Hybrid Rocket Engine Research Articles

하이브리드 로켓엔진 액체산소 재생냉각 노즐 연구 - Part 1

하이브리드 로켓엔진 액체산소 재생냉각 노즐 연구 - Part 1

Bridging the Technology Gap: Strategies for Hybrid Rocket Engines

Hybrid rocket propulsion, first demonstrated by the Russian GIRD-09 rocket in 1933, combines liquid oxidizer and solid fuel for thrust generation. Despite numerous advantages, such as enhanced safety, controllability, and potential environmental benefits, hybrid propulsion has yet to achieve its full potential in space applications. In recent years, the research on hybrid propulsion has gained enormous momentum in both academia and industry. Recent accomplishments such as the altitude record for student rockets (64 km), the launch of the first electric pump-fed hybrid rocket, and a successful 25 s hovering test highlight the potential of hybrid rockets. However, although the hybrid community is growing constantly, industrial utilizations and in-space validations do not yet exist. In this work, we reassess the possibilities of hybrid rocket engines by presenting potential fields of applications from the literature. Most importantly, we identify the technical challenges that hinder the breakthrough of hybrid propulsion in the space sector and evaluate the technologies and approaches necessary to bridge the gaps in hybrid rocket development.

Ignition delay and low-frequency coupling in hybrid rocket combustion

Hybrid rocket engines (HREs) require a post chamber to allow additional enthalpy production of unburned mixture, which may be expelled without participating in combustion. Many studies have confirmed that the post chamber is responsible for generating low-frequency instability (LFI). This study performed various tests with increasing post-chamber lengths. The test results show that post-chamber length is an important factor that determines the occurrence of the LFI. While the LFI persists, there is an observed positive coupling between pressure oscillations (p′) and heat release oscillations (q′), which share the same high-frequency band (400–500 Hz). But a phase difference changes periodically over time, it forms pressure beats of approximately 15–20 Hz. Video images (with LFI) show that combustion flows bounce back upstream after impinging on the nozzle wall. The bounced-back flow then collides with the incoming flow and forms a counter flow in the post chamber. In addition, re-ignitions with delays are observed in the counter-flow region, which is discontinuous and periodic. This study revealed that the periodic positive coupling of p′ and q′ is one of the manifestations of thermoacoustic instability. A modified kick oscillator model is used to account for the ignition delay and successfully predicted the pressure beats.

3D printed different polymer fuel grains for hybrid rocket engine

The advent of various additive manufacturing technologies, such as 3D printing, has changed the structural design and preparation process of rocket fuels. In order to investigate the combustion properties of various common polymer materials that can be additively manufactured by fused deposition, the combustion tests of hybrid rocket fuels prepared by 3D printing were carried out. These materials include polylactic acid (PLA), wood-like polylactic acid (Wood), acrylonitrile-butadiene-styrene (ABS), acrylonitrile-styrene-acrylate (ASA), copolymers of nylon 6 and nylon 6,6 (CoPA), polycarbonate-polybutylene terephthalate (PC-PBT), flame retardant polycarbonate (PC-FR) and polyethylene terephthalateco-1,4-cylclohexylenedimethylene terephthalate (PETG). Thermogravimetric-differential scanning calorimetry (TG-DSC) analysis was carried out on these materials, the printing effect of the fuel grains was observed by three-dimensional X-ray microscopy (µCT) and the combustion performance of these fuels in gaseous oxygen flow (GOX) was recorded by high-speed photography at a constant pressure of 1 MPa. The results show that ASA and ABS exhibit good printing results. The regression rates of PC-PBT, PETG, ABS, ASA, CoPA and PLA are 0.792 mm/s, 0.592 mm/s, 0.536 mm/s and 0.477 mm/s, 0.368 mm/s, 0.339 mm/s (AtGOX=220kg/(m2·s)), respectively.

Testing of the N2O/HDPE Vortex Flow Pancake Hybrid Rocket Engine with Augmented Spark Igniter

The paper is part of the research aimed at determining if the vortex flow pancake (VFP) hybrid rocket engine is feasible as green in-space chemical propulsion. The objective of this study is to test an N2O/HDPE VFP hybrid ignited with N2O/C3H8 torch igniter. The N2O is used in self-pressurizing mode, which results in two-phase flow and varying inlet conditions, thus better simulating real in-space behavior. The study begins with characterizing the torch igniter, followed by hot-fire ignition tests of the VFP. The results allow for the improved design of the torch igniter and VFP hybrid. The axial regression rate ballistic coefficients are reported for the N2O/HDPE propellants in the VFP configuration.

Hybrid Rocket Engine Noise: Measurements and Predictions of Acoustic Environments from Horizontal Static Fire

A rocket’s turbulent jet radiates intense acoustic waves, which are an acoustic load for structural components like payload, launch structure, and rocket avionics, and impact communities near the launch site. Therefore, a careful characterization of the acoustic field produced by a rocket engine can provide crucial information during the design phase. In particular, this deals with improving the understanding of the acoustics of low-thrust hybrid rocket engines. Since an accurate jet noise detection around the entire launch site is time-consuming and extremely cost-prohibitive, a fast and reliable predictive tool is invaluable. For this purpose, a semi-empirical model was employed, using the exhaust plume property and geometric characteristics of the nozzle as input. Experimental data collected during a firing test campaign, conducted in the framework of HYPROB-NEW project by the Italian Aerospace Research Center, were decisive to discuss the validity of the model also for low-thrust hybrid propulsion and support the goodness of the noise curves and metrics estimated for nearby regions and provide considerations about the implications of engine geometric characteristics on noise emissions.

An HTS Module for Magnetic Combustion Control in a Lab-Scale Hybrid Rocket Engine

Magnetohydrodynamic control of diffusive flames can provide a contactless method for altering combustion in hybrid rocket engines. The presence of a magnetic field modifies conditions within the thermal boundary layer and provides an opportunity for active control of fuel release. However, the realization of the process requires the application of substantial magnetic fields, beyond the capability of copper electromagnets. This work presents a feasibility study of an HTS magnet designed for a lab-scale technology demonstration. The presented magnetohydrodynamic combustion model shows that a magnetic field in the range of 1 T is required to enhance the conditions in the combustion chamber. The realization of cooling to sustain the superconducting state is assumed to be achieved through the utilization of the existing cryogenic environment needed for the liquid oxygen in combustion. The numerical analysis shows that the required magnetic field can be delivered with a mid-sized racetrack coil magnet made of two double pancake coils operating at 65 K. The coil arrangement is shown to be able to withstand the heat generated by the combustion of a 10 s burn. For the lab-scale application, a ferromagnetic core will be used to reduce the amount of superconducting tape required, while retaining a suitable magnetic field. The analysis is concluded with the scaling studies for an HTS coil without an iron core, which shows that a magnetic field of up to 2 T can be achieved.

Numerical modeling of hybrid rocket engine

Numerical modeling of hybrid rocket engine

Numerical and experimental analysis of fuel regression rate in a lab-scale hybrid rocket engine with swirl injection

In this work the regression rate performance and flow physics of a lab-scale hybrid rocket engine burning gaseous oxygen and paraffin-based fuels are experimentally and numerically investigated. Regression rates are obtained by thickness-over-time averaging procedures and through a non-intrusive optical method enabling fuel grain port diameter tracking. A numerical rebuilding of the experimental data is performed with axisymmetric Reynolds-averaged Navier-Stokes simulations, using sub-models accounting for the effects of turbulence, chemistry, radiation, and fluid-surface interaction. Simulations are performed with different computational setups, also considering the fuel grain shape variation over time, obtaining a fairly good agreement between the numerical and experimental data. A parametric analysis is also performed to assess the variation of the fuel regression rate with swirl intensity.

Test Activities on Hybrid Rocket Engines: Combustion Analyses and Green Storable Oxidizers—A Short Review

Hybrid rocket engines (HREs) offer a low-cost, reliable, and environmentally friendly solution for both launch and in-space applications. Hybrid propellants have been identified as green thanks to their use of non-toxic, non-carcinogenic oxidizers. Of particular relevance are storable oxidizers, namely high-concentration (≥90 wt.%) hydrogen peroxide (HP, H2O2) and nitrous oxide (N2O). This work provides a survey of experimental activities based on H2O2 and N2O for hybrid rocket propulsion applications. Open literature data are completed with original thermochemical calculations to support the discussion.

Preliminary sizing and study of a hybrid rocket based combined cycle

This study investigates the feasibility and performance of a missile propelled by a Hybrid Rocket-Based Combined Cycle (HRBCC) engine, which combines hybrid rocket and air-breathing engines. The HRBCC engine could enable intra and extra-atmospheric hypersonic flight in the future. Using a hybrid rocket engine could bring several advantages as reduced costs, increased reliability and simplicity. This research presents the unique feature of the hybrid rocket engine developed by HyPrSpace and explains its benefits for the conception of a HRBCC system. The performance of the missile was studied in two operational modes: pure rocket and ramjet. The results were then used to develop an in-house code that compared the trajectory of different missiles. Our findings indicate that HRBCC engines provide several advantages over other propulsive technologies and can meet the requirements of various military applications.

Experimental Firing Test Campaign and Nozzle Heat Transfer Reconstruction in a 200 N Hybrid Rocket Engine with Different Paraffin-Based Fuel Grain Lengths

Firing test campaigns were carried out on a 200 N thrust-class hybrid rocket engine, using gaseous oxygen as an oxidizer and a paraffin-wax-based fuel. Different fuel grain lengths were adopted to extend the fuel characterization under different operating conditions, and to evaluate rocket performances and internal ballistics in the different configurations. In addition to data collected under a 220 mm propellant grain length, two further test campaigns were carried out considering 130 mm and 70 mm grain lengths. Two different injector types were adopted in the 130 mm configuration; in particular, a showerhead injection system was used with the aim to contain high-amplitude pressure oscillations observed during some firing tests in this engine configuration. Parameters such as the chamber pressure and temperature inside the graphite nozzle, space-averaged fuel regression rate and nozzle throat diameter were measured. The results allowed for the investigation of different issues related to hybrid rockets (e.g., fuel regression rate, engine performance, nozzle ablation under different conditions). The focus was mainly directed to the nozzle heat transfer, through the reconstruction of the convective heat transfer coefficient for different tests in the 70 mm grain length engine configuration. The reconstruction took advantage of the experimental data provided by the nozzle embedded thermocouple. Then, the experimental convective heat transfer coefficient was used to validate the results from some empirical correlations. The results showed significant differences between the experimental convective heat transfer coefficients when considering tests with different oxidizer mass flow rates. Furthermore, the predictions from the empirical correlations proved to be more reliable only in cases characterized by oxidizer-rich conditions.

Numerical study of the effect of fuel droplets on the combustion proceeding under typical conditions of hybrid rocket engines

Numerical study of the effect of fuel droplets on the combustion proceeding under typical conditions of hybrid rocket engines

Data acquisition unit for suborbital rocket vibration measurement

Suborbital rockets and space launch vehicles experience high levels of vibration. They are caused by various factors, such as aerodynamic effects on the external surfaces and engine combustion instability. The second factor is especially pronounced in hybrid rocket engines, such as the main engine of the ILR-33 AMBER 2K rocket. In order to ensure the correct operation of all the equipment and payload on board, vibration testing of the devices is necessary as part of pre-flight qualification. For such tests to be truly indicative of in-flight performance, the vibration environment of the vehicle must be analysed. A conceptual measurement method was proposed, along with a prototype measurement system. The system is adapted to be integrated into the modular ILR-33 AMBER 2K on-board computer or work independently as a separate module. The device measures vibration in three axes in several key points onboard the rocket, with the possibility of connecting multiple units in a daisy-chain configuration. This paper presents the design of the module, including the key aspects of integrating it as part of the avionics package. The means of acquiring and storing of measurement data are presented. The method of validating the system’s data acquisition capabilities is described in detail. Qualification testing results and identified issues are discussed. Possible future improvements are proposed.

Performance of Additively Manufactured Fuels for Hybrid Rockets

Hybrid rocket engine (HRE) performance is dependent on fuel/oxidizer selection and fuel grain geometry. A literature review was performed to identify key trends and findings related to the application of the additive manufacturing (AM) of fuel systems for HREs. The effects of complex combustion port geometries, embedded structures, and end-burning systems, along with the use of metallic additives, turbulators, diaphragms, gel-like fuels, powdered fuels, liquid fuels, and liquifying fuels and their impact on regression rates, combustion efficiencies, and/or mechanical strength are thoroughly documented here. In general, the application of AM to HRE fuels can be implemented to increase regression rates and combustion efficiency, and tailor HRE designs. Chemical equilibrium analysis computations were completed to characterize the theoretical performance of HTPB and common AM fuels (ABS, PLA, PC, PMMA, Nylon 6, and a UV-based fuel) with common oxidizers (LOX and N2O). AM fuels exhibit a similar theoretical performance as the commonly used HTPB fuel, and proper selection of the fuel can yield improved performance and design metrics. Development of AM approaches for HRE fuel design have significantly expanded their design trade space and should enable the competitive application of HREs for future propulsion missions.

Numerical Simulations of Fuel Shape Change in Paraffin–Oxygen Hybrid Rocket Engines

Reynolds-averaged Navier–Stokes simulations with submodels of turbulence, chemistry, fluid–surface interaction, and radiation are performed in this work to rebuild the internal ballistics of an experimentally tested hybrid rocket engine with paraffin and gaseous oxygen as propellants. Firstly, the effects of the prechamber and postchamber cavities at the initial, average, and final diameter of a reference burn are assessed to be negligible. Then, numerical simulations modeling the fuel shape change in space and time are compared to simulations performed at uniform port radius. The latter provide reasonable regression rate, pressure, and final grain profile predictions with reduced computational cost. On the other hand, the more computationally expensive fuel shape change simulations improve the model prediction capabilities providing a more accurate comparison with experimental data. The fuel shape change approach is finally applied with success to simulations of a throttled burn.

Model for Combustion Instability in Hybrid Rocket Engines Burning Liquefying Fuels

Starting from the data measured in several static firings of two laboratory-scale hybrid rocket engines of different sizes burning gaseous oxygen with either a liquefying fuel or a pyrolyzing polymer, a mechanism for the combustion instability observed only with liquefying fuel is suggested. With the latter type of fuels, regression rate is increased for the formation of an unstable liquid layer on the grain surface, for which liquid fuel structures are entrained into the gas stream and burn, upon vaporization and mixing, far from the surface. Based on this circumstance, the hypothesis of a time delay to burn, with the resulting effect of triggering longitudinal acoustic modes of the system, is formulated. A simplified analytical acoustic model of a series of interconnected ducts with different cross sections and gas properties is preliminarily addressed, which allowed demonstrating the acoustic nature of the main pressure oscillation peaks measured in three firing tests along with the influence of the oxygen-feeding line on the system acoustics. Finally, by leveraging an existing acoustic model of the hybrid rocket combustion chamber including the effects of cross-sectional discontinuities and distributed fuel mass addition and combustion, the linear stability of the system is analyzed, whose results are apt to reproduce the experimental findings.

Influences of steps in hybrid rocket engines:Simulation and validation on simplified geometries

In this study, we investigate the effect of single steps inside the fuel grains of typical hybrid rocket motors numerically. Steps can be a promising candidate to increase the regression rate of hybrid engines. We show that simulations on simplified geometries yield surprisingly close numerical regression rate profiles when compared to experimental data. Moreover, there is no necessity to calculate the progression of the fuel port at different time steps, as a single simulation on the average diameter can match the experiments sufficiently well. This observation allows for larger parametric studies on optimal positioning of the steps while limiting computational resources. The potential to increase the average regression rate through steps can be predicted numerically, and the simulations are following the same trend as the experiments. The augmented mixing and heat transfer inside the turbulent boundary layer can be predicted. It is concluded that backward facing steps increase the regression rate proportional to the step height up to 10 mm, and are preferably utilized upstream. For forward facing steps, the regression rate augmentation is saturated starting with 5 mm step height, and the position should be further downstream.

Charactering Catalytic Ignition and Combustion for the Ninety-Percent Hydrogen-Peroxide/Paraffin-Based Hybrid Rocket Engine

The present work performed a combined experimental and numerical investigation of the catalytic ignition and combustion characteristics for a hybrid rocket engine (HRE) using 90% hydrogen peroxide () with a paraffin-based fuel as the propellant system. The variation of the decomposition temperature used as the analytical premise was monitored by a thermocouple and laser absorption spectroscopy based on water absorption. The effects of preheating, the number of silver mesh layers, and the usage count for the catalytic bed on the decomposition characteristics were analyzed. A preheated 60-layer silver catalyst bed provided the best decomposition performance, including the shortest catalytic decomposition delay time and the highest decomposition efficiency. A two-dimensional numerical model was built to determine the effect of decomposition efficiency decay on the HRE combustion characteristics. Variations in the combustion temperature, combustion chamber pressure, and product distribution were compared for decomposition efficiencies ranging from 70 to 100%. As the decomposition efficiency of decreased to 80%, part of the oxidizer was blown out the chamber before participating in the combustion, yielding a significant decline in the combustion performance of the hybrid rocket engine. Static firing tests were also used to determine the catalytic ignition and reignition performance of the HRE. The results showed that the performance of the catalytic bed was stable while the reignition process was rapid without a pressure peak.

A new strategy for the reinforcement of paraffin-based fuels based on cellular structures: The armored grain — Ballistic characterization

Slow regression rate of the solid fuel is the main limitation for the use of hybrid rocket engines in high thrust applications. Paraffin-based fuels tackle this limitation thanks to the entrainment mass transfer. In this study, ballistic behaviors of conventional polymeric fuel (ABS) and paraffin-based blends are studied and compared with those of the armored grains. These latter are a new generation of fuels featuring 3D printed cellular structures embedded in the wax-based grain. The ballistic characterization focuses on the evaluation of the regression rate (rf) and its dependence on the oxidizer mass flux. Relative ballistic grading of the formulations is pursued via thickness over time methods and an optical technique for rf determination. The armored grains are reinforced by gyroid structures that are 3D printed using three different polymers (ABS, PLA, and Nylon 6) and two relative densities (10% and 15%). Despite the slow burning behavior of the printing polymers, the embedded reinforcement enhances the rf of the paraffin-based formulations, with percent increases ranging from +48% to +91%. This result could be explained by the uneven and irregular texture of the burning surface promoting turbulence (and therefore, propellant mixing) and convective heat transfer. For both the armored grains and the paraffin-based formulations, blending the pristine paraffin wax with polymeric additives results in more viscous formulations and in a rf reduction. Armored grain combustion performance makes this novel fuel an interesting candidate for high-thrust hybrid rockets.