Bipropellant Thruster Research Articles

Modeling of Spray Combustion and Heat Transfer of MMH/N2O4 in a Small Rocket Engine Using Different Mechanisms

Although various hypergolic propellants like MMH/N2O4 (monomethylhydrazine/dinitrogen tetroxide) are widely used in small rocket engines, there remains a lack of in-depth study conducted on their chemical reactions and spray combustion behaviors. To fill this research gap, a simplified chemical kinetic model that is suitable for three-dimensional simulation was proposed in this paper for MMH/N2O4. Then, numerical investigation was conducted using the Volume of Fluid (VOF) model to explore the transient injection and atomization of MMH/N2O4 impinging jets in a small bipropellant thruster. Also, the instantaneous formation and evolution of the fan-shaped liquid film were analyzed. With the spray distribution determined, the proposed kinetic model and two existing mechanisms were applied to simulate spray combustion and heat transfer within the thruster, respectively, under the Euler–Lagrange framework. According to the research results, the liquid film covered nearly the entire chamber wall with a sawtooth pattern, which protected against the high temperatures of the engine wall. Notably, the two existing mechanisms showed significant errors in predicting temperature changes around the wall due to the excessively simplified reaction pathways. In contrast, the proposed model enabled the accurate prediction of the chamber pressure, wall temperature, and thrust with an error of less than 10%. Given the high accuracy achieved by the proposed numerical method, it provides a valuable reference for the development of advanced space engines.

Transient effects by pintle motion in cavitating venturi tube for throttled bipropellant thruster

An experimental study was conducted to investigate the behavior of an H2O2/kerosene bipropellant thruster during throttling. The thruster used 90 wt% hydrogen peroxide, and thrust was controlled using a variable area cavitating venturi valve equipped with a linear motor translating a pintle in a venturi geometry. In particular, transient effects were compared for three settings of pintle motion. Transient effects impacting pressure, temperature, thrust, efficiency, and instability data were analyzed. The results show that even at low chamber pressure, ignition occurred due to throttling, and no instabilities were created by changing the throttling speed. On the contrary, transient movements affected the recorded pressure data, impacting the characteristic velocity and thrust. Finally, this study provided new insight into H2O2/kerosene bipropellant throttling for deep throttling missions. Moreover, using the results from this study, the precision of future control algorithms will be improved.

Effect of swirl injection parameters on cooling performance of vortex-cooled bipropellant thruster

The cooling performance of a 10N-class gaseous Propane/GOx bi-propellant thruster with vortex cooling was investigated in this experimental study. Injecting oxidizer with a vortex was proposed as a potential cooling method for a semi-thermoplastic constructed thruster and demonstrated as a promising integrated cooling method that can be fabricated using 3D printing technology. Compared to other cooling types, vortex cooling has low system complexity and can be 3D printed as a single solid piece along with a thruster, reducing design complexity, weight, and manufacturing time. The experimental testing examined various injection configurations and their associated parameters to determine the most optimal injection combination. The test collected a comprehensive amount of data to evaluate the cooling performance with respect to varying operational parameters: oxygen-to-fuel ratio, propellant mass flow rate, and injection configuration. The study found that the acrylic chamber operated safely with a proper injection configuration identified maintaining the temperature of the thruster chamber walls below 60 °C, lower than the melting point of the acrylic material, with no damage observed. It was found that the downstream swirl injection had more favourable vortex cooling performance than the upstream injection for the scale of the thruster considered in this study.

Experimental investigation and numerical simulation of intense plume impingement effects from a 10 N bipropellant thruster

The intense interaction between thruster plumes and spacecraft surfaces presents substantial challenges for the advancement of space technologies. The impingement effects from a 10 N bipropellant thruster plume on a vertical instrumented plate were experimentally and numerically investigated in this study. Impingement pressure and heat flux were measured using micro differential pressure transducers and Gardon gauges, respectively, while the spatial distribution of the plume was captured using the high-speed camera. The phenomena and the distribution characteristics of plume impingement have been summarized, thereby providing a comprehensive validation dataset for subsequent predictions. Numerical simulations are conducted utilizing the coupled hybrid Computational Fluid Dynamics and Direct Simulation Monte Carlo (CFD-DSMC) algorithm. The dependence of the simulation outcomes on grid resolution was critically assessed, and the accuracy of the results was validated by comparison with experimental data. An integrated methodology, combining experimental and numerical approaches, was developed to calculate the maximum heat flux under realistic grid configurations. This method has proven to be effective and supports the accurate and efficient evaluation of intense plume impingement effects for engineering applications.

Optimal Liquid Engine Architecture by Performance-Cooling Tradeoff Analysis

We perform physics-based theoretical analyses on the tradeoff between the overall performance of specific impulse and the thermal protection ability of a coolant liquid film to identify the optimal configuration of a liquid rocket engine. A bipropellant thruster is set as the target, which uses a hypergolic propellant mixture of nitrogen tetroxide as the oxidant and monomethyl hydrazine as the fuel. By considering the practical nonuniform distribution of the local mixture ratio produced in the thrust chamber, the maximum specific impulse is achieved when the fuel and oxidizer spray widths become identical, allowing for the specification of the diameter of the impinging-type injector. The heat balance between the convective heat transfer from the combustion gas and the latent heat of the three-dimensionally wavy liquid film provides the optimal diameter of the combustion chamber. The longest liquid film is found to be achieved when half of the initial film is entrained by the fast gas, correspondingly mitigating the heat transfer area due to the liquid film waviness. We successfully demonstrate the optimal architecture of the liquid engine based on the tradeoff, in which an improvement of specific impulse by 1 s shortens the film length by 2 mm.

THEO & MUFN: Defending Earth against the 2023 PDC hypothetical asteroid impact

Planetary defense is an emerging field among the aerospace community. As the rate of near-Earth object (NEO) discoveries increases, so too does the probability that we will discover a NEO on a collision course with Earth. This paper documents a planetary defense mission campaign design to defend against the hypothetical Earth impactor 2023 PDC. At its 90th percentile mass estimate, the asteroid could have a diameter of approximately 1120 m. The asteroid makes no close approaches until Earth impact and has the potential to affect over a billion people, depending on its actual size. This paper describes a mission campaign consisting of both reconnaissance and mitigation missions that will be deployed to neutralize the threat posed by the asteroid. The reconnaissance orbiter named THEO (Terrestrial Hazard Exploration Orbiter) will survey the asteroid to clarify its morphological, gravitational, and dynamical properties. This information will be used to inform the design of the subsequent mitigation mission. THEO will be designed to have a mission lifetime of 12 years which will enable it to observe the mitigation mission and confirm its success. Two 21-day launch periods have been identified that minimize post-launch ΔV requirements and include current launch vehicle capabilities. For the purposes of this paper, the mitigation mission has been designed assuming a geocenter impact for the 90th percentile mass estimate of the asteroid because a precise impact location and an accurate mass estimate will not be available until THEO reconnoiters 2023 PDC. The mitigation mission, named MUFN (Mitigation Using a Fission Nuclear device), will send three spacecraft to rendezvous with 2023 PDC and alter the asteroid’s trajectory prior to Earth impact through a series of standoff nuclear detonations. The MUFN spacecraft are identical in design, each employing a hypergolic bi-propellant thruster. Each MUFN spacecraft launches on a SpaceX Falcon Heavy Expendable launch vehicle between 2028 and 2030 and carries four 340 kt nuclear explosive device (NED) equipped orbiters, named BELA-Os (“Bull’s Eye, Later Asteroid” Orbiters). The BELA-Os are designed to sequentially deploy, maneuver to an optimal standoff distance anti-parallel to the asteroid velocity vector, and detonate the NED to impart a change-in-velocity to the asteroid. After arrival at 2023 PDC, each MUFN spacecraft will station-keep above the asteroid. The NED-equipped BELA-Os will deploy from the MUFN spacecraft and maneuver 180 degrees out of phase from MUFN to the ‘front’ of the asteroid. Once there, each BELA-O will detonate and impart 2.32 mm/s of velocity change to the asteroid. After the ablation debris has cleared, MUFN will complete a single orbit around the asteroid to survey the blast site and assess the effectiveness of the detonation. After MUFN has completed its survey, it will release another BELA-O. This process will repeat until all BELA-Os have been released by all MUFN spacecraft. All denotations will occur between 2029 and 2031. After all BELA-Os have detonated their NED payloads near the 2023 PDC asteroid to deflect its orbit, the asteroid will safely fly by Earth at a minimum altitude of 3500 km instead of colliding on October 22, 2036.

Bipropellant Enabled Solid Oxide Fuel Cell Stack Power Supply: Energy Agility for Space Vehicles

Spacecraft often have instances where they are required to operate for long periods of time with significant power requirements and no access to solar power. While a battery system could be designed to meet these occasional, and possibly unexpected requirements, creating a power supply to meet these infrequent demands results in having a battery system that is excessively large, massive, and inefficient. A Solid Oxide Fuel Cell (SOFC) system capable of running on the same fuel and oxidizer as the onboard thruster of a bipropellant thruster equipped spacecraft can eliminate much of the added mass and volume needed to enable batteries to meet extreme power demands and create more operational flexibility on how the chemical energy in stowed bipropellants are used. If the spacecraft needs thrust, the bipropellant can be ignited to provide thrust. If the spacecraft needs electric power, the spacecraft can consume bipropellant to generate auxiliary power. In this work, the proven bipropellant pair ammonia (NH3) and nitrous oxide (N2O) are shown to be a viable fuel and oxidant for hydrogen (H2) consuming air breathing SOFCs, as spontaneous NH3 decomposition will generate H2 and spontaneous N2O decomposition will generate O2. A scale up to a 5-cell SOFC stack fed NH3 and N2O produced power up to ~89W. Higher power can be achieved by scaling up the number of cells in a stack to achieve required power demands. The results of this study demonstrate the potential and feasibility of integrating SOFCs with a bipropellant thruster to creating a dual functioning spacecraft system that can satisfy operational power requirements and on-orbit maneuverability.This material is based on research sponsored by Air Force Research Laboratory under agreement number FA9453-21-C-0060. The views expressed are those of the authors and do not reflect the official guidance or position of the United States Government, the Department of Defense or of the United States Air Force. Approved for public release; distribution is unlimited. Public Affairs release approval #AFRL20230023. Figure 1

Contamination assessment of a freely expanding green propellant thruster plume

A number of propellant/thruster combinations are under development in recent years that aim to replace the prevailing hydrazine-driven reaction control thrusters with less hazardous substances (“green propellants”). With some of these systems already in orbit, characterizing their contamination potential in a space environment becomes relevant. In this paper we discuss experiments on plume induced contamination from a novel propene/nitrous oxide bipropellant thruster, including high-speed imaging, SEM-EDS analysis, QCM measurements and in-situ mass spectrometry. Additional measurements of combustion chamber pressure complement the overall characterization of the thruster performance in a high vacuum environment. The main findings of this exploratory study are strong indications of solid and liquid phase particles being ejected from the nozzle, which will be investigated in a subsequent phase of the activity.

Bipropellant Enabled Solid Oxide Fuel Cell Stack Power Supply: Energy Agility for Space Vehicles

A Solid Oxide Fuel Cell (SOFC) system capable of running on the same fuel and oxidizer as an onboard bipropellant thruster can create enhanced operational flexibility. If a spacecraft needs thrust, the bipropellant can be ignited to provide thrust. If a spacecraft needs electric power, the spacecraft can consume bipropellant to generate auxiliary power. In this work, the proven bipropellant pair ammonia (NH3) and nitrous oxide (N2O) are shown to be a viable fuel and oxidant for SOFCs, as in situ thermal NH3 cracking will generate H2 and N2O decomposition will generate O2. A 5 cell SOFC stack produced 89 W when fueled with NH3 and N2O. The results of this study demonstrate the potential for a bifunctional spacecraft system that can satisfy both operational power and on-orbit maneuverability requirements.

Liquid film and heat transfer characteristics during superheated wall cooling via pulsed injection of a liquid jet

Pulse firing is one of the major operation modes of bipropellant thrusters for the attitude control of small-scale spacecrafts. In the pulse-firing mode, the operational range of the thruster depends on the deterioration of liquid film cooling performance. Because cooling performance is determined by the balance of heat removed by the liquid film and transferred to the manifold, the behavior and heat transfer characteristics of liquid films under the intermittent injection of liquid jets must be understood to enable a wider range of operational patterns. We conducted cooling tests on a superheated metal plate via the pulsed injection of a liquid jet for better understanding of the pulsed cooling process. Different injection patterns, including continuous injection, with the same injection quantity were examined on two types of metal plates (aluminum alloy and copper), because the thermal properties of metal plates affect both the heat transfer characteristics of the liquid film and temperature rise of the plate during the inter-pulse duration. The cooling process was evaluated based on the evolution of the liquid film on the metal plate and rear-side temperature of the metal plate. The evolution and temperature were simultaneously visualized using a high-speed camera and infrared camera, respectively. To link the liquid film state to the heat transfer process, the temperature and heat flux on the cooled surface were estimated by solving the inverse problem of three-dimensional transient heat conduction. The results indicate that the wetting front position corresponds to the position of the maximum temperature gradient. Additionally, the residual liquid film is consumed through the evaporation and the droplet dispersion related to the nucleate boiling. The effects of thermal inertia and diffusivity of the metal plate highlight the differences in the amount of heat removed by the liquid film. The heat removed by the liquid film during pulsed cooling exhibited a peak and was higher than that removed by continuous injection on the aluminum alloy plate. Less heat was removed under low-duty-cycle conditions, and the amount of heat removed by the liquid film was lower than that removed by continuous injection on the copper plate.

Catalyst bed behavior of hydrogen peroxide/kerosene bipropellant thruster in monopropellant and bipropellant modes with cavitating venturi valve

For typical hydrogen peroxide/kerosene bipropellant thruster applications, hydrogen peroxide is injected into a catalyst bed prior to an injection of kerosene. During this period, the bipropellant thruster operates in a monopropellant mode. For ignition, kerosene is injected into hot oxygen gas and water vapor, which are the decomposition products of hydrogen peroxide. After ignition, the bipropellant thruster operates in a bipropellant mode. During this ignition sequence, the catalyst bed switches from the monopropellant mode to the bipropellant mode. The pressure drop in the catalyst bed differs between these modes. It is necessary to understand why this difference exists. In this study, I examined the cause of the difference in the pressure drop of the catalyst bed by catalyst bed modeling. The catalyst bed model was tuned according to the results of experiments conducted using a 40 N monopropellant thruster. Subsequently, the model prediction and experimental results of a 330 N bipropellant thruster were compared. Using the model prediction, the cause of the difference in the catalyst bed pressure drop was examined. A possible cause of the pressure drop difference between the two modes is the difference in gas density owing to the injection pressure difference of hydrogen peroxide.

Simulation Study of the Swirl Spray Atomization of a Bipropellant Thruster under Low Temperature Conditions

The spray atomization of an injector significantly influences the performance and working life span of a bipropellant thruster of a spacecraft. Deep space exploration requires the thruster to be able to operate reliably at a low temperature range from −40 °C to 0 °C, so the effect of low temperature conditions on the atomization characteristics of injector spray is motivated to be comprehensively investigated. To study the swirl atomization characteristics of MMH (methylhydrazine), which is more difficult to atomize than NTO (nitrogen tetroxide), numerical simulations were conducted, employing the methods of VOF (volume of fluid) and LES (large eddy simulation) under low temperature conditions. The physical model with a nozzle size of 0.5 mm and boundary conditions with a velocity inlet of 3.89 m/s both follow the actual operation of thrusters. The development of spray atomization at low temperatures was observed through parametric comparisons, such as spray velocity, liquid total surface area, droplet particle size distribution, spray cone angle and breakup distance. When the temperature decreased from 20 °C to −40 °C at the same condition of flowrate inlet, those atomization characteristics of MMH propellant vary following these rules: the spray ejection velocity of MMH is significantly reduced by 7.7%, and gas-liquid disturbance sequentially decreases; the liquid film development is more stable, with a negative influence on atomization quality, causing difficulties for primary and secondary breakup, so the total surface area of droplets also decreases by 6.4%; the spatial distribution characteristics, spray cone angle and breakup distance vary less than 5%. Therefore, the low temperature condition can directly lower the combustion efficiency of thrusters with obvious performance degradation, but there are no significant changes in the propellant mixing and liquid film cooling. It is concluded that the bipropellant thruster can reliably work at low temperatures around −40 °C for deep space probe operation.

Analysis and Prediction Research for Bipropellant Thruster Mixture Ratio Based on BP-RNN Chain Method

In order to improve the mixture ratio accuracy of the bipropellant thruster and prolong the service life of a satellite on orbit, a variety of different liquid-flow tests and the final firing test are critical in the production testing process. However, after comparison with a large number of tests data, it was found that the liquid-flow test results were far away from matching the firing-test data without clear regularity, resulting in a qualified rate of mixture ratio of less than 40%. This study developed a BP-RNN (back propagation—recurrent neural network) chain method based on machine learning, which uses multi-dimensional nonlinear parameters to construct the specific dataset after data enhancement. Then, the mapping characteristic of the neural network was used to fit the historical data for the weight analysis and mixture ratio prediction, and effectively improved the qualified rate of the mixture ratio. The back propagation neural network was used to learn the association rules of the 10-dimension characteristic data and the firing test results generated in the historical process of thruster production. Then, the features with high influencing weight were extracted and sorted, so the “many-to-one” mixture ratio prediction was conducted through the subsequent recurrent neural network. The accepted prediction accuracy could reach around 75% after the test data verification. By using this method, most of bipropellant thrusters could directly reach the qualified mixture ratio in the firing test after adjusting the throttle orifice size in the liquid-flow tests. This chain method first bridges the data between the liquid-flow test and the firing test.

Unsteady Stream-Tube Model for Pulse Performance of Bipropellant Thrusters

We propose a theoretical model for predicting the impulse bit of bipropellant thrusters under pulse-firing operations. The present theoretical model, which considers the nonuniformity of the mixture ratios created inside the thrust chamber, extends the stream-tube approach, which is limited to the prediction of the steady performance. In pulse-firing operation, the fuel or the oxidizer alone can be injected in isolation due to the mismatched injection timing before the rated injection, which leads to the deterioration of performance as compared with the steady operation. The present approach (the unsteady stream-tube model) successfully implements a time-dependent stream-tube structure inside the thrust chamber, allowing for the prediction of the impulse bit as a straightforward function of the injection conditions. Three different pulse-firing tests using distinct hypergolic propellants demonstrate the validity of this model, typically reproducing the notable trend of deterioration in the impulse bit in the short-pulsed mode. We also examine the time-averaged specific impulse and mass flow rate to improve the impulse bit during short-pulsed operations.

Effect of Upstream Valve Opening Process on Dynamic Spray Atomization of Bipropellant Thruster Injector.

In order to develop a new generation of intelligent satellites, fast-response bipropellant thrusters are required to work in minimum impulse mode without limitation. When a valve is opening, the fluctuation affects downstream spray atomization at the injector, which determines the thruster’s impulse performance, involving combustion efficiency and impulse repeatability. Accordingly, the spray atomization under impulse working condition was investigated to optimize the thruster’s dynamic response. The effects of propellant property, switch speed, valve stroke, and throttle orifice layout are respectively compared in simulation cases using OpenFOAM. The fluctuating flowrate caused by valve opening was simulated and then used as boundary conditions for downstream spray. Among these factors, orifice layout plays the most significant roles in transient spray development. Compared with MMH spray, NTO spray from outer swirl injector is more sensitive to upstream fluctuation. When the upstream flowrate stabilizes faster, the atomization stability can also be enhanced, thereby improving the impulse repeatability of thrusters in combustion. This experimental result was in good agreement with the simulation, thereby showing that only when atomization of MMH spray and NTO spray both develop into a steady state within 5 ms after valve opening can the impulse performance be reliably achieved.

Evaporation of Three-Dimensional Wavy Liquid Film Entrained by Turbulent Gas Flow

A film cooling technique using a liquid film subjected to a hot gas flow in a turbulent condition is theoretically investigated. We successfully incorporate the two essential factors for the evaporating liquid film, droplet entrainment and three-dimensional film architecture, allowing for the physically consistent straightforward formulation. The validity of the present model is conveyed by reproducing combustion test results conducted for two types of flight-model bipropellant thrusters, in which the film length or dryout point shortens approximately inversely proportional to the combustion pressure. The underlying scenario to determine the film length is revealed. The gas flow initiates the originally smooth liquid film to be destabilized by Kelvin–Helmholtz instability in the axial direction and accelerates the wave crests leading to Rayleigh–Taylor instability in the transverse direction. Superposing the two types of waves produces three-dimensional cusps on the film as roots of entrained droplets. The convective heat transfer evaporating the liquid film is enhanced by the entrainment, reducing the net film flow rate, and by the cusp structure, enlarging the area of liquid/gas interface along the transverse direction in particular.

Energy Agility for Space Vehicle Applications Enabled by Bipropellant Consuming Solid Oxide Fuel Cells

The difficulty and cost of deploying space assets coupled with the variety of operations a space vehicle may be tasked to perform, the range of challenges a space vehicle may face, and the extended duration a space vehicle is often expected to operate without servicing make it desirable to create robust space vehicles that are capable of adapting to evolving circumstances or mission requirements. To help realize this vision, a research effort was begun to develop low profile Solid Oxide Fuel Cells (SOFCs) that can be used to generate electricity, on demand, by consuming fuel and oxidizer used in bipropellant thrusters for on-orbit maneuvering and station keeping. Developing a fuel cell to pair with a bipropellant thruster is of interest as such a combination can allow satellite operators to utilize the fuel and oxidizer stowed onboard a spacecraft equipped with a bipropellant thruster to generate electricity if the need arises. Moreover, creating such a fuel cell using solid oxide technology is of interest as the high operating temperature of SOFCs make them less susceptible to poisoning by contaminants which can be found in reactant flows, the nature of SOFCs make them easily customizable to meet specific electrical needs, and the large amount of energy which can be stowed in chemical storage tanks that can be subsequently converted into electricity by SOFCs can enable a spacecraft loaded with bipropellant chemicals to operate without tapping any other sources of energy for extended periods of time. Results from this work indicate that fuel cells assembled using commercial SOFC components designed to run on hydrogen (H2) and oxygen (O2) are often also effective at generating electricity by consuming fuels and oxidizers under consideration as green alternatives to the hazardous hydrazine (N2H4) fuel and nitrogen tetroxide (N2O4) oxidizer pair commonly used in on-orbit bipropellant thrusters. In this presentation, how fuel cells made with commercial components performed when fed different fuel and oxidizer combinations, each suitable for use in bipropellant thrusters, and the implications of these performances on efforts to create space deployable bipropellant driven SOFCs will be detailed.

Visualization of Coolant Liquid Film Dynamics in Hypergolic Bipropellant Thruster

We conduct a high-speed visualization of coolant liquid film dynamics inside a 10N-class bipropellant thruster using monomethylhydrazine and a mixture of nitrogen tetroxide with approximately 3% nitric oxide as propellants at equivalent conditions with the flight model. The direct visualization of the liquid film inside a quartz glass chamber demonstrates the transient dynamics of film flow from ignition to cutoff for the first time. The scenario for the life of the film flow is found as follows: the coolant fuel jet impinges on the chamber wall being the liquid film; the friction force from the chamber wall rapidly decelerates the film as one-tenth of the injection velocity; next, the film moves sheared by the fast combustion gas downstream; and eventually the film completely evaporates by heat transfer from the hot combustion gas. The liquid film presents the typical ripple wave structure, covered by the velocity and thermal boundary layers. The effect of the mixture ratio is significant as the film flow rate increases at a small mixture ratio, leading to the longer film length.

Effect of High-speed Solenoid Valve Dynamic Characteristics on Downstream Transient Spray Atomization of Bipropellant Thruster

Effect of High-speed Solenoid Valve Dynamic Characteristics on Downstream Transient Spray Atomization of Bipropellant Thruster

Direct Formulation of Bipropellant Thruster Performance for Quantitative Cold-Flow Diagnostic

We present a straightforward formulation predicting the characteristic velocity and specific impulse for bipropellant thrusters as a direct function of injection conditions, propellant combination, and nozzle expansion ratio. The theoretical formulation deduces a framework for a quantitative noncombustion test or cold-flow test to predict the performance indices. We simply employ water and dyed water as simulant liquids and then measure the local ratios of mixture and flow rate using an absorbance spectrometer. Density ratio mismatch between hypergolic propellants and water can be reasonably convertible. Combined with chemical equilibrium analysis, we obtain characteristic velocity and specific impulse across wide injection mixture ratios. The validity of the quantitative water-flow diagnostic is evidenced by comparing the results with those of corresponding combustion tests using nitrogen tetroxide and monomethylhydrazine as the propellants under several injector unlike-doublet and triplet configurations, showing that the mixing states of bipropellant thrusters under combustion can be reproduced using the water-flow diagnostic.