional flow model for the design of a small rocket motor. Combustion of polyethylene as solid propellant with oxygen is used as a candidate hybrid fuel to test and evaluate the performance of this hybrid system. To assess the performance under different operating conditions, a computer program has been developed, which facilitates inputs to be varied and effects assessed. A system of governing equations is summarized in the main body of this paper and is numerically solved by the computer program. The results of the modelling are then used to design and build a small low-cost rocket motor for experimental verification. Therefore, the materials presented herein could be used in the future design of hybrid rocket motors. [ABSTRACT FROM AUTHOR]

Highlights three jet propulsion systems of rockets in the order of their feasibility automatic control over the thrust force vector. Partial control permitted by one of the magnetic propulsion system; Importance of the other magnetic propulsion systems; Magnetic propulsion system which can have total control over the direction and modulus of the electrodynamic thrust force vector.