Hot-fire Tests Research Articles

Experimental Study on the Mass Flow Rate Characteristics of a Fluidized Bed Powder Fuel Feeding System

In this study, a fluidized bed powder fuel feeding system that controls flow using the principle of two-phase gas-solid choking is proposed. Experiments based on real-time weighing methods were conducted to explore the flow rate characteristics under various total pressure and throttle-orifice area conditions. The experimental results showed that as the total pressure increased, the powder flow rate first decreased and then increased. To achieve the same increase in powder mass flow rate by altering the pressure, fine powder requires a greater pressure increment than coarse powder. A choked flow formula based on the gas-solid two-phase nonequilibrium model was derived, for investigating the influence of these two factors on the powder flow rate. Compared with the experimental results, this formula predicts the powder flow rates within a 20 % error margin. This model provides more precise guidance for engineering designs. The studied powder fuel-feeding system was connected to a powder engine, and hot-fire tests were conducted. The results showed that the pressure fluctuation in the powder engine combustion chamber did not exceed 2.5 %, confirming the feasibility and stability of the powder supply in the fluidized bed powder fuel feeding system.

Longitudinal combustion instability in a hypergolic liquid bipropellant combustor with single dual-swirl coaxial injector

Self-excited longitudinal combustion instabilities were investigated in a hypergolic liquid bipropellant combustor, which applied single dual-swirl coaxial injector. Hot-fire tests were conducted for four different injector geometries, while extensive tests on injection conditions were carried out for each injector geometry. The synchronous measurement of the pressure and heat release rate was applied, successfully capturing the process of the pressure and heat release rate enhanced coupling and developing into in-phase oscillation. By calculating Rayleigh index at the head and middle section of the chamber, it is shown that Rayleigh index of the middle section is even higher than that of the head, indicating a long heat release zone. When the combustion instability occurs, the pressure in propellant manifolds also oscillates with the same frequency and lags behind the oscillation in the combustor. Compared to the oscillation in the outer injector manifold, the oscillation in the inner injector manifold shows a higher correlation with that in the chamber in amplitude and phase. Based on numerical simulations of the multiphase cold flow inside the injector and combustion process in the chamber, it is found that injector geometries affect longitudinal combustion instability by changing spray cone angle. The spray with small cone angle is more sensitive to the modulation of longitudinal pressure wave in combustion simulations, which is more likely to excite the longitudinal combustion instability. Meanwhile, the combustion instability may be related to the pulsating coherent structure generated by the spray fluctuation, which is determined by injection conditions. Besides, a positive feedback closed-loop system associated with the active fluctuation and passive oscillation of the spray is believed to excite and sustain the longitudinal combustion instability.

Effect of the combustion chamber length in combustion instability of low-toxic hypergolic thruster

In this study, the effect of combustion chamber length on the hypergolic combustion instability of hydrogen peroxide-based propellants was investigated. Twenty-four hot-firing tests were conducted using a combination of 95 wt.% hydrogen peroxide and amine-based fuel with a drop test ignition delay of 5.65 ms and an adjustable length hypergolic thruster. When the chamber length was changed from 80 mm to 120 mm, the root mean square (RMS) combustion instability decreased drastically from 24 % to 9 %. The measured high-frequency instability was considerably consistent with the longitudinal resonance mode of each combustion chamber geometry. Low-frequency instability, that is, the rate of popping, occurred predominantly in all hot-firing tests. Within the 245–418 Hz range, its frequency increased as the chamber length decreased or the chamber pressure increased. The high-speed camera image of the exhaust plume coincided with the period of low-frequency instability, which was confirmed by the periodic popping of the propellant. Combustion instability was analyzed in depth by performing power spectral density (PSD), wavelet synchro squeezed transform (WSST), dynamic mode decomposition (DMD), and image intensity analyses using the chamber pressure and exhaust plume images. DMD decomposed the plume behavior into one expansion mode and three plume decay modes, and it also matched the low-frequency instability of the chamber pressure with an error of less than 5 %.

Early detection of thermoacoustic instability in an O2/CH4 single-injector rocket combustor using analysis of chaos and deep learning models

Thermoacoustic instability (TAI) presents a critical challenge for lean-burning combustors and rocket engines. The early detection of instability is crucial, and to address this, a data-driven prediction framework has been established for TAI in a sub-scale rocket combustor with variable chamber length. Nonlinear combustion features are generated from time series of dynamic pressure using recurrence matrices. Deep learning models are then utilized to train these features and predict the proximity of impending TAI. The performance of the proposed method is investigated through cross-validations of 12 groups of hot-fire test datasets. Remarkably, the prediction performances are in good agreement with measured experimental data, with most instabilities being predicted dozens of milliseconds in advance. This capability paves the way for the early implementation of active control systems in full-scale combustors in the future. The prediction performances are also compared with state-of-the-art TAI prediction methods.

Advancement of extreme environment additively manufactured alloys for next generation space propulsion applications

The National Aeronautics and Space Administration (NASA) has been involved in the development and maturation of metal additive manufacturing (AM) for space applications since the late 2000's. Several efforts have focused on the understanding of AM processes through material characterization and testing, standards development, component fabrication, and infusion into propulsion development and flight applications. NASA matured commonly used aerospace alloys from various alloy families (Nickel, Copper, Stainless and Steel, Aluminum, and Titanium-based) through detailed AM process and heat treatment characterization, in addition to mechanical and thermophysical testing. While these alloys are actively used in many propulsion applications, there is a need for ongoing AM optimized alloys using integrated computational materials engineering (ICME) and process development for high performance applications. The applications targeted are liquid rocket engines; advanced propulsion systems; and in-space propulsion with high heat fluxes, high pressure, and/or that use propellants that can degrade alloys (e.g., hydrogen). This paper highlights the characterization and physical properties of the more common AM alloys using laser powder bed fusion (L-PBF) and laser powder directed energy deposition (LP-DED) processes. Additionally, this paper discusses some of the ongoing novel alloy development and maturation using AM for use in these harsh environments, such as GRCop-42, GRCop-84, NASA HR-1, GRX-810, and C-103. The results from these processes demonstrated that AM could enable rapid development, and that optimized alloys could be developed using ICME, yielding higher performances. These alloys have undergone modeling, fundamental metallurgical evaluations, heat treatment studies, detailed microstructure characterization, and mechanical testing campaigns. This, combined with direct application-specific component fabrication and hot-fire testing, enabled the increase of the Technology Readiness Level (TRL) through high duty-cycle testing. A background and overview of these novel AM-enabled alloys and AM processing developments including metallurgical and mechanical property studies is presented here. The latest advancement in the parallel component development and hot-fire testing and future developments for these alloys is also discussed.

Two-Stage Impinging-Jet Injector Flow Dynamics and Mixing: Kerosene and Hydrogen Peroxide Propellants

ABSTRACT Kerosene/hydrogen-peroxide (H2O2) green propellant system has relatively high operational ratio of oxidizer (O) to fuel (F) that complicates the injector design and its layout. In this research, a procedure of two-stage, like-doublet (O-F-F-O) impinging-jet injector design is presented, where mixing of H2O2 and kerosene occurs between O-O and F-F sprays. Investigations at O/F ratios between 3.75 and 6.25 are performed. The impinging distance, angle, and jet velocity of the two-stage impinging-jet module are the design parameters examined. The PLIF technique is used to observe the spatial droplet distribution of the individual spray. The predicted mixing behavior by overlapping the individual fuel and oxidizer sprays is compared to the actual mixing of the O-F-F-O spray to justify the spray/spray interactions. Detailed behaviors of the fuel and oxidizer sprays are revealed. The effects of O/F ratio and other design parameters on the injector flow dynamics and propellant mixing are examined systematically. With experimentally defined shape factors, an optimum design of the injector element can be determined for a given injector plate. The design of the injector plate for a 450 N thrust engine is demonstrated, and the design is justified by the hot-fire test of the engine.

Simulation and analysis of low-frequency instability in hot-fire test of a cryogenic hydrogen-oxygen preburner

Low-frequency chamber pressure oscillation is observed in the hot-fire test of a subscale cryogenic hydrogen-oxygen preburner. A lumped parameter simulation model with bipropellant injector is established by solving differential equations directly. Under the condition of constant inlet pressure, the frequency of oscillation of combustion chamber pressure increases with the decrease of combustion delay, while the variation trend of amplitude of oscillation of chamber pressure is opposite. As the combustion delay decreases to 1.76 ms, the pressure oscillation gradually converges and stabilizes, but its initial oscillation frequency is still about 2.7 Hz lower than that of the test. In addition, the influence of the volume ratio of liquid propellant in the combustion chamber is analyzed. The combustion time delay is set to 1.76 ms, and the frequency and amplitude of the pressure oscillation increase with the increase of the liquid volume ratio. With the liquid volume ratio set at 2.8%, the frequency and amplitude obtained by simulation are in good agreement with the test. In order to reproduce the continuous oscillation of the combustion chamber pressure during the 30 s process in the test, the pre-injection pressure test curve varying with time is used as the inlet condition. The liquid volume ratio and combustion delay are empirically corrected in real time using propellant flow rate and mixture ratio, respectively. The continuous oscillation of the simulated pressure during the 30 s test was obtained, and its frequency value and trend are completely consistent with the test data. It can be inferred that low-frequency unstable combustion related to combustion delay may have occurred in the hot-fire test of the preburner.

Experimental Investigation of Flame Anchoring Behavior in a LOX/LNG Rocket Combustor

Hot fire tests of a multi-injector research combustor were performed with liquid-oxygen and liquefied-natural-gas (LOX/LNG) propellants at chamber pressures from 30 up to 67 bar, hence at conditions similar to an upper stage rocket engine. Within these tests shear coaxial injectors were tested with and without a recessed LOX post. In both configurations, operating conditions with flames anchored at the LOX post tip and thus, if available, pre-combustion in the recess volume as well as lifted flames were observed. Flame anchoring was indirectly detected via acoustic measurements, using mean speed of sound to indicate the presence of flame in the head end of the combustion chamber. While the injector without recess showed only stable combustion irrespective of the flame anchoring behavior, the recessed injector featured short-lived bursts of oscillatory combustion and sustained combustion instabilities. Analysis of the test data showed that stable flame anchoring could not be ensured at momentum flux ratios below 20 for a non-recessed and below 45 for a recessed injector.

Numerical and experimental investigation of a Mg/N2O powdered fuel rocket engine

Powdered metal fuel rocket engines, utilizing high energy density metal powder as fuel, have great potential in the development of propulsion systems for upper-stage rockets, and thrusters for attitude or trajectory control. The main challenges of this rocket engine technology currently lie in the realization of stable powder fuel supply, quick engine start-up, and high combustion efficiency. In this paper, we present a computational analysis of the reactive flow in a Mg/N2O powdered metal fuel rocket engine. Consideration in this work includes effects of a flame holder, mass flow rate ratio and entry angle of N2O on combustion efficiency. Based on the numerical simulation results, an experimental Mg/N2O powdered metal fuel rocket engine was designed, and working processes of the engine were investigated through hot fire tests with Mg powder as fuel and N2O as oxidant. A ground test system, consisting of a high-precision test bench, a fuel feed unit, an oxidant feed unit, and a measurement and control unit, was established. The test results show that the combustion of Mg powder and N2O can be sustained.

Flame dynamics of an injection element operated with LOX/H2, LOX/CNG and LOX/LNG in a sub- and supercritical rocket combustor with large optical access

Hot fire tests were performed using a single-injector research combustor featuring a large optical access window ([Formula: see text] mm) for flame visualisation. Three test campaigns were conducted with the propellant combination of liquid oxygen and hydrogen, liquid oxygen and compressed-natural-gas, as well as liquid oxygen and liquefied-natural-gas at conditions relevant for main- and upper-stage rocket engines. The large optical access enabled synchronised flame imaging using ultraviolet and blue radiation wavelengths covering a large portion of the combustion chamber for various sets of sub- and supercritical operating conditions. Combined with temperature, pressure and unsteady pressure measurements, this data provides a high-quality basis for the validation of numerical modelling. Flame width, length and opening angle as features describing the flame topology were extracted from the imaging. The suitability of flame imaging using ultraviolet and blue radiation wavelengths as qualitative markers of heat release was evaluated. Two-dimensional distributions of the Rayleigh Index were calculated for intervals with and without high-amplitude, self-excited oscillations of the longitudinal acoustic resonance modes. The calculated Rayleigh Index values properly reflect the thermoacoustic state of the chamber, indicating that both types of imaging may be used for qualitative study of thermoacoustic coupling of high-pressure cryogenic flames.

Rocket noise predictions and scaling methods based on outdoor acoustic measurements

This paper will present follow-on work to the outdoor acoustic measurements that were presented at the May 2022 ASA conference for a static hot fire test of a rocket engine. The focus is on scaling the measured sound pressure levels to predict the acoustic performance of a different sized engine. Source models were developed based on the near field acoustic measurements and compared to far field measurements.

Conceptual Design of Small-Sized Thruster Using Laser Ignition of High-Energy Monopropellant

Ammonium dinitramide (ADN)-based energetic ionic liquid propellants (ADN-EILPs) present considerable potential as monopropellants due to high energy density, high thermal/chemical stability, and low toxicity. Consequently, a chemical thruster system based on ADN-EILPs can be employed in the development of high-performance ultrasmall/small satellites. However, the characteristics of the ADN-EILPs present various problems in their ignition during the thruster development. To solve this problem, the authors have previously achieved the continuous-wave (CW) laser ignition of ADN-EILPs using laser absorbers composed of carbon wools. However, the conventional propellant injection of ADN-EILPs with carbon wool using high-pressure gas faces several limitations. Therefore, a propellant feed system suited for the ignition method is proposed here, as well as a conceptual model of a 0.5 U/0.5 N-class thruster operated by the CW laser ignition of ADN-EILPs (). Additionally, an attempt is made to manufacture a laboratory model (LM) thruster, and its fundamental operation properties are determined. The future research implications of this study include the further observation of the combustion of the decomposed gas before its emission from the throat of the LM thruster, along with the further development of the propellant feeding system proposed in this study and the hot-fire tests performed for the feeding systems.

Design of pintle injector using Kerosene-LOx as propellant and solving the problem of pintle tip thermal damage in hot firing test

The pintle injector is considered to have high reliability because it can control engine thrust by adjusting propellant flow rate via adjusting the orifice area and is strong against combustion instability. Although many studies have been conducted on pintle injectors, their accessibility is limited and studies on performance optimizations for propellant type and propellant mixing efficiency according to pintle shape and atomization are incomplete. In this study, a 1.5-tonf class liquid–liquid pintle injector with rectangular two-row orifices that uses kerosene/liquid oxygen as the propellant was designed and manufactured. Combustion tests were performed on the pintle injector to verify performance and stability under supercritical conditions, which are the actual operational conditions of liquid rocket engines. From the combustion tests on the initial prototype, the pintle tip was observed to be damaged by heat, so the pintle injector design was changed, and thermal fluid analysis was performed to analyze the pintle tip cooling and combustion performances. To increase the cooling performance of the pintle tip, we devised a method of changing the shape of the pintle orifice and inserted a cooling device called an insert nozzle into the pintle without material changes or applied coatings, as in previous studies. The thermal fluid analysis results showed that there was a difference of up to 147–400 K in the pintle tip cooling performance depending on the insert nozzle and blockage factor, which was verified through combustion tests. The characteristic velocity efficiency and heat flux showed increasing tendencies with increase in total momentum ratio, and a difference of up to 2.0% was confirmed for the characteristic velocity efficiency.

Dynamic characterization of wall temperature in LOX/CH[formula omitted] rocket engine operating conditions using phosphor thermometry

Accurate surface temperature prediction is essential for the lifetime and efficiency of a rocket engine. The new ambition of future reusable engines requires perfect knowledge of the thermal environment encountered. In the design phase of such systems and to improve numerical simulations, more resolved experimental data are needed. Experiments are thus conducted by Lab. EM2C, ONERA, CNES, and ArianeGroup on the cryotechnic test bench MASCOTTE operated by ONERA. Conditions representative of those encountered in the gas generator of a rocket engine are investigated. Among them is the cryogenic injection of transcritical methane and oxygen at a very low mixture ratio (0.3) and high pressure (up to 60 bar) through a single coaxial injection element. The present work aims at characterizing the temperature of several surfaces of the combustion chamber in a series of hot fire tests performed in the transient thermal regime and in a short period of time (∼30 s). For this purpose, dynamic Laser Induced Phosphorescence (LIP) is implemented with the Full Spectrum Fitting method (FSF method), a dedicated approach developed by Lab. EM2C. Single-shot measurements are performed to simultaneously track the temperature of five surfaces at 10 Hz located near and downstream of the flame with a 30 K uncertainty. Data indicate that the temperature peak at ignition can be recorded with LIP measurements in contrast to conventional sensors. The analysis of gas and surface temperatures shows that convection is the predominant heat transfer mode and drives the surface temperature. Thanks to LIP thermometry on the internal and external faces of one of the silica windows, a 1D semi-infinite medium model is applied, showing excellent agreement with the measurements. Fitting this model on the measurements allows the estimation of the local convection coefficient and the prediction of the temperature at any depth inside the walls.

Experimental Investigation of Throttleable and Polypropylene Hybrid Rocket Motor

This study experimentally investigated the propulsion performance of a swirling-type throttleable hybrid rocket motor using hydrogen peroxide and polypropylene. An in-house throttle valve consisted of a servomotor, and an industrial ball valve was integrated to instantaneously control the oxidizer mass flow rate. Several ground static hot-fire tests with fixed oxidizer mass flow rates of (100%) and (82%) and many step-varying oxidizer mass flow rates were performed to examine its propulsion performance. The hybrid rocket motor achieved an impressive sea-level of 240 s, an engine efficiency of more than 97%, and a thrust uncertainty of less than 5% when the oxidizer flow rate was fixed as . For the multiple-step hot-fire test results, the thrusts and chamber pressures were both found nearly proportional to the oxidizer mass flow rates with different proportionalities. The overall ratio of the tests is 5.61, which is hardly shifted, unlike most hybrid rocket data reported previously. In summary, the total impulse and thrust are both linear to total oxidizer mass consumption in the range of 50–100% of . This major finding would indeed broaden the application of hybrid rocket motors to future space technology.

Flame Characteristics and Response of a High-Pressure LOX/CNG Rocket Combustor with Large Optical Access

Hot-fire tests were performed with a single-injector research combustor featuring a large optical access (255 × 38 mm) for flame imaging. These tests were conducted with the propellant combination of liquid oxygen and compressed natural gas (LOX/CNG) at conditions relevant for main- and upper-stage engines. The large optical access enabled synchronized flame imaging using OH* and CH* radiation wavelengths covering an area of the combustion chamber from the injection plane to shortly before the contraction section of the nozzle for two sets of operating conditions. Combined with temperature, pressure and unsteady pressure measurements, these data provide a high-quality basis for validation of numerical modeling. Flame width and opening angle were extracted from the imaging in order to determine the flame topology. A two dimensional Rayleigh Index was calculated for an acoustically unexcited and excited interval. These Rayleigh Indices are in good agreement with the thermoacoustic state of the chamber.

Energy Estimation and Testing Verification on Ignition of a Torch Ignition System

The torch type electric ignition system is a desirable choice for reusable liquid vehicle engine. It is generally believed that the ignition and start-up process of engine are crucial and high-risk. However, due to the complexity of combustion experiments, the ignition and start-up process can be efficiently analysed from the perspective of energy exchange. By building the energy analysis equations of electric spark plug, ignition chamber and thrust chamber, the energy estimation method of torch type electric ignition system is established. Furthermore, the time scheduling and examining method for multi-injector ignition and start-up process of large flow liquid rocket engine with multi-injector are established. The hot-firing test of LOX/CH4 rocket engine is carried out according to the design parameters to validate the effectiveness of the energy estimation method.

Assessment of thrust chamber stability margins to high-frequency oscillations based on mathematical modeling of coupled ‘injector – rocket combustion chamber’ dynamic system

High-frequency instability of a liquid-propellant rocket engine (LRE) during static firing tests is often accompanied by a significant increase in dynamic loads on the combustion chamber structure, often leading to a chamber destruction. This dynamic phenomenon can also be extremely dangerous for the dynamic strength of a liquid-propellant rocket engine. The calculation of acoustic combustion product oscillation parameters is important in the design and static firing tests of such rocket engines. The determination of the oscillation parameters (natural frequencies and stability margins such as oscillation decrement) is one of the problems solved in the LRE design period as part of the development of measures to ensure the engine stability. The main aim of the paper is to develop a numerical approach to determining the parameters of acoustic oscillations of combustion products in liquid-propellant rocket engines combustion chambers taking into account the features of combustion space configuration and the variability of gaseous medium physical properties depending on the axial length of the chamber, acoustic impedance in critical throat and dissipation effects (damping experimental values) in the shell structure and the gas media in the chamber. The approach is based on mathematical modeling of the coupled ‘chamber shell structure – gas’ dynamic system by using the finite element method and the CAE (Computer Aided Engineering) system. The developed approach testing and further analysis of the results for the RD 253 engine using nitrogen tetroxide and unsymmetrical dimethylhydrazine as a propellant pair were carried out. The dynamic system shapes and frequencies of longitudinal, tangential and radial modes are determined. The results of mathematical modeling of the dynamic system indicate a satisfactory agreement of the calculated decrements of the first longitudinal oscillation mode and third tangential oscillation mode with the experimental decrements obtained by hot-fire tests data. From system harmonic analysis of the thrust chamber, it follows that the dynamic pressure gain factor of the gas media in the chamber at the first longitudinal mode frequency is 1.6 times greater than the system dynamic gain in the tangential mode. At the same time, the oscillation decrement for the system tangential mode is 2 times smaller than that of the first longitudinal mode. This means that the thrust chamber tangential mode is more dangerous and can lead to rocket engine combustion instability. The effect of the injector on the high-frequency stability of the combustion chamber and the possibility of partial suppression of combustion chamber thermoacoustic oscillations by adjusting the high-frequency dynamics of the injector are shown theoretically.

Demonstration of ammonia borane-based hypergolic ignitor for hybrid rocket

The low toxicity hypergolic solid fuels (HSFs) and hydrogen peroxide (H2O2) oxidizer have been demonstrated as environment-friendly hybrid propellants. In the present study, a new HSF containing ammonia borane (AB), Pd–C, and paraffin wax, was fabricated in the form of the hypergolic ignitor for the hybrid rocket. Furthermore, the atmospheric combustion test (ACT) and hot-fire test (HFT) with HSF ignitor, PMMA or PE fuel, and 95% H2O2 oxidizer were performed using fuel-exposed-hybrid-combustor; studied the ignition reliability and combustion characteristics of HSF. The hypergolic ignition of HSF initiated through Pd–C followed by AB combustion with H2O2 was observed by using the high-speed camera images of ACT. A HFT with a ground thrust of 120 N exhibited the rising time and pressure instability of 137 ms and ±4.00%, respectively. Moreover, smooth combustion transition from HSF to PE was achieved with stable chamber pressure and no hard-start. The results suggest that the HSF ignitor containing AB may be the suitable preference for the hypergolic hybrid rocket.

Experimental study of the effects of Weber number and recess length on the combustion characteristics for liquid-liquid swirl coaxial injector

The injector is a vital component of a liquid rocket engine that controls the atomization and mixing of propellant, which significantly on combustion performance. In this study, two liquid-liquid swirl coaxial injectors (LLSCI) with different recess lengths were developed, and the effects exerted by fuel (liquid methane LCH4) and oxidizer (liquid oxygen LOx) Weber numbers on combustion efficiency were examined by performing hot-firing test experiments. In addition, a heat transfer model of the combustion chamber was developed. By measuring the temperature at a spot on the combustion chamber wall, the temperature of any depth at that point and the temperature near the chamber wall could be deduced. The calculated results well complied with the experimental measurements and the maximum error was 5.8%. Furthermore, an investigation was conducted on the effects of Weber number and recess length on the combustion flow field of the internal mixing swirl coaxial injector. As indicated from the results, the combustion efficiency generally increased with the increase in LOx Weber number and the decrease in LCH4 Weber number. However, under the higher Weber numbers of the two propellants, the combustion efficiency and the gas temperature distribution were affected by the location of the combustion area and the degree of mixing. The combustion efficiency was elevated by 3.65% when the recess length increased by 2 mm. However, the high-temperature area of combustion close to the head would be caused, as well as ablation damage.