Hot-fire Tests Research Articles

Outdoor acoustic measurements of a rocket engine hot fire test

The primary focus of this paper is to present the results from outdoor acoustic measurements from a rocket engine hot fire test. The two objectives of the test were to quantify the overall sound pressure level at large distances from engine to quantify the impact of the events on the surrounding community, and to better understand the source distribution in the engine plume. Microphone arrays were positioned around the property to be able to capture the sound propagation, estimate source levels, and conduct phased array-based plume source localization.

Analysis of self-excited transverse combustion instability in a rectangular model rocket combustor

A methane/oxygen mixture is considered to be an appropriate propellant for many future rocket engines due to its practicality and low cost. To better understand the combustion instability in methane/oxygen-fed rocket engines, the spontaneous transverse combustion instability in a rectangular multi-element combustor (RMC) was analyzed both experimentally and numerically. Severe combustion instabilities occurred in the RMC during repeatable hot-fire tests. The physical mechanisms were systematically investigated through numerical simulations based on the stress-blended eddy simulation and flamelet-generated manifolds method with detailed chemical mechanisms (GRI Mech 3.0). The numerical results for the frequency spectrum and spatial modes agree well with the theoretical analysis and experimental data. The driven regions of the combustion instability were identified on both sides of the combustion chamber through a Rayleigh index analysis. The longitudinal pressure oscillations in the oxidizer post were found to be coupled with the transverse pressure waves in the combustion chamber and led to periodic oscillations of the mass flow rate of propellant. Moreover, the mixing was highly enhanced when the pressure wave interacted with walls of the combustion chamber. Therefore, a sudden release of heat occurred. The pressure oscillations were enhanced by pulsated heat release. A closed-loop system with positive feedback associated with periodic oscillations mass flow rate of the propellant, and sudden heat release, was believed to account for the present combustion instability.

In-Situ Optical Measurements of Solid and Hybrid-Propellant Combustion Plumes

A method for in-situ optical measurements of solid and hybrid propellant rocket plumes is developed, and results from proof of concept tests are presented. The developed method inserts fiber-optic cables acting as radiation conduits into the solid-fuel combustion port, allowing optical signals to be transmitted from the flame zone to externally-mounted spectrometers. Multiple hot-firings using a using a lab-scale gaseous-oxygen, thermo-plastic fueled hybrid rocket system were performed to validate the sensing method. Burn durations varied from 5 to 25 s, and the inserted fiber optic sensors survived for all of the hot fire tests. The obtained optical spectra were curve-fit to Planck’s black-body radiation law, and Wien’s displacement law was used to estimate the internal flame-temperature. Optically-sensed flame-temperatures are correlated to analytical predictions, and shown to generally agree within a few degrees. Additionally, local maxima in the optical spectra are shown to correspond to emission frequencies of atomic and molecular hydrogen, water vapor, and molecular nitrogen; all species known to exist in the hybrid combustion plume. Based on these preliminary test results, it is concluded that this simple in-situ measurement system operates as designed, and it shows considerable promise for future applications to a wide swath of gas-generator systems.

Hot Firing Tests of a Novel Green Hypergolic Propellant in a Thruster

This study investigates the ignition and combustion behavior of the recently developed green hypergolic combination named HIP_11 in a small thruster. The combination HIP_11 consists of highly concentrated hydrogen peroxide as oxidizer and a fuel based on the ionic liquid 1-ethyl-3-methylimidazolium thiocyanate (EMIM SCN) with addition of copper thiocyanate (Cu SCN) to facilitate the hypergolic ignition. A dedicated thruster, featuring a 2-on-1 unlike impinging jet injector, was designed and manufactured. The initial tests were conducted with the injector only, without the combustion chamber mounted. These tests were conducted with two different amounts of catalyst: 1% and 5% Cu SCN. Subsequently, three hot firing series (around 35 tests) were conducted with the complete thruster, using the propellant combination 97% high-test peroxide and EMIM Cu SCN. The propellant mass flow rates were varied between 7 and and the oxidizer-to-fuel ratio from 3 to 7.

Experimental investigation on effects of injection distribution on combustion instability in a model rocket combustor

Self-excited combustion instabilities of transverse modes were experimentally investigated in a rectangular multi-injector model combustor, operating with the bipropellants O2/CH4. The propellants were injected through a linear array of five oxidizer-centered shear coaxial injectors into the combustor. High-amplitude limit cycles obtained in hot-fire tests were analyzed in detail. Different combustion instability modes, including first and second width modes, were observed in cases with three different injection distribution schemes. Hence, the injection distribution strongly determined the combustion dynamics. One insight can be gained that the stable combustion could be achieved by properly designing the propellants' injection distributions.

Design and experimental research of a sub-Newton N2O monopropellant thruster with inner-heater

Nitrous oxide (N2O) is a green propellant with excellent application prospects. A sub-Newton N2O monopropellant thruster with inner-heater and a N2O self-pressurization stable supply system with regenerative heat compensation are designed in this paper. The experimental research of the thruster is described, including measurements of preheating power, activation temperature, vacuum thrust, specific impulse, life-span and pulsed operation performance. By inserting the heater into the catalyst-bed, preheating efficiency of the heater is significantly improved compared to the thruster with outer-heater. Thus, the preheating power demand of the thruster is successfully reduced to around 10 W. The mean vacuum thrust of 322 mN is attained and the corresponding specific impulse reaches 162 s at the mass flow rate of 0.2 g/s. Successful activation temperature of 523 K is achieved, and the activation performance of the thruster is affected by the loading factor. A long term hot-firing test longer than 12000 s is attained. The pulsed operation performance of the inner-preheating thruster is also studied by measuring chamber pressure. Impulses with different magnitudes are produced by adjusting the opening duration of the control valve. A minimum impulse of 81 mN∙s is attained. Finally, the performance of the thruster is evaluated by comparison with other thrusters of the same type. The results indicate that the proposed thruster with inner-heater is superior in terms of preheating power, activation temperature and specific impulse performance.

Flame characteristics of a high-pressure LOX/H2 rocket combustor with large optical access

Hot fire tests were performed with a newly developed single-injector research combustor featuring a large optical access (255 × 38 mm) for flame imaging. Within these tests the propellant combination of liquid oxygen and cold gaseous hydrogen (LOX/H2) has been used at chamber pressures up to 70 bar and ratios of oxidizer to fuel (ROF) at the main injector of 3.4–6. The large optical access enabled synchronized flame imaging using OH* and blue radiation wavelengths covering an area of the combustion chamber from the injection plane to shortly before the contraction section of the nozzle. High-speed flame imaging was realized for three different operating conditions. Combined with data from temperature, pressure and unsteady pressure sensors high-quality validation data for numerical modeling has been gained. Several key dimensions of the flame topology were extracted from the imaging. The change in flame topology in response to different operating conditions is consistent with expectations from the literature, in particular the dependence of flame opening angle on injection momentum flux ratio. Using OH* imaging for such topology measurements was found to have less uncertainty than with blue radiation imaging.

Experimental and numerical investigation of swirling H[formula omitted]O2 and polypropylene hybrid rocket motor with regenerative cooling

In this paper, the experimental and numerical investigations of both swirling hybrid rocket motors without and with the design of regenerative cooling (RC hereafter), using 90%-wt H2O2 and polypropylene (PP) as the liquid oxidizer and the solid fuel, respectively, were performed and compared. We implemented a new design of regenerative cooling using the oxidizer to transfer the heat generated from combustion. A series of ground horizontal hot-fire tests, which include the catalytic-bed only and the motor without/with RC were performed to measure the propulsion performance related parameters. The resulting C* efficiency of the RC design with catalytic-bed only remained in comparison with the one without RC. With a similar initial combustion chamber geometry, the RC implementation resulted in a better engine Isp efficiency from 96.3 to 98.6 with a lower O/F ratio from 6.49 to 5.94. In addition, the topologies of the burned fuel grain surface were optically measured by a 3D laser scanner. The resulting variation of regression rates between the upstream and downstream fuel for the chamber with the RC was smaller than case without the RC. Moreover, a simplified semi-empirical computational fluid dynamics (CFD) model based on the measured regression rates was constructed for investigating the physical phenomenon inside the hybrid rocket combustion chamber. The simulations show that a higher swirling number in the later portion of the port was found for the RC case. By concerning about the different O/F ratios and port size of the burned fuel, a series of simplified numerical CFD investigation with a fixed O/F and a fixed port size varying the regression rates from upstream to downstream fuel systematically with the same average region rate was performed to complement the semi-empirical CFD simulations. The results showed that there were two distinct regions including an upstream non-classical region dominated by the gas impingement of gas and a downstream classical region caused by the swirling effect of gases. In the non-classical region, the calculated distributions of the swirling numbers of the former are essentially the same for all the cases no matter what the magnitude of the regression rate is, while in the classical region higher swirling numbers are found for the case of more uniform distribution of regression rates. Heat generated near the upstream of the fuel grain was transferred more efficiently through the design of RC, which makes the resulting fuel regression more uniform than that without RC. The resulting swirling efficiency in the RC case was higher than that without RC. The HRM with RC hence had a better engine Isp efficiency with a lower O/F ratio.

Результаты испытаний модельной кислородно-метановой камеры сгорания жидкостного ракетного двигателя, созданной с использованием методов аддитивного производства

The paper focuses on an experimental unit developed for modeling combustion characteristics in a model oxygen-methane combustion chamber of a liquid rocket engine. The key components of the unit, i.e., the mixing head of the combustion chamber and the regeneratively cooled nozzle, were manufactured using advanced methods of additive manufacturing. The paper emphasizes the specific character of the combustion chamber components made with the use of additive technology and introduces hot-fire test results of the model combustion chamber as part of the experimental unit. The study shows the durability of the mixing head and combustion chamber nozzle under hot-fire test conditions, as well as the reliable operation of the experimental unit as a whole, which confirms the selected design and technological solutions. Within the study, we analyzed the cooling system of the experimental unit for the test conditions, estimated the thermal state of the nozzle, with account for the features of the additively manufactured cooling path. To increase the cooling system’s reliability and expand the combustion chamber pressure application, it is recommended to apply a heat-shielding coating on the firewall of the nozzle. Using new experimental data, we analyzed the parameters of improving the efficiency of the model combustion chamber with the additively manufactured components and corresponding in scale and consumption characteristics to the combustion chamber of the liquid rocket engine

Numerical investigation of pre-detonator in rotating detonation engine

Pre-detonators are commonly used in rotating detonation engine (RDE) experiments. Current experimental studies focus on the performance of pre-detonator while ignoring the influence of pre-detonator on the flow field. In numerical simulations one-dimensional detonation wave is usually used to ignite the fresh gas in RDE. This is a simplification of the pre-detonator used in practical hotfire tests. But the coupling between the pre-detonator and the combustor is ignored. The aim of the present study is to study the influence of pre-detonator on the flow field in the RDE. A model of RDE with a pre-detonator is built, in which three-dimensional numerical simulations fueled with hydrogen/air is performed. The influence of pre-detonator on the combustor in different stages is studied. After initiation, detonation wave from the pre-detonator forms two counter-rotating detonation waves. The tangential installation of pre-detonator fails in directional initiation of detonation wave. The coupling effect is shown as the reflection and expulsion of shock wave. Detonation wave or oblique shock wave in the combustion chamber enters the pre-detonator and turns into shock wave before colliding with the end and re-entering the combustion chamber. Under some circumstances, the reflected shock wave will initiate a detonation wave and affect the wave structure in the combustion chamber. In the stable stage, the reflected shock wave has no effect on the flow field. However, periodic collision of reflected shock wave with detonation wave at the junction causes ablation in long-time experiments. Increasing the axial distance between pre-detonator and injection wall is expected to be a solution for the ablation problem.

Computational and Experimental Study of Nozzle Performance for Rotating Detonation Rocket Engines

A computational and experimental study was conducted on nozzle geometries for rocket application rotating detonation engines (RDEs). Three geometries, including a nozzleless blunt body typically employed in RDE combustor hot-fire testing and two aerospike nozzles, were investigated. Simulations of the exhaust flow of a rotating high-frequency, high-pressure ratio wave based on rocket RDE test results were related to comparable constant-pressure conditions. Computational and experimental results showed the high momentum added by the highest-pressure detonation products influences the exhaust plume differently than a comparable steady flowfield fed by the same average product gas flow rate. In particular, the RDE exhaust flow tended to enhance entrainment on the nozzleless blunt body recirculation region and delay flow separation on nozzle expansion surfaces due to overexpansion compared to a constant-pressure engine. Results have important ramifications for isolating RDE combustor performance from nozzle effects and must be considered for future design of nozzle geometries to exploit the high-frequency, high-pressure ratio outflow of an RDE.

Speckle Pattern Inversion in High Temperature DIC Measurement

During hot fire rocket engine testing, non-contacting measurements are superior to bonded gauges because they are immune to burning, shaking loose, or damage due to the harsh testing conditions. Additionally, when compared to instruments which register at single points, Digital Image Correlation (DIC) has the added benefit in that it collects full-field displacement and strain maps over the duration of the test. However, for certain materials and paints under some circumstances of temperature and camera sensitivity, portions of the speckle pattern which were darker at room temperature may emit more light compared to the initially lighter portions of the pattern, resulting in a high temperature pattern which is inverted in comparison with that at room temperature. To address this inversion, a post-processing method is introduced wherein an inverted image containing only emitted light is subtracted from an image containing both emitted and reflected light, thereby generating an un-inverted image. The artificial high temperature image is subsequently correlated against the room temperature image to obtain full-field strains. The subtraction technique is then validated using optical bandpass filters to prevent significant amounts of emitted light from reaching the camera sensor. The two methods are mapped onto common coordinates and shown to produce comparable results. The subtraction method sufficiently mitigates speckle pattern inversion, but its key drawback is that it only works when there is negligible displacement between the subtracted images (i.e. quasi-static loading). It is therefore preferable to eliminate inversion from reaching the camera in the first place by using optical bandpass filters.

Performance of a Rotating Detonation Rocket Engine with Various Convergent Nozzles and Chamber Lengths

A rotating detonation rocket engine (RDRE) with various convergent nozzles and chamber lengths is investigated. Three hundred hot-fire tests are performed using methane and oxygen ranging from equivalence ratio equaling 0.5–2.5 and total propellant flow up to 0.680 kg/s. For the full-length (76.2 mm) chamber study, three nozzles at contraction ratios ϵc = 1.23, 1.62 and 2.40 are tested. Detonation is exhibited for each geometry at equivalent conditions, with only fuel-rich operability slightly increased for the ϵc = 1.62 and 2.40 nozzles. Despite this, counter-propagation, i.e., opposing wave sets, becomes prevalent with increasing constriction. This is accompanied by higher number of waves, lower wave speed Uwv and higher unsteadiness. Therefore, the most constricted nozzle always has the lowest Uwv. In contrast, engine performance increases with constriction, where thrust and specific impulse linearly increase with ϵc for equivalent conditions, with a 27% maximum increase. Additionally, two half-length (38.1 mm) chambers are studied including a straight chamber and ϵc = 2.40 nozzle; these shortened geometries show equal performance to their longer equivalent. Furthermore, the existence of counter-propagation is minimized. Accompanying high-fidelity simulations and injection recovery analyses describe underlying injection physics driving chamber wave dynamics, suggesting the physical throat/injector interaction influences counter-propagation.

Experimental study on the rocket-ejector system with a throat in the secondary stream

This study concentrated on hot-fire tests to demonstrate the flow characteristics that influenced by the throttle of the secondary stream in the rocket-ejector system. The experiment was conducted under the sea level static (SLS) condition. A small rocket used gaseous oxygen / liquid kerosene propellant as a primary stream supply system, and the high-pressure air bottles were employed as the secondary stream supply system. The equivalence ratio of the rocket, the air mass flow rate and the mixing length were varied to study the performance. The results show that there is a mass flow rate mismatch between the air flow entering through the inlet and the mass flow permitted due to aerodynamic choking. If the air mass flow rate through the throat is more, the inlet may unstart. If the air mass flow rate through the throat is less, the total pressure loss of the air due to the pseudo-shock would be caused under the effect of the suction performance. The unsuitable air mass flow rate is detrimental to RBCC engine operation in the ejector mode. With the increasing equivalence ratio of the rocket, the pressurization performance is improved accordingly, but the suction performance is obviously decreased. When the equivalence ratio is more than 1.1, the reacting shear layer would affect the suction performance.

Numerical modeling and experimental investigation on the rocket-ejector system with limited mixer length

Rocket-ejector system as a pressurization unit is of critical importance to the rocket-based combined-cycle engine under the low flight Mach number conditions. Due to the restriction of engine scale, an insufficient mixing process exists in the limited mixer length, leading to a non-uniform stream at the mixer exit. This study concentrated on hot-fire tests to demonstrate the flow characteristics of the rocket-ejector system with the limited mixer length. The results show that shortening the length reduces the critical backpressure of the rocket-ejector system. The rocket-ejector system with a mixer length of 6 length-to-diameter ratio to the straight section can resists the critical pressure ratio of 1.25 in which the entrainment coefficient is 3.432, the total pressure ratio is 36.4 and the occupied area ratio of the rocket exit to the straight section is about 32%. Although the non-uniformity of mixed stream at the mixer exit section would reduce the critical pressure to some extent, the limited mixer length also meet the operation requirements of the rocket-based combined-cycle engine with the diffusion and afterburning cycle.

An experimental study on the spray characteristics of splash platelet injector

An experimental investigation was performed to explore the characteristics of sprays produced by the splash platelet injector in a spray test facility. A high-speed camera was used to measure the flow and spray characteristics. During the experiments, four cases with different lengths of trapezoidal orifice and expanding angle were analyzed. The experimental results showed that the mass flow rate and flow coefficient decrease when the length of the trapezoidal orifice and expanding angle increase. The influence of the length of the trapezoidal orifice on the spray morphology is significant. The spray region increases when the length of the trapezoidal orifice decreases. The spray images showed that no liquid film (liquid sheet) existed, but an obvious liquid band was formed on the impact surface, which increased the transverse thickness and was not conducive to the propellant atomization. Furthermore, an increase in injection pressure caused a direct increase in impinging angle and a decrease in impinging distance. Among the four injectors, the SP1 injector has a shorter impinging distance and a larger impinging angle. The results presented in this paper may provide valuable reference for future hot-firing tests.

Comparative study of failure mechanisms of MoSi2 coating on Mo1 wire mesh under isothermal oxidation and hot-fire test using a hydroxylammonium nitrate based monopropellant

A MoSi2 coating was prepared on the Mo1 wire mesh via pack cementation method, and its failure mechanisms under isothermal oxidation and hot-fire test using a hydroxylammonium nitrate based monopropellant were comparatively studied. Under isothermal oxidation at 1300 °C and 1400 °C, degradation of MoSi2 into Mo5Si3 caused failure of the coating, and interdiffusion made a much larger effect relative to oxidation. However, the MoSi2 coating failed because of the synergy of oxidation, ablation, and interdiffusion under hot-fire test. Besides, dissolution of mullite into SiO2 and ablation of high velocity flame contributed to the failure of the coating as well.

Closed Loop Throttle Control of a Liquefying Fuel Hybrid Rocket Motor

The thrust produced by a hybrid rocket motor (HRM) can be controlled by varying the oxidizer flow rate to the combustion chamber. This feature is useful in shaping motor thrust profiles and optimizing a vehicle flight trajectory, but propellant throttling in solid-fuel hybrids is limited to the liquid component only, complicating the control scheme and potentially destabilizing combustion in the motor. While hybrid motor throttling ability remains a subject of considerable interest, there has been little investigation of throttling in motors that use high regression rate, liquefying fuels such as paraffin wax. This article describes the development and implementation of a closed loop thrust control scheme for a laboratory-scale paraffin wax/nitrous oxide HRM using a low-cost ball valve as the controlling hardware element. A model of motor performance is first developed from which proportional-integral-derivative (PID) controller constants are obtained through experimental testing. The control scheme is demonstrated through closed loop hot fire tests of a laboratory-scale motor in which thrust tracks a set-point value with feedback provided through a load cell. Upon reaching the setpoint, the motor remains throttled within ± 2.4% of the maximum thrust of the motor. Constant and ramping thrust profiles are demonstrated.

Fundamental Study on Injector Flow Characteristics of Self-Pressurizing Fluid for Small Rocket Engines

Abstract Nitrous oxide is a suitable propellant for rocket engines and has been widely used in various countries, given its high saturated vapor pressure, which enables self-pressurization. Because nitrous oxide is in a state of vapor–liquid equilibrium in tanks, it is easy to form a gas–liquid two-phase flow by cavitation in feed line. Since accurately estimating the performance of rocket engines requires evaluating the characteristics of propellant flows, tests reported in this paper were conducted using hybrid rocket engines under three conditions: cold flow test, hot firing test at low back pressure, and hot firing test at high back pressure. With consideration to the subcooling degrees, nitrous oxide may be in an unsteady superheated state in the upstream flow of the injector. In a comparison of the pressure ratios between the injector in each test condition, it is observed that a critical two-phase flow was formed in the injector in the cold flow test and in the low backpressure firing test. In the high backpressure hot firing test, the injector flow may be choked, but the large oscillations were observed in chamber pressure and thrust. According to the FFT analysis results, these oscillations were caused by chugging and acoustic oscillation. In light of these experimental results, it is suggested that when the chamber pressure fluctuates due to combustion instability such as chugging and acoustic oscillation, it may affect the injector flow characteristics and the critical two-phase flow.

Copper-based alloys for structural high-heat-flux applications: a review of development, properties, and performance of Cu-rich Cu–Cr–Nb alloys

ABSTRACT This review examines the development and current state of Cu-rich Cu–Cr–Nb alloys commonly referred to as GRCop or Glenn Research copper alloys with emphasis on Cu–8Cr–4Nb (at%), or GRCop-84, and Cu–4Cr–2Nb, or GRCop-42. Recent additive manufacturing efforts have increased interest in GRCop alloys, and full-scale hardware has been fabricated using AM techniques and practical hot-fire tests have been conducted, but structure–property relationships are still under development. The development, processing, and current microstructure-property relationships of GRCop alloys are reviewed along with comparisons to similar high-heat-flux Cu alloys including NARloy-Z, GlidCop Al-15, AMZIRC, Cu–1Cr–0.1Zr, and Cu–0.9Cr. The review concludes with an assessment of future prospects for GRCop alloys and overview of advantages provided by additive manufacturing.