Hypergolic Propellants Research Articles

Trialkylsulfonium Thiocyanate Ionic Liquids: Investigation on Temperature‐Dependent Ignition Behavior of Green Hypergolic Propellants

AbstractThree trialkylsulfonium thiocyanate ionic liquids (ILs) are presented as fuel candidates for novel green hypergolic propellants. The physical and chemical properties such as density, viscosity, melting point and decomposition temperature as well as the enthalpy of formation of the ILs were determined. Further, the research focused on the potential of the ILs as fuel components for hypergolic propellants with highly concentrated hydrogen peroxide as oxidizer. For this, the theoretical maximum specific impulses of the propellants were calculated with NASA CEA. To evaluate the ignition delay times of the hypergolic propellants, drop tests were carried out. These were conducted not only at ambient conditions but also at lower levels of fuel temperature. It was found that all three propellant combinations still ignited reliably at temperature levels down to −25°C, however, the average ignition delay increased at lower temperatures. Overall, not only three promising hypergolic fuels were identified, but also important insights into the temperature‐dependent ignition behavior of hypergolic ionic liquid‐based propellants with thiocyanate anions were obtained.

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.

Analysis of unstable combustion caused by the Gas–Liquid two-phase flow in small hypergolic propellant engines

Hypergolic bipropellant engines are extensively used in spacecraft operations. In these engines, the combustion reaction is initiated by the impinging jets of the oxidizer and fuel in the liquid phase. The most commonly used hypergolic fuels, hydrazine and monomethylhydrazine, as well as the oxidizer, nitrogen tetroxide, are all in liquid state under room temperature and atmospheric pressure conditions. However, as these engines are operated in a space vacuum, a portion of the propellant inevitably evaporates because of low-pressure boiling immediately after engine start-up. The mixing of gas with liquid propellant results in unstable combustion, and minimizing this effect is critical for the engine design. Particularly, preventing high-frequency combustion instability is essential as it can be detrimental to the engine. This paper presents a mechanism that activates high-frequency combustion instability, realized by analyzing the pressure within the combustion chamber during artificially induced unstable combustion. Furthermore, the methodology for establishing operational constraints based on the proposed mechanism is clarified, along with the method for collecting test data. The study findings provide important indicators for the design criteria of bipropellant engines, contributing to the diversification and complexity of space development programs.

Selection of the method for determination of ignition delay of hypergolic propellants

Ignition delay is one of the most important parameters characterising hypergolic propellants. This parameter has a strong im-pact on thruster operation, especially during the cold start. Ignition delay influences the intensity of pressure rise and its peak values during the start of a thruster. High-pressure levels cause stress inside the chamber wall, which directly affects durability and safety. One of two measurement techniques is usually chosen to determine the ignition delay: visual and pressure-based methods. Visual methods are based on high-speed imaging and subsequent image analysis. In the pressure-based method, the pressure trace is analysed. In this study, both techniques were used together and compared in terms of ignition delay determination of hypergolic propellants igniting during the drop tests. The advantages and disadvantages of both techniques were indicated and described. In the setup used in the study, the visual method was found to be more accurate and reliable.

Numerical Study on the Design of an Anti-Backflow Injector for Combustion Chambers

The backflow of high-temperature products in an engine’s combustion chamber is a key issue which can significantly reduce combustion efficiency. This is particularly problematic for hypergolic propellants, as the high-temperature products may still contain fuel or an oxidizer. If either the fuel or the oxidizer backflows into the manifold of the other, it can easily lead to micro-explosions, thereby creating a threat. To address this problem, this paper proposes a new design of an anti-backflow injector, aimed at effectively preventing the backflow of combustion products to the propellant manifold. A steady-state, non-reactive Computational Fluid Dynamics (CFD) model is employed to evaluate the steady internal flow characteristics of the proposed anti-backflow injector. Additionally, a complementary transient, multiphase fluid dynamics simulation is carried out to assess the response characteristics of the anti-backflow injector. Our analysis focuses on the response characteristics of the concave valve core. The study also explores the impact of different expansion port angles on the injection effect, finding that the vortex diameter at the injector outlet is positively related to the expansion port angle. It is also shown that the injection angle becomes more stable as the expansion angle increases. When the expansion port angles are 10° and 20°, the injection angles show similar trends. In terms of anti-backflow effect, taking the injection stiffness of 150% and a time range of 300 µs as examples, the response time of the anti-backflow models with expansion ports of 10°, 15°, and 20° is increased by 67 µs, 99 µs, and 213 µs, respectively, compared to the base models with the same expansion angles. Meanwhile, when the injection stiffness is 50%, the response time of the anti-backflow models with expansion ports of 10°, 15°, and 20° is increased by 207 µs, 210 µs, and 207 µs, respectively. When the injection stiffness is 20%, the response recovery speed of the anti-backflow models with expansion ports of 10° and 15° is increased by 41 µs and 96 µs, respectively. However, the performance of the anti-backflow model with a 20° expansion port is 216 µs of the base model. The optimized design of the anti-backflow injector has potential applications in solving the propellant backflow problem and contributes to the advancement of combustors.

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.

Droplet collision of hypergolic propellants

AbstractIn the present mini‐review, droplet impacting on a liquid pool, jet impingement, and binary droplet collision of nonreacting liquids are first summarized in terms of basic phenomena and the corresponding nondimensional parameters. Then, two representative hypergolic bipropellant systems, a hypergolic fuel of N,N,N′,N′‐tetramethylethylenediamine (TMEDA) and an oxidizer of white fuming nitric acid (WFNA) and a monoethanolamine‐based fuel (MEA‐NaBH4) and a high‐density hydrogen peroxide (H2O2), are discussed in detail to unveil the rich underlying physics such as liquid‐phase reaction, heat transfer, phase change, and gas‐phase reaction. This review focuses on quantifying and interpreting the parametric dependence of the gas‐phase ignition induced by droplet collision of liquid hypergolic propellants. The advances in droplet collision of hypergolic propellants are important for modeling the real hypergolic impinging‐jet (spray) combustion and for the design optimization of orbit‐maneuver rocket engines.

Engineering High-Performance Hypergolic Propellant by Synergistic Contribution of Metal-Organic Framework Shell and Aluminum Core.

Hypergolicity is a highly desired characteristic for hybrid rocket engine-based fuels because it eliminates the need for a separate ignition system. Introducing hypergolic additives into conventional fuels through physical mixing is a feasible approach, but achieving highly reliable hypergolic ignition and energy release remains a major challenge. Here, the construction of core-shell Al@metal organic framework (MOF) heterostructures is reported as high-performance solid hypergolic propellants. Upon contact with the liquid oxidizer the uniformly distributed hypergolic MOF (Ag-MOF) shell can induce the ignition of hypergolic-inert fuel Al, resulting in Al combustion. Such a synthetic strategy is demonstrated to be favorable in hotspot generation and heat transfer relative to a simple physical mixture of Al/Ag-MOF, thus producing shorter ignition delay times and more efficient combustion. Thermal reactivity study indicated that the functionalization of the Ag-MOF shell changes the energy release process of the inner Al, which is accompanied by a thermite reaction. The synergistic effect of implantation of hypergolic MOF and high energy Al contributes to high specific impulses of 230-270 s over a wide range of oxidizer-to-fuel ratios.

Symmetry-Breaking-Induced Internal Mixing Enhancement of Droplet Collision

Binary droplet collision is a basic fluid phenomenon for many spray processes in nature and industry involving lots of discrete droplets. It exists an inherent mirror symmetry between two colliding droplets. For specific cases of the collision between two identical droplets, the head-on collision and the off-center collision, respectively, show the axisymmetric and rotational symmetry characteristics, which is useful for the simplification of droplet collision modeling. However, for more general cases of the collision between two droplets involving the disparities of size ratio, surface tension, viscosity, and self-spin motions, the axisymmetric and rotational symmetry droplet deformation and inner flow tend to be broken, leading to many distinct phenomena that cannot occur for the collision between two identical droplets owing to the mirror symmetry. This review focused on interpreting the asymmetric droplet deformation and the collision-induced internal mixing that was affected by those symmetry breaking factors, such as size ratio effects, Marangoni Effects, non-Newtonian effects, and droplet self-spin motion. It helps to understand the droplet internal mixing for hypergolic propellants in the rocket engineering and microscale droplet reactors in the biological engineering, and the modeling of droplet collision in real combustion spray processes.

Hypergolic Propellants Based on High-Test Hydrogen Peroxide and Organic Compounds

Hypergolic Propellants Based on High-Test Hydrogen Peroxide and Organic Compounds

Hypergolic ignition response to oxidizer droplet properties

The hypergolic ignition response to oxidizer droplet properties is an important consideration for oxidizer injector design in hybrid rocket motors. Hence, the present work focused on hypergolic ignition characterization as a function of droplet properties and impact dynamics for a promising propellant. The propellant for the present work consisted of rocket grade hydrogen peroxide (RGHP) as oxidizer, high-density polyethylene (HDPE) as fuel and sodium borohydride (NaBH4) as the hypergolic additive embedded in the HDPE matrix. This propellant was chosen for this detailed study since it showed promising results in prior studies, exhibiting low (<10 ms) ignition delays. For this work, in a drop-test setup, RGHP droplet diameters between 1.90±0.11 mm and 4.35±0.26 mm were dropped on the HDPE + NaBH4 fuel samples at velocities between 0.3 m/s to 1.93 m/s. The ensuing ignition event was studied using both visible and infrared high-speed imaging. Post-impact droplet parameters were determined to enable a deeper analysis of observed phenomena. While ignition delay was the focus of this work, flame spread areas and heat release were also studied. Higher impact velocities generally led to lower ignition delays while splashing oxidizer drops increased probability of larger flame areas and higher heat release. Additionally, the initial ignition/heat release event was found to originate in a small region at the thinnest area of the post-impact droplet. Surface temperature measurements were also performed using infrared imaging. The data suggested that the initial heat release was found to be likely due to RGHP-NaBH4 reaction with no contribution from RGHP decomposition. Regression and random forest machine learning analysis found ignition delay depended strongly on the minimum post-impact drop layer thickness. Based on the data analysis, a simple mathematical model was developed whose results matched experimentally observed trends. The study is expected to further understanding of the processes that control hypergolic ignition in addition to aiding rocket engineers design better oxidizer feed systems. Novelty and SignificanceThis work constitutes a novel in-depth study on the influence of drop impact properties on hypergolic ignition. Using innovative techniques, this work is the first to show the overwhelming importance of residual droplet layer thickness on ignition delay. Additionally, a simple mathematical model based on that finding performed well against experimental data which is of substantial significance, given the complex and multi-parametric nature of hypergolic ignition. Finally, the work establishes that operating in the splashing regime enhances heat release and flame spread. The candidate propellant for this work was hydrogen peroxide and high-density polyethylene with embedded NaBH4. However, the findings of this study are likely generalizable over a range of solid-liquid hypergolic propellants involving solid fuels with embedded energetic materials, of which there are many. The findings of this work can enhance understanding of hypergolic ignition in addition to aiding in better rocket injection system design.

Experimental research on the detonation behavior in annular combustors utilizing liquid hypergolic propellants

The development of rotating detonation engine utilizing hypergolic propellants is necessary for practical application of detonation-based rocket engines, but this liquid-liquid rotating detonation has been rarely researched due to its difficulty of well atomization and mixing in very short time. This study aims to experimentally investigate the propagation characteristics of rotating detonations using non-premixed hypergolic propellants in annular combustors. Here the monomethyl hydrazine (MMH) and nitrogen tetroxide (NTO) have been used as oxidizer and fuel, respectively. Twenty-four unlike impinging elements have been used for injection and mixing of liquid propellants. The orifice diameter is 0.4 mm for oxidizer and 0.3 mm for fuel. Three typical combustion modes such as deflagration mode, one-wave rotating detonation mode and two-wave collision mode have been captured by dynamic pressure transducers and the high-speed camera. The influence of many factors, such as mass flow rate, mixing ratio, annular combustor width and chamber throat, on rotating detonation behaviors have been investigated. The propagation velocity of rotating detonations is around 1328 m/s∼1405 m/s, including one-wave rotating detonation mode and two-wave collision mode. When the mass flow rate and mixing ratio is not appropriate, the hypergolic combustion becomes obviously unstable and combustion modes vary very quickly, especially at the start-up stage. The chamber throat plays a negative role on rotating detonation which suppresses the detonation propagation and can lead to detonation fail. Besides, the wider annular combustor width is beneficial for rotating detonation propagating, which can eliminate the negative influence of the chamber throat on the detonation propagation. In addition, increasing mass flow rate would lead to higher injection velocity and thus better impinging injection, atomization and mixing, which is positive for increasing the stable operation range of mixing ratio for rotating detonations.

ATOMIZATION AND MIXING OF BLENDED ENERGETIC IONIC LIQUID HYDROXYETHYLHYDRAZINIUM NITRATE

Hypergolic propellants are widely used in liquid rocket propulsion. Instantaneous control of combustion, lack of requirement of an ignition system, and storability for long durations are the major advantages of hypergolic propellants. The propellant community is on the lookout for replacements for the currently utilized toxic hydrazine and its derivatives. Energetic hypergolic ionic liquids are excellent candidates for such replacement. However, their high viscosity and surface tension render atomization difficult, and hence, increase the ignition delays in the combustion chamber and reduce combustion efficiency. The current study focuses on the geometrical aspects of injector design and other injection conditions that control the ignition process and combustion efficiency. These mainly depend on the extent of mixing and atomization. Shadowgraphy and patternator-based experiments were performed on doublet and triplet impingement injectors under variable injection conditions, such as impingement angle, jet diameters, and injection velocities. A blend of 50&amp;#37; unsymmetrical dimethylhydrazine in the energetic ionic liquid hydroxyethylhydrazinium nitrate (UHN50) shows promise as a candidate for a hydrazine replacement. A blend of 38&amp;#37; ethanol in glycerol (GE6238) was used as a surrogate for UHN50, whereas water was used as a surrogate for nitrogen tetroxide. The effect of variation of injection conditions was found to be contradictory for achieving atomization and mixing. An optimum criterion was obtained for both atomization and mixing.

Hypergolic ignition behaviors of green propellants with hydrogen peroxide: The TMEDA/DMEA system

For nearly-six decades, hydrazine and its derivatives have been the standard fuels for rockets and spacecrafts. However, their attractive features are offset by their toxicity and associated high handling and storage costs. This work presents a novel green fuel system based on N,N,N’N’-Tetramethylethylenediamine (TMEDA), dimethylaminoethanol (DMEA) and methanol or ethanol with high test peroxide (HTP). Drop tests were performed with 27 samples using both high speed photography and transient infrared imaging. Results have shown that neither pure TMEDA nor pure DMEA is hypergolic with high test peroxide (HTP) even with copper chloride (anhydrous or hydrated) as the catalyst. Interestingly, their combination provided a synergistic effect achieving hypergolic behavior with a consistently low ignition delay time (IDT) using only half a percent of catalyst, and the near-optimal hypergolic ignition behavior was achieved with a volume ratio of TMEDA to DMEA of 50:50. In addition, in order to adjust some properties of this blend, TMEDA/DMEA (50:50) was further mixed with methanol or ethanol with different ratios and they showed even better hypergolic ignition performance with IDT as low as 10 ms. Besides their good hypergolic ignition performance characteristics, these new fuel systems based on two propagators (TMEDA/DMEA) and a solvent (methanol or ethanol) also present low viscosity and comparable theoretical specific impulses compared to the conventional hydrazine-based systems. It is believed that this catalytically promoted hypergolic systems with HTP open up a new avenue to the replacement of conventional highly toxic hypergolic propellants.

Influence of Flow Rate Distribution on Combustion Instability of Hypergolic Propellant

Combustion instability is the biggest threat to the reliability of liquid rocket engines, whose prediction and suppression are of great significance for engineering applications. To predict the stability of a combustion chamber with a hypergolic propellant, this work used the method of decoupling unsteady combustion and acoustic system. The turbulence is described by the Reynolds-averaged Navier–Stokes technique, and the interaction of turbulence and chemistry interaction is described by the eddy-dissipation model. By extracting the flame transfer function of the combustion field, the eigenvalues of each acoustic mode were obtained by solving the Helmholtz equation, thereby predicting the combustion stability for the combustion chamber. By predictions of the combustion chamber instability with different flow rate distributions, it was found that the increasing of inlet flow rate amplitude will improve the stability or instability of combustion. The combustion stability of the chamber was optimized when the flow rate distribution for the oxidant was set more uniform in the radial direction. The heterogeneity of the flow rate distribution in the circumferential direction is not recommended, considering that a homogeneous flow rate distribution in the circumferential direction is beneficial to the combustion stability of the chamber.

Comprehensive ignition characterization of a non-toxic hypergolic hybrid rocket propellant

This paper describes a comprehensive characterization of ignition properties of a metal-hydride based non-toxic hypergolic hybrid rocket propellant. The propellant consists of Rocket Grade Hydrogen Peroxide (RGHP) as oxidizer, high-density Polyethylene (HDPE) as fuel and sodium borohydride (NaBH4) as the additive, embedded in the HDPE matrix. Ignition quality was characterized as ignition delay, ignition probability and flame spread. In a drop-test setup, ignition characteristics were determined as a function of seven parameters: RGHP concentration, additive loading, oxidizer droplet impact velocity, oxidizer droplet volume, pressure, diluent gas, and environmental exposure. The parameters encompass thermo-chemical, fluid/droplet dynamics and environmental factors affecting ignition. Ignition delays as low as 3 ms were observed, one of the lowest using non-toxic hypergolic hybrid propellants in open-air. An overwhelming majority of conditions tested yielded <10 ms ignition delays and 100% ignition success. All conditions tested affected ignition to varying degrees with RGHP concentration, NaBH4 loading and drop impact velocity significantly affecting ignition. Further, contrary to expectations, exposing sanded fuel samples to humidity for a few h enhanced ignition instead of hampering it and exposure for 24 h did not lead to ignition degradation. Tests with diluent gases other than air (at atmospheric and elevated pressures) revealed that atmospheric oxygen played a negligible role in the reaction process. This proved that oxygen for the initial ignition event was obtained from RGHP decomposition, with atmospheric oxygen playing no role in ignition performance. Aside from demonstrating excellent ignition characteristics, our results further show a need to go beyond thermo-chemical properties and to consider aspects of ignition other than ignition delay in hypergolic propellant research to enable a complete understanding of the ignition processes. The comprehensive ignition characterization demonstrates the chosen propellant's ability to overcome ignition challenges in hybrid rockets and serves as a proof of concept for its further development.

Reactivity of hypergolic hybrid solid fuel with industrial grade hydrogen peroxide

Hypergolic solid fuels and hydrogen peroxide oxidizers have attracted considerable attention for the development of safe hybrid rocket propulsion systems. In this study, a new approach was explored to develop a safe hypergolic hybrid propulsion system using industrial-grade hydrogen peroxide (∼70 wt%). Initially, the propulsion performance of different hypergolic solid fuels with an industrial-grade hydrogen peroxide was calculated using the NASA CEA code to find an appropriate propellant combination. The theoretical specific impulse of the hypergolic solid fuel was 27.8 % over hydrazine and 13.4 % over LMP-103S. Based on theoretical investigations, hypergolic solid fuels, named such as HSF-1 to HSF-10, were fabricated using molding and pressing methods. The ignition delay time was measured using the drop-test method. Interestingly, HSF-3 and HSF-4 exhibited very short ignition delay times of 4.92 and 8.75 ms, respectively, with industrial-grade hydrogen peroxide. In addition, the drop test results of hypergolic solid fuel with varying compression pressure in the pressing method confirmed that a pressure above the compressive strength of the binder shortens the ignition delay times deviation. Further, HSF was also ignited even at low concentration of IGHP (>40 wt%); opens the new area of research in combustion science. Overall, low-toxicity hypergolic solid fuels and industrial-grade hydrogen peroxide are promising alternatives to conventional toxic hypergolic propellants.

Effects of oxidizing additives on the physical properties and ignition performance of hydrogen peroxide-based hypergolic propellants

Herein, the concept of novel hypergolic pairs for the enhanced ignition performance of hypergolic propellants is proposed. The proposed hypergolic combinations comprised ionic liquid fuels, namely, 1-ethyl-3-methyl-imidazolium-thiocyanate (EMIM SCN) and 1-butyl-3-methyl-imidazolium-thiocyanate (BMIM SCN), and hydrogen peroxide (H2O2) with oxidizing additives. LiNO3 and NH4NO3 were used as oxidizing additives, which were dissolved in 60–95 wt% H2O2 to enhance the physical properties and ignition performance. Adding the oxidizing additives decreased the theoretical performance of the hypergolic pairs, however, the presence of oxidizing additives drastically reduced the freezing point of the mixture with 95 wt% H2O2. Particularly, an abrupt change was observed in the freezing point of the mixture with the LiNO3 additive. At only 5 wt% LiNO3, the freezing point reached −30 °C. The ignition performance of the novel hypergolic pairs was significantly enhanced by the introduction of oxidizing additives. In the drop test, the LiNO3 additive exerted a substantially greater effect on the enhancement of ignition performance than that exerted by NH4NO3. Furthermore, this positive contribution was maximized with the decrease in the hydrogen peroxide concentration. Although the addition of NH4NO3 to 90 and 95 wt% H2O2 adversely affected the ignition performance, the combination of BMIM SCN and 60 wt% H2O2–containing LiNO3 extended the hypergolicity limit. This study is the first attempt to introduce nitrate salts as oxidizing additives in hydrogen peroxide with ionic fuels, including SCN anions. Nitrate salts exhibit potential to serve as additives that can be used with hydrogen peroxide to enhance the properties of oxidizers. The nitrate salt usage helps lower the freezing point of oxidizers. Further, it shortens the ignition delay time of the hypergolic pairs.

Detonation cell size of liquid hypergolic propellants: Estimation from a non-premixed combustor

An experimental approach for estimating the detonation cell size for liquid hypergolic propellants is presented and applied for monomethylhydrazine (MMH) and a MON-3 variant of nitrogen tetroxide (NTO) in a non-premixed combustor. The method utilizes a correlation between cell size and reactant fill height in an annular combustor geometry. Reactant fill height is inferred using a control-volume analysis to relate the geometry of the reactant fill zone to the propellant flow rate, detonation wave-speeds and number, and annular gap size. High-speed videography is used to measure the number of waves and their speed over a range of quasi-steady continuous detonation conditions. The detonation criteria is also proven valid in the transient conditions as the decrease of the fill height below a critical value identified in this work matches modal transitions with decreasing number of waves. A power law relating the cell width with induction length further supports the validity of the present technique down to cell sizes of 1–10 µm, a scale not practically resolvable with conventional methods.

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.