Low Thrust Levels Research Articles

Turbopump Parametric Modelling and Reliability Assessment for Reusable Rocket Engine Applications

The development of modern reusable launchers, such as the Themis project with its LOX/LCH4 Prometheus engine, CALLISTO—a reusable VTVL-launcher first-stage demonstrator with a LOX/LH2 RSR2 engine, and SpaceX’s Falcon 9 with its Merlin 1D engine, underscores the need for advanced control algorithms to ensure reliable engine operation. The multi-restart capability of these engines imposes additional requirements for throttling, necessitating an extended controller-validity domain to safely achieve low thrust levels across various operating regimes. This capability also increases the risk of component failure, especially as engine parameters evolve with mission profiles. To address this, our study evaluates the dynamic reliability of reusable rocket engines (RREs) and their subcomponents under different failure modes using multi-physics system-level modelling and simulation, with a particular focus on turbopump components. Transient condition modelling and performance analysis, conducted using EcosimPro-ESPSS software (version 6.4.34), revealed that turbopump components maintain high reliability under nominal conditions, with turbine blades demonstrating significant fatigue life even under varying thermal and mechanical loads. Additionally, the proposed predictive model estimates the remaining useful life of critical components, offering valuable insights for improving the longevity and reliability of turbopumps in reusable rocket engines. This study employs deterministic, thermally dependent structural simulations, with key control objectives including end-state tracking of combustion chamber pressure and mixture ratios and the verification of operational constraints, exemplified by the LUMEN demonstrator engine and the LE-5B-2 engine class.

A Methodology for the Preliminary Design of a High-Efficiency Multistage Plasma Thruster

AbstractSpace electric propulsion represents a class of power-limited systems that utilize the interaction of electromagnetic fields with ionized inert gas propellants to generate thrust. This technology has emerged as a highly fuel-efficient and sustainable alternative to chemical propulsion systems, particularly for satellite constellations. However, the miniaturization potential of EP systems is impeded by certain limitations, necessitating the exploration of novel architectures. The high-efficiency multistage plasma thruster (HEMP-T) stands as a promising contender for stand-alone missions due to its employment of a cusped magnetic-field topology, which effectively mitigates plasma-wall interactions and enhances overall efficiency even at low thrust levels. Despite the growing interest in HEMP-Ts, there is a dearth of comprehensive and streamlined preliminary design procedures for these systems. Prior research has predominantly focused on extensive numerical analyses, neglecting the development of efficient and accessible design tools. To bridge this gap, this paper presents a novel preliminary design tool derived from integrating established analytical models available in the literature. The proposed design tool also incorporates an iterative procedure that refines geometric properties using a 2D magnetostatic solver. Through the application of this tool, a 4 mN HEMP thruster was analyzed. This finally exhibited a specific impulse of approximately 2000s and a good efficiency level of 23%. Also, the results obtained for a 10 mN application align closely with those achieved by other types of EP thrusters.

A novel calculation method for low-thrust transfer trajectories in the Earth-Moon restricted three-body problem

This paper introduces a novel trajectory calculation method for achieving continuous low-thrust transfers from the geostationary orbit to libration point orbits within the Earth-Moon restricted three-body framework. Low-thrust engines can offer benefits such as high specific impulse and long operating lifetimes, making them suitable for deep space exploration. However, the primary challenge lies in the low thrust level, leading to prolonged flight duration for spacecraft to overcome Earth's gravitational pull. Furthermore, the transfer trajectories exhibit multiple revolutions around the Earth, adding complexity to the calculations. To solve these difficulties, a novel trajectory computation method is presented, which for the first time discusses dynamical evolution beyond a predefined time interval in the context of an optimal control problem specific to low-thrust engines. This method employs a backward integration technique applied to the known optimized solution, coupled with a deliberate extension of the flight time to generate an initial guess for subsequent trajectory optimization, significantly simplifying the computational procedure. Several application scenarios involving transfer from the geostationary orbit to libration point orbits are demonstrated. The results confirm the effectiveness of this method, providing a rapid computational strategy for similar low-thrust transfer missions.

Very low thrust station-keeping for low Earth orbiting satellites

Traditional station-keeping for Earth observation satellites with chemical thrusters generally involves maneuvers every couple months that are able to change significantly the semi-major axis and the inclination. These strategies do not scale down to very low thrust level (a few hundreds of μN) electrical thrusters. This paper presents both in-plane and out-of-plane strategies that spread corrections over very long arcs and discretize them to tiny maneuvers every couple orbits, taking into account mission-constraints on maneuvers locations. These strategies scale up to medium thrust strategies, filling the gap between propulsion technologies. The out-of-plane strategy although features a new no-deadband property and controls the full orbital momentum. All strategies allow control very close to the reference (a few hundreds meters in osculating parameters) and very low cost.

Using Wavy Fuel Surface to Improve the Performance of the Solid Fuel Ramjets

AbstractThis paper introduces a new idea to augment the fuel regression rate in solid fuel ramjets. The heat transfer rate from the reacting flow field into the surface of the solid fuel grain determines directly its regression rate in this propulsion system, which is very low in comparison with solid rocket motors leads the low thrust levels. This innovative simple practical method is based on the geometrical treatment of the fuel grain, making the smooth ribs as a wavy surface, to increase the turbulence and the heat transfer rate into the evaporating surface where the fuel flow rate is determined. The finite volume solver of the turbulent reacting compressible flow is developed to study the flow field where the burning rate is computed using the conjugate heat transfer for the solid fuel. The influence of the amplitude and the period of the cosine function as the fuel surface geometry are investigated on the regression rate, and it is shown that the amplitude is more effective, and utilizing such surface treatments helps to augment considerably the fuel average regression rate.

Optimal Control of Transfer to Vertical Orbits from Lyapunov Orbits Using Low-Thrust Engine

There are many different families of periodic orbits in the Earth-Moon system, such as Lyapunov orbits, halo orbits, vertical orbits, etc. The establishment of a lunar space station requires a spacecraft to be able to transfer among these orbits. Lyapunov orbits have been used by some missions and are well-studied orbits, while periodic vertical orbits can provide large amplitudes of spacecraft motion outside the plane of the Moon’s motion, which makes it possible to avoid shadowing of the orbits and use them as relay satellite in cislunar space. Modern researchers mainly consider the use of high-thrust engines for transfer. With the development of electric propulsion technology, the use of low, but long-acting thrust for deep space exploration has become especially relevant. This is due to the high specific characteristics of the propulsion systems of this type. In this article, an algorithm has been developed for determining the optimal control with a low-thrust engine for a transfer from Lyapunov orbit to a vertical orbit. The minimum time of flight or the minimum costs of the working body are used as criteria for optimality. In the calculation for solving the two-point boundary value problem of the optimal control theory, the parameter continuation algorithm is used, which allows to gradually get the transfer from some simple results to the final transfer trajectory. The results obtained make it possible to assert that the use of intermediate axial orbits allows the use of propulsion systems with lower thrust levels. In this case, the duration of the flight increases slightly with an almost unchanged consumption of the working body. Moreover, the homotopy method makes it possible to reduce the consumption of working body, while the control of the engine throttling becomes discrete. The results of this study and the algorithms proposed in this article can be used to determine the optimal program control and ballistic design of lunar missions.

Conceptual Study on Low-Melting-Point Thermoplastic Fuel/Nitrous Oxide Hybrid Rockoon

A hybrid rockoon, which is a hybrid rocket launched from a high-altitude balloon, is proposed. The rocket system configuration was studied for single-stage or multiple-stage rockets, and the performance and range safety requirements were considered. The propellants of the hybrid-rocket motor are a low-melting-point thermoplastic fuel and nitrous oxide, which are beneficial, owing to their mechanical properties at cold temperatures experienced in high-altitude environments. Three-dimensional launch trajectory analyses were performed for science missions aimed at sampling cosmic dust levitating in the upper stratosphere. The apogee altitude can be significantly increased by elevating the altitude of the launch point because the problem of low thrust level of the hybrid rocket is solved by increasing the nozzle expansion ratio. Suborbital trajectories with multiple apogee points and orbital missions to extremely low Earth orbits are presented, which provide a long flight path in the upper atmosphere. The sequence of events and flight characteristics of the proposed hybrid rockoon system are discussed, and necessary technological improvements in the structure and propulsion systems are presented.

Parametric optimization of aircraft arrival trajectories for aviation noise mitigation using BADA4 performance model

Successful mitigation of aviation noise is a key enabler for sustainable aviation growth. A key focus of this effort is the noise arising from aircraft arrival operations. Arrival operations are characterized by the use of high-lift devices, deployment of landing gear, and low thrust levels, which results in the airframe being the major component of noise. In order to optimize for arrival noise, management of the flap schedule and gear deployment is crucial. This research aims to create an optimization framework for evaluating various aircraft trajectories in terms of their noise impact. A parametric representation of the aircraft arrival trajectory will be created to allow for the variation of aircraft's flap schedule. The Federal Aviation Administration's Aviation Environmental Design Tool will be used to simulate the aircraft trajectory and performance, and to compute the noise metrics. Specifically, the latest performance model from EUROCONTROL called "Base of Aircraft Data - Family 4" will be used. This performance model contains higher fidelity modeling of aircraft aerodynamics and other characteristics which allows for better parametric variation.

Numerical investigation of effect of crossflow transition on rotor blade performance in hover

In the present study, the effect of crossflow transition on the rotor blade performance in hover was investigated numerically. The simulations were conducted for the Pressure Sensitive Paint (PSP) rotor in hover by using a Reynolds-averaged Navier-Stokes (RANS) solver based on unstructured meshes. The k−ωSST model was used for simulating the turbulent flow fields and for calculating the turbulent eddy viscosity. For predicting the laminar-turbulent onset phenomena involving crossflow-induced transition, the γ−Reθt−CF+ transition model was adopted. For the comparison of the transition locations without considering the crossflow transition effect, simulations were also carried out using the γ−Reθt transition model. The calculations were made for the PSP rotor at collective pitch angles from four to 12 degrees with an interval of two degrees. To investigate the effect of blade tip Mach number, two blade tip Mach numbers of 0.585 and 0.65 were tested. The predicted results such as the transition onset location and the rotor aerodynamic performance in terms of thrust coefficient, torque coefficient and figure of merit were compared with experimental data. It was found that proper consideration of the effect of crossflow transition is critical for the accurate prediction of the laminar-turbulent transition onset location on rotor blades. The figure of merit also compares better with the experimental data when the effect of laminar-turbulent transition is included. With the addition of the crossflow transition, the figure of merit is slightly decreased, particularly at low thrust levels.

Performance Analysis and Mass Estimation of a Small-Sized Liquid Rocket Engine with Electric-Pump Cycle

The propellant supply system of a liquid rocket engine using an electric pump has high reliability because of the relatively small number of components. The system also has the merit of a quick response and ease of control owing to its simple configuration. Recently, the rocket lab developed the Rutherford engine, which has an electric pump cycle, because of the improved technology in the electric motor and battery. This paper examined the development of the electric-pump cycle and compared the performance with other cycles for a small-sized low-thrust rocket engine. Performance analysis and mass estimation were conducted using the developed analysis program, in which reliability in mass estimation was improved based on the designed configuration or real performance data from commercial products. In addition, the modeling method and analysis procedure were described in detail. The results showed that it is possible to develop a small-sized engine with an electric-pump cycle when the present technologies are applied. The electric-pump cycle had a smaller dry mass than the gas-generator cycle, even at a low thrust level of 500 N, and showed higher performance in specific impulse and speed increments.

Effects of the O/F Ratio on the Performance of a Low Thrust LOX/Methane Rocket Engine with an ElecPump-fed Cycle

Recently, electrically driven pump-fed cycle rocket engines (i.e., ElecPump engines) have received considerable attention. ElecPump engines are superior to pressure-fed engines and will likely replace gas generator cycle engines under a 100 kN thrust level. This study focuses on LOX/methane ElecPump engines with a low-thrust level and estimates the nominal operation point of the engine. The objective variable was set to the maximum velocity increment of the stage. The mixture ratio and combustion pressure were explored as the design variables. Thus, the optimal mixture ratio was proposed as the design variable, considering not only the engine weight but also the structural weight.

Nonvolatile Particulate Matter Emissions of a Business Jet Measured at Ground Level and Estimated for Cruising Altitudes.

Business aviation is a relatively small but steadily growing and little investigated emission source. Regarding emissions, aircraft turbine engines rated at and below 26.7 kN thrust are certified only for visible smoke and are excluded from the nonvolatile particulate matter (nvPM) standard. Here, we report nvPM emission characteristics of a widely used small turbofan engine determined in a ground test of a Dassault Falcon 900EX business jet. These are the first reported nvPM emissions of a small in-production turbofan engine determined with a standardized measurement system used for emissions certification of large turbofan engines. The ground-level measurements together with a detailed engine performance model were used to predict emissions at cruising altitudes. The measured nvPM emission characteristics strongly depended on engine thrust. The geometric mean diameter increased from 17 nm at idle to 45 nm at take-off. The nvPM emission indices peaked at low thrust levels (7 and 40% take-off thrust in terms of nvPM number and mass, respectively). A comparison with a commercial airliner shows that a business jet may produce higher nvPM emissions from flight missions as well as from landing and take-off operations. This study will aid the development of emission inventories for small aircraft turbine engines and future emission standards.

Chemical composition and radiative properties of nascent particulate matter emitted by an aircraft turbofan burning conventional and alternative fuels

Abstract. Aircraft engines are a unique source of carbonaceous aerosols in the upper troposphere. There, these particles can more efficiently interact with solar radiation than at ground. Due to the lack of measurement data, the radiative forcing from aircraft exhaust aerosol remains uncertain. To better estimate the global radiative effects of aircraft exhaust aerosol, its optical properties need to be comprehensively characterized. In this work we present the link between the chemical composition and the optical properties of the particulate matter (PM) measured at the engine exit plane of a CFM56-7B turbofan. The measurements covered a wide range of power settings (thrust), ranging from ground idle to take-off, using four different fuel blends of conventional Jet A-1 and hydro-processed ester and fatty acids (HEFA) biofuel. At the two measurement wavelengths (532 and 870 nm) and for all tested fuels, the absorption and scattering coefficients increased with thrust, as did the PM mass. The analysis of elemental carbon (EC) and organic carbon (OC) revealed a significant mass fraction of OC (up to 90 %) at low thrust levels, while EC mass dominated at medium and high thrust. The use of HEFA blends induced a significant decrease in the PM mass and the optical coefficients at all thrust levels. The HEFA effect was highest at low thrust levels, where the EC mass was reduced by up to 50 %–60 %. The variability in the chemical composition of the particles was the main reason for the strong thrust dependency of the single scattering albedo (SSA), which followed the same trend as the fraction of OC to total carbon (TC). Mass absorption coefficients (MACs) were determined from the correlations between aerosol light absorption and EC mass concentration. The obtained MAC values (MAC532=7.5±0.3 m2 g−1 and MAC870=5.2±0.9 m2 g−1) are in excellent agreement with previous literature values of absorption cross section for freshly generated soot. While the MAC values were found to be independent of the thrust level and fuel type, the mass scattering coefficients (MSCs) significantly varied with thrust. For cruise conditions we obtained MSC532=4.5±0.4 m2 g−1 and MSC870=0.54±0.04 m2 g−1, which fall within the higher end of MSCs measured for fresh biomass smoke. However, the latter comparison is limited by the strong dependency of MSC on the particles' size, morphology and chemical composition. The use of the HEFA fuel blends significantly decreased PM emissions, but no changes were observed in terms of EC∕OC composition and radiative properties.

Difficulties in Dual-Mode Operation of Nontoxic Hypergolic Thruster Using Decomposition of

This study aimed to conceptually introduce a novel method of dual-mode operation of nontoxic hypergolic thruster using decomposition gas of hydrogen peroxide () and to experimentally assess the technical feasibility of the concept. An engineering model of a 500-N-scale nontoxic hypergolic thruster was tested with a nontoxic hypergolic combination that consisted of hydrocarbon fuel (“stock 3”) and 95 wt % as an oxidizer. Technically, dual-mode operation involves transition from a bipropellant mode to a monopropellant mode, or vice versa, which can provide great flexibility in terms of the deep throttling of the thrust level and the suppression of chugging instability at low thrust levels. The phase change of the oxidizer played a key role in the dual-mode operation. In particular, through hot-firing tests, system instabilities related to the mode transition were experimentally investigated. It was also confirmed that deep throttling of the thrust level was feasible through the mode transition.

Analysis of Nonisothermal Rarefied Gas Flow in Diverging Microchannels for Low-Pressure Microresistojets

Heat transfer and fluid flow through different microchannel geometries in the transitional regime (rarefied flow) are analyzed by means of direct simulation Monte Carlo (DSMC) simulations. Four types of three-dimensional microchannels, intended to be used as expansion slots in microresistojet concepts, are investigated using nitrogen as working fluid. The main purpose is to understand the impact of the channel geometry on the exit velocity and the transmission coefficient, parameters which are well known to affect directly the thruster performance. Although this analysis can be applied in principle to several possible microfluidics scenarios, particular focus is given to its application in the field of space propulsion for micro-, nano-, and picosatellites, for which the requirements ask for low thrust levels from some micronewtons to a few millinewtons and moderate specific impulse, as well as a low power consumption in the order of a few watts. Analysis shows that the thrust produced by one single microchannel can be increased by about 480% with a careful selection of the channel geometry, decreasing at the same time the specific impulse by just 5%, with a power consumption decrease of more than 66.7%.

Preliminary conceptual design of a new moderated reactor utilizing an LEU fuel for space nuclear thermal propulsion

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 klbf ∼ 11.9 klbf), 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 H2 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.

Conjunction challenges of low-thrust geosynchronous debris removal maneuvers

The conjunction challenges of low-thrust engines for continuous thrust re-orbiting of geosynchronous (GEO) objects to super-synchronous disposal orbits are investigated, with applications to end-of-life mitigation and active debris removal (ADR) technologies. In particular, the low maneuverability of low-thrust systems renders collision avoidance a challenging task. This study investigates the number of conjunction events a low-thrust system could encounter with the current GEO debris population during a typical re-orbit to 300km above the GEO ring. Sensitivities to thrust level and initial longitude and inclination are evaluated, and the impact of delaying the start time for a re-orbiting maneuver is assessed. Results demonstrate that the mean number of conjunctions increases hyperbolically as thrust level decreases, but timing the start of the maneuver appropriately can reduce the average conjunction rate when lower thrust levels are applied.

INVESTIGATION OF PARAFFIN-BASED FUELS IN HYBRID COMBUSTORS

Hybrid propulsion is a promising alternative for space propulsion; however, due to the low regression rate and hence low thrust levels of the classic polymeric fuels, new alternatives are sought. In recent years, liquefying, paraffin-based fuels have been investigated for their much higher regression rate compared to that of polymeric fuels. This paper presents experimental and theoretical research conducted at the Technion. Both plain paraffin fuel and paraffin-polymer mixtures have been tested. Plain paraffin exhibits a high regression rate but poor mechanical properties. The addition of polymer improves the mechanical properties but lowers the regression rate compared to plain paraffin fuel. Nevertheless, it indicates a promising direction to pursue a relatively high burning rate fuel of good mechanical properties. A model accounting for heat and mass transfer in the presence of liquefying fuel has been developed to predict the regression rate of the paraffin-based fuel. The model developed includes an additional mass loss mechanism, i.e., the liquid melt flowing along the grain. It is shown that this mechanism might have a noticeable effect on the overall regression rate and a potential for decreasing system performance. The test results exhibit good correlation with the model predictions.

Chemical Properties of Aircraft Engine Particulate Exhaust Emissions

The chemical properties of the particulate exhaust emissions from an in-use commercial aircraft engine were characterized in April 2004 as part of the Aircraft Particle Emissions Experiment. The test aircraft was the NASA DC-8 equipped with CFM56-2-C1 engines and the test matrix included 11 different engine throttle levels, three fuel compositions, and three sampling distances. The variations in particle emissions number, size, mass, and chemical composition were measured using a suite of instruments, including an aerosol mass spectrometer. The particle emissions were characterized by a trimodal size distribution. The largest mode was dominated by ambient accumulation mode particles mixed into the plume. The middle mode consisted of carbon soot with sulfate and organic coatings. The smallest mode was completely volatile and consisted of sulfate and organic components. The soot emission indices increased with power from 2―120 mg/kg fuel. The semivolatile components increased with distance and decreased with power from 33―5 mg/kg fuel. The sulfate emissions increased with distance and fuel sulfur content. The emissions under low power were dominated by organics, and the high-power conditions were dominated by soot. The CFM56 engine was less efficient at the low thrust levels typically used on the ground at an airport.

Investigation of Improved Hall Thruster Configurations for Low-Power Operation

<para xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink"> This paper deals with the performance degradation of Hall thruster, when it is operated at reduced power levels of 200–300 W. The standard approach to overcome the ionization efficiency problem by scaling down the thruster is limited by the magnetic material and circuit properties. In order to improve the thrust-measurement accuracy, a new calibration system was designed and built. Results indicate that, in the 200–350-W input power range, higher thrust, specific impulse, and efficiency are achieved. The effect of the magnetic-field distribution on the thruster performance was also examined and included a preliminary investigation of a novel idea of a new type of Hall thruster, which uses a reversed magnetic field near the anode. In order to implement a reversed magnetic-field configuration, a modified magnetic circuit was built and tested. Preliminary results demonstrated that indeed this configuration operated as a thruster, although at somewhat lower thrust levels. On the other hand, plume measurements indicate for a better focusing of the jet with this configuration. </para>