Orbital Transfer Vehicle Research Articles

A holistic approach for efficient greener in-space propulsion

This paper presents a comprehensive framework for designing in-space propulsion systems, integrating four criteria: global propulsive performance, environmental impact, cost efficiency, and architectural reliability. The study focuses on the emerging class of Orbital Transfer Vehicles to illustrate the application of this method. By examining the synergistic potential of OTVs and greener propellants, the paper addresses different mission scenarios, including LEO, GEO, and lunar missions, with both scientific and commercial objectives. The proposed framework aims to go beyond traditional cost-centric approaches, offering a more complete evaluation method for early design phases. A case study comparing three liquid bipropellant options, pressure-fed MON-3/MMH, 98%-HTP/RP-1, and self-pressurizing N2O/Ethane, demonstrates the utility of the tool. Findings suggest that scientific missions benefit most from 98%-HTP/RP-1, while traditional propellants remain preferable for cost-driven commercial missions to GEO and the Moon, though greener alternatives are competitive for less demanding LEO missions. This innovative framework aims to guide the selection of propulsion systems to achieve greener space missions, aligning traditional performance figures with environmental responsibility.

Design of Orbital Payload transfer Vehicle

Design of Orbital Payload transfer Vehicle

Performance comparison of green propulsion systems for future Orbital Transfer Vehicles

This paper explores the relevance of a new class of upper stages commonly referred to as Orbital Transfer Vehicles (OTVs) or Kick-Stages, and the importance of green propellants in the development of such systems. Orbital Transfer Vehicles are currently developed for diverse applications by many entities and are expected to have a major impact in the future of space missions and overall space sector. Such systems are commonly connected to the topic of On-Orbit Servicing with the promise of unlocking new missions, including active space debris removal, multi-payload to multi-orbit delivery, in-orbit experiments, in-orbit refuelling and other generic services. Green propulsion is an increasingly prevalent trend in the space industry that has experienced significant growth in recent decades with the main objective of identifying suitable alternatives to commonly used, toxic liquid propellants for in-space applications. The study investigates the synergies between the two themes, underlining the advantages that a widespread use of green technologies could bring to the overall sector, but also examining the final benefits to the system design of OTVs. Due to the numerous and challenging propulsive system requirements associated with these systems, green technologies are analysed to determine their advantages and disadvantages in comparison with current concepts, outlining current trends and expected future developments. The focus is on attainable performances of propulsion systems with respect to the required dry and inert mass. Various architectures of the different green alternatives based on hydrogen peroxide and the more peculiar self-pressurized propellants are proposed, exploring green propellants-based designs that can offer synergies and advantages over classical ones. These new architectures offer performance comparable to classical and widely available toxic alternatives and present the important advantage of using easy-to-handle propellants. The development of such new options has the potential to enable innovative future designs by improving mass and size requirements of new propulsion systems, thereby enhancing the current landscape.

Employing a multi-sensor fusion array to detect objects for an orbital transfer vehicle to remove space debris

PurposeThis article takes into account object identification, enhanced visual feature optimization, cost effectiveness and speed selection in response to terrain conditions. Neither supervised machine learning nor manual engineering are used in this work. Instead, the OTV educates itself without instruction from humans or labeling. Beyond its link to stopping distance and lateral mobility, choosing the right speed is crucial. One of the biggest problems with autonomous operations is accurate perception. Obstacle avoidance is typically the focus of perceptive technology. The vehicle's shock is nonetheless controlled by the terrain's roughness at high speeds. The precision needed to recognize difficult terrain is far higher than the accuracy needed to avoid obstacles.Design/methodology/approachRobots that can drive unattended in an unfamiliar environment should be used for the Orbital Transfer Vehicle (OTV) for the clearance of space debris. In recent years, OTV research has attracted more attention and revealed several insights for robot systems in various applications. Improvements to advanced assistance systems like lane departure warning and intelligent speed adaptation systems are eagerly sought after by the industry, particularly space enterprises. OTV serves as a research basis for advancements in machine learning, computer vision, sensor data fusion, path planning, decision making and intelligent autonomous behavior from a computer science perspective. In the framework of autonomous OTV, this study offers a few perceptual technologies for autonomous driving in this study.FindingsOne of the most important steps in the functioning of autonomous OTVs and aid systems is the recognition of barriers, such as other satellites. Using sensors to perceive its surroundings, an autonomous car decides how to operate on its own. Driver-assistance systems like adaptive cruise control and stop-and-go must be able to distinguish between stationary and moving objects surrounding the OTV.Originality/valueOne of the most important steps in the functioning of autonomous OTVs and aid systems is the recognition of barriers, such as other satellites. Using sensors to perceive its surroundings, an autonomous car decides how to operate on its own. Driver-assistance systems like adaptive cruise control and stop-and-go must be able to distinguish between stationary and moving objects surrounding the OTV.

Development of a lanthanum hexaboride hollow cathode for a magnetic octupole thruster

Electric propulsion has undergone significant advancements, leading to increased usage in satellite constellation management and orbital transfer vehicles. This study describes the design and development of a lanthanum hexaboride (LaB6) hollow cathode for the Magnetic Octupole Plasma Thruster (MOPT), a novel electric propulsion concept for future space missions. Its operational characteristics are also provided. To gain insights into the coupling behavior between the cathode plume and the octupole magnetic field, the hollow cathode was operated with and without an octupole magnet over a range of discharge currents and flow rates. The results indicate that the discharge voltage was anti-correlated with the discharge current and flow rate. Notably, the octupole magnetic field observed reduced discharge voltage oscillations in both high- and low-discharge-current regimes. Nonetheless, the existence of the magnetic field was associated with the amplification of ionization-like instabilities in the high-flow-rate regime. Finally, the paper also delves into the plasma behavior under the influence of the octupole magnetic field, signifying an advancement in developing LaB6 hollow cathodes for their potential use in space-based MOPT applications.

Nuclear fusion powered Titan aircraft

This paper discusses a system for Titan exploration enabled by nuclear fusion power. Titan is one of the most interesting locations in the solar system with a thick atmosphere, surface oceans, under-ice oceans and complex terrain. This paper provides a conceptual design of a fusion-powered system to explore many parts of Titan and enable the use of high-power instruments. The design includes a fusion-powered orbital transfer vehicle and an electric Titan science aircraft. A Direct Fusion Drive (DFD) propulsive engine could bring a sizable spacecraft to Titan orbit in less than two years. A second fusion reactor, configured as a closed-loop power generator, would be used for an electric Titan science aircraft. Both reactors are based on the Princeton Field-Reversed Configuration (PFRC) concept which combines an FRC with a magnetic mirror. PFRC uses a novel radio-frequency plasma heating system and deuterium-helium-3 fuel. A lower temperature plasma flows around the closed-field FRC region removing the fusion products. In the DFD propulsive configuration, this secondary flow permits direct and variable thrust and exhaust velocity. The science aircraft would do a powered entry to Titan and then have the capability to fly anywhere on the moon at subsonic speeds. The DFD-powered transfer vehicle would allow the in-orbit transfer stage to change inclination as needed to cover different areas of the surface.

Mitigating CubeSat confusion: Results of in-flight technical demonstrations of candidate tracking and identification technologies

CubeSats and nanosats offer affordable access to space. Their standardized size and shape have allowed a CubeSat industrial ecosystem to flourish, which has lowered costs for acquisition and launch, and vastly shortened development times. This has enabled access to space for many nontraditional actors in the space arena. What previously might have taken years to develop can now be accomplished in months, and for significantly less money. This has allowed educational institutions to fly their own CubeSats. It has also allowed numerous developing countries their first satellite. It has allowed more traditional aerospace companies and government agencies very rapid cycle times for research and development efforts. However, when scores of CubeSats are launched en masse, many cannot be immediately identified, some not even after months, if ever. This inability to identify the launched CubeSats from their cohort is what gives rise to “CubeSat confusion”. There are many factors to mitigate this issue, but chief among them would be the ability of a CubeSat owner/operator to independently identify their CubeSat, perhaps using one of various specialized techniques or technologies to aid in this regard.One of the major launch aggregators, Spaceflight Inc., has announced the opportunity for identification and tracking technology providers to fly candidate devices aboard their Sherpa orbital transfer vehicle as hosted payloads, allowing on-orbit technical demonstration and raising the technology readiness level for the devices. We report the results of several technical demonstrations, their potential utility to the CubeSat community, and future plans. The demonstrations will be of interest to both CubeSat/nanosat owner/operators, as well as Space Traffic Management government regulators.

Quasi-steady testing approach for high-power Hall thrusters

Hall effect thrusters operating at power levels in excess of several hundreds of kilowatts have been identified as enabling technologies for applications such as lunar tugs, large satellite orbital transfer vehicles, and solar system exploration. These large thrusters introduce significant testing challenges due to the propellant flow rate exceeding the pumping speed available in most laboratories. Even with proposed upgrades in mind, the likelihood that multiple vacuum facilities will exist in the near future to allow long duration testing of high-power Hall thrusters operating at power levels in excess of 100 kW remains extremely low. In this article, we numerically explore the feasibility of testing Hall thrusters in a quasi-steady mode defined by pulsing the mass flow rate between a nominal and a low value. Our simulations indicate that sub-second durations available before the chamber reaches critical pressure are sufficiently long to achieve the steady-state current and flow field distributions, allowing us to characterize thruster performance and the near plume region.

The on-orbit mission analysis of OTV based on DoDAF

PurposeThe purpose of this paper is to improve the efficiency of on-orbit operations through the top-level task design based on DoDAF. Based on the existing upper stage rocket technology, orbit transfer vehicles (OTVs) have developed rapidly in recent years. However, the lack of decision guidance based on overall task analysis requires integrating top-level analysis and bottom-level execution to achieve the smooth development of full-process tasks.Design/methodology/approachUsing the Department of Defense Architecture Framework (DoDAF) as a reference, this paper performs the top-level mission analysis modeling of the on-orbit rendezvous and capture of the OTV. Moreover, the typical operational view products are obtained, and the cooperative relations among the mission requirements, the system requirements, and the functional requirements are also analyzed.FindingsThe results show that the attitude of the OTV changes violently during the maneuver and rendezvous phases. In addition, the view products can be optimized based on the results.Originality/valueThe proposed DoDAF-based on-orbit task integration analysis method achieves the effective fusion of high-level analysis and bottom-level execution of OTV on-orbit rendezvous and capture task through the top-level task modeling, operation view generation and task relationship analysis. According to the requirements and constraints of the on-orbit rendezvous and capture task, the control instructions of the vehicle are efficiently generated under the DoDAF framework.

Analytical Solutions to Aeroassisted Orbital Transfer Problem

In this article, the atmospheric flight of aeroassisted orbit transfer vehicle is investigated analytically. As the flight speed is greater than the first cosmic velocity, the centrifugal force is also larger than gravity. Due to this, the vehicle has to fly with a downward lift and the flight altitude is extremely unstable. To cope with this challenge, the high-precision formulae of altitude and flight-path angle for steady glide are deduced and used to develop the steady glide control technique to stabilize the flight altitude. Upon steady glide, the analytical solutions for 3-D aeroassisted maneuvering trajectory are derived. In the derivation, a reduced-order system with negative lift-to-drag ratio as control variable is obtained by subtly simplifying the nonlinear flight dynamics. To address the complex influence of Coriolis force, the regular perturbation method is used such that the system is further decomposed into two special linear time-varying systems, which can be solved analytically by the spectral-decomposition-based method. Due to the full consideration of the influence of the Earth's curvature and rotation, the analytical solutions can predict the downrange, crossrange, and heading angle of the 3-D trajectory accurately.

Method of Optimizing Maneuvers of an Orbital Transfer Vehicle in a Strong Central Field of Gravity

A method of constructing the optimal trajectories of interorbital flights of an orbital transfer vehicle with a low thrust propulsion system is described. Testing the method against the problem of delivering the maximum payload from a low orbit to an orbit close to the geostationary one confirms its effectiveness.

Optimization of Liquid Rocket propulsion system for Reusable Orbital Transfer Vehicle

The work is devoted to the investigation of possible ways to increase the efficiency of reusable orbital transfer vehicle (R-OTV) by using a multi-criteria optimization procedure of liquid rocket propulsion system (LRPS) design parameters. The “Federation” spacecraft is considered as an example payload of R-OTV for orbital transfer mathematical modeling, due to its known initial mass of 16,500 kg and other design parameters. Payload mass and specific cost of payload transfer are chosen as optimality criteria.Multiple tasks are solved during the work: mathematical model and corresponding software for determining the optimal design parameters of R-OTV LRPS by mass and cost criteria are developed, R-OTV and LRPS calculation and optimization are conducted, the technical appearance of LRPS is designed.Mathematical modeling is used to solve the presented problem. A developed mathematical model has a modular structure, in particular, it has modules for liquid rocket engine (LRE) mass calculating, OTV (Orbital Transfer Vehicle) mass calculating module, payload mass, design cost and manufacturing cost calculating modules for LRE and OTV and module for calculating the specific cost of payload transfer to target orbit. The intermediate result is used by LRPS project parameters design optimization module, then by technical appearance design module and LRPS units design module.Multiple calculations are conducted with a developed mathematical model. Verification of the model is carried out by developing the technical appearance of real OTVLRPS. Several variants of LRPS for spacecraft “Federation” type OTV are designed for low lunar orbit, OTV parameters are compared with optimized LRPS and with existing LRPS.

Trajectory Optimization of a Low-Thrust Geostationary Orbit Insertion for Total Ionizing Dose Decrease

The paper considers a method of the dose exposure decrease from the charged particles of Earth’s radiation belts (ERBs) affecting a reusable orbital transfer vehicle, which is launched from a low circular orbit into geostationary orbit using a nuclear electric propulsion system. The main idea of the method consists in numerical continuation of a solution of the minimum time problem with respect to an ionizing radiation dose accumulated at the end of a transfer. To do this, equations of motion of the orbital transfer vehicle are supplemented by an additional equation for the radiation dose, and the boundary condition for the dose at the right end is introduced. In calculating the dose, the AE8/AP8 MIN, AE8/AP8 MAX, and AE9/AP9 models of fluxes of charged particles of ERBs were used. By changing the insertion trajectory shape, it became possible to lower the radiation dose by 25–38% relative to the minimum time trajectory. At the same time, the transfer time increased no more than by 7% of the minimum time of launching into geostationary orbit, and the characteristic velocity expenses increased by 320–560 m/s.

Development of a new MLI for orbital cryogenic propulsion systems -thermal performance under one atmosphere to a vacuum

The efficient storage of cryogenic propellants is among the key technologies for long-duration space exploration missions. For the orbital transfer vehicle, the required thermal insulation performance is more than ten times higher than that of conventional spray-on foam insulation. Conventional multi-layer insulation blankets are used as excellent insulation for spacecraft in the vacuum environment, but are not usable in the atmospheric environment. A new type of insulation-a load-bearing, non-interlayer-contact spacer MLI (LB-NICS MLI)-has been developed. The insulation performance in both air and vacuum environments is measured with a boil-off calorimeter. According to the test results, LB-NICS MLI is about 3 times superior under 1 atmosphere and 13 times superior under a vacuum, compared to foam insulation.

Conceptual design of the space disposal system for the highly radioactive component of the nuclear waste

A system for space disposal of the most radioactive component of the spent fuel produced by the nuclear reactors is designed conceptually with an intent of ensuring safety. Unlike the earlier design by NASA, this system is to use a commercially available launch vehicle and be launched from the western Pacific Ocean. It will fly low until it acquires Earth-escape velocity, so that the nuclear package will not fall on land in case of system malfunctioning. The package will be parked in one of the lunar libration points L4 or L5, and kept vigilant by a watch-dog orbital transfer vehicle which will push the package away in case an asteroid approaches. The calculation shows that such a safe system will be feasible and relatively cheap. The details of how the nuclear package is made, how the launch vehicle is made and flown, and how the nuclear package is managed in the final parking orbit are described.

Orbital transfer vehicle (OTV) system sizing study for manned GEO satellite servicing

ABSTRACTIn an effort to quantify the feasibility of candidate space architectures for astronauts servicing Geosynchronous Earth Orbit (GEO) satellites, a conceptual assessment of architecture-concept and operations-technology combinations has been performed. The focus has been the development of a system with the capability to transfer payload to and from geostationary orbit. Two primary concepts of operations have been selected: (a) Direct insertion/re-entry (Concept of Operations 1 – CONOP 1); (b) Launch to low-earth orbit at Kennedy Space Center inclination angle with an orbital transfer to/from geostationary orbit (Concept of Operations 2 – CONOP 2). The study concludes that a capsule and de-orbit propulsion module system sized for the geostationary satellite servicing mission is feasible for a direct insertion/re-entry concept of operation CONOP 1. Vehicles sized for CONOP 2 show overall total mass savings when utilising the aero-assisted orbital transfer vehicle de-orbit propulsion module options compared to the pure propulsive baseline cases. Overall, the consideration of technical, operational and cost factors determine if either the aero-assisted orbital transfer vehicle concepts or the re-usable/expendable ascent/de-orbit propulsion modules is the preferred option.

Bi-objective optimization of a multiple-target active debris removal mission

The increasing number of space debris in Low-Earth Orbit (LEO) raises the question of future Active Debris Removal (ADR) operations. Typical ADR scenarios rely on an Orbital Transfer Vehicle (OTV) using one of the two following disposal strategies: the first one consists in attaching a deorbiting kit, such as a solid rocket booster, to the debris after rendezvous; with the second one, the OTV captures the debris and moves it to a low-perigee disposal orbit. For multiple-target ADR scenarios, the design of such a mission is very complex, as it involves two optimization levels: one for the space debris sequence, and a second one for the “elementary” orbit transfer strategy from a released debris to the next one in the sequence. This problem can be seen as a Time-Dependant Traveling Salesman Problem (TDTSP) with two objective functions to minimize: the total mission duration and the total propellant consumption.In order to efficiently solve this problem, ONERA has designed, under CNES contract, TOPAS (Tool for Optimal Planning of ADR Sequence), a tool that implements a Branch & Bound method developed in previous work together with a dedicated algorithm for optimizing the “elementary” orbit transfer. A single run of this tool yields an estimation of the Pareto front of the problem, which exhibits the trade-off between mission duration and propellant consumption. We first detail our solution to cope with the combinatorial explosion of complex ADR scenarios with 10 debris. The key point of this approach is to define the orbit transfer strategy through a small set of parameters, allowing an acceptable compromise between the quality of the optimum solution and the calculation cost. Then we present optimization results obtained for various 10 debris removal scenarios involving a 15-ton OTV, using either the deorbiting kit or the disposal orbit strategy.We show that the advantage of one strategy upon the other depends on the propellant margin, the maximum duration allowed for the mission and the orbit inclination domain. For high inclination orbits near 98 deg, the disposal orbit strategy is more appropriate for short duration missions, while the deorbiting kit strategy ensures a better propellant margin. Conversely, for lower inclination orbits near 65 deg, the deorbiting kit strategy appears to be the only possible with a 10 debris set. We eventually explain the consistency of these results with regards to astrodynamics.

A passive thermal control method by radiative shield with multi-parameter adjustment capacity

To ensure the stability and reliability of the orbit transfer vehicle and its equipments, a multi-parameter adjustable radiative shield is adopted in the passive thermal control method. The principle of the radiative shield by radiation heat transfer is first expounded and the multi-parameter adjustment strategy is then thoroughly analyzed. Taking a typical power system which is used in orbital transfer vehicle as an example, temperature evolutions of the storage tank, pipelines and electrical equipments are obtained by numerical simulations. The thermal control results of the radiative shield method, in which the multi-parameter adjustable characteristic is dominated, is compared with those of thermal insulation technique. Comparison results indicated that the proposed passive thermal control method completely takes advantage of the thermal inertia and radiative properties of the radiative shield, which extends the high thermal flux at both time and spatial domains.

Behavior of tethered debris with flexible appendages

Active exploration of the space leads to growth of a near-Earth space pollution. The frequency of the registered collisions of space debris with functional satellites highly increased during last 10 years. As a rule a large space debris can be observed from the Earth and catalogued, then it is possible to avoid collision with the active spacecraft. However every large debris is a potential source of a numerous small debris particles. To reduce debris population in the near Earth space the large debris should be removed from working orbits. The active debris removal technique is considered that intend to use a tethered orbital transfer vehicle, or a space tug attached by a tether to the space debris. This paper focuses on the dynamics of the space debris with flexible appendages. Mathematical model of the system is derived using the Lagrange formalism. Several numerical examples are presented to illustrate the mutual influence of the oscillations of flexible appendages and the oscillations of a tether. It is shown that flexible appendages can have a significant influence on the attitude motion of the space debris and the safety of the transportation process.

Computational Study on the Influence of jet on Reduction of Drag Over Cone Flare Bodies in Hypersonic Turbulent Flow

Recently, Nhypersonic research activities around the universe have been in major focus because of the milestone developments in hypersonic reentry vehicles, orbital transfer vehicles, reusable launch vehicles and space recovery experimental modules. One of the major problems in hypersonic flight is drag and with the tremendous progress in computational powers it can be analysed. In this work, the effects of counter flow jet on reduction of drag around two blunt cone flare bodies in the hypersonic turbulent flow are investigated through a numerical study. Flow field around the blunt bodies is calculated numerically for the free stream Mach number of 6.5. Numerical solutions of Navier-Stokes equation and energy equation governing the turbulent flow of compressible fluid are obtained by adopting Finite Volume Method (FVM) through an industry standard CFD code, FLUENT 6.3.26 package. Shear Stress Transport (SST) model was used in the computation of pressure drag, skin friction drag and total drag for both the absence of jet and the presence of jet cases for both the configuration. Contours of Mach number are presented to describe the flow patterns. It is clear that the reduction of drag was greatly influenced by the jet conditions and body shapes.