Static Firing Tests Research Articles

Scaling of catalyst bed for hydrogen peroxide monopropellant thrusters using catalytic decomposition modeling

Hydrogen peroxide or H2O2 has been widely used as a green propellant in various types of thrusters based on catalytic decomposition. However, no standardized design methods for H2O2 catalyst beds exist at present. In this study, we scaled a catalyst bed for a H2O2 monopropellant thruster using catalytic decomposition modeling. We adopted the model developed by Pasini et al. (2010) [1] to simulate the catalytic decomposition of H2O2 in the catalyst bed. The one-dimensional profiles of temperature, pressure, species concentration, and other properties were estimated using the model. Additionally, static firing tests were conducted under various conditions using a 100-N H2O2 monopropellant thruster with a MnO2/PbO/Al2O3 catalyst. The simulation results were compared with those obtained from the static firing tests to validate the accuracy of the model. We observed that the temperature estimations from the model concurred with those of the experimental data; however, pressure estimations deviated slightly. Furthermore, the model was used to obtain the design parameters for scaling based on catalyst capacity and pressure drop analyses under various conditions. We determined that the pressure drop can be scaled and expressed as a constant. Thus, the catalyst bed can be scaled precisely by analyzing the catalyst capacity and pressure drop constants through catalytic decomposition modeling.

Investigation of the Ablative Properties of an EPDM\Kevlar Insulator in a Solid Rocket Motor

AbstractA study of the influence of the reinforced fiber orientation and the propellant aluminum content on the ablation and erosion performance of Ethylene Propylene Diene Monomer type insulation is presented. A subscale solid rocket motor has been used, with ammonium perchlorate and hydroxyl‐terminated polybutadiene propellant composition. Seven static firing tests were conducted: six with low Al and one with high Al content. The EPDM articles had four different Kevlar fiber orientations: circumferential, longitudinal, diagonal (45°), and random. Numerical simulations were supplemented to decipher the experimental outcome. The numerical results exhibit an over‐prediction tendency, yet they show that the examination of post‐fire charring of the insulation is insignificant. The test results indicate that the longitudinal fiber orientation has the least ablative resistance.

Design and Development of DALBIT-1: Earth-Based Lander Test Bed

DALBIT-1, a miniature-sized, fully integrated, Earth-based lander test bed was designed and developed for the vertical takeoff and vertical landing flight demonstration. Mission and design requirements and vehicle subsystems were selected to accommodate the vertical takeoff and vertical landing preliminary flight demonstration. Subsystems, including the propulsion system, structure, avionics, and ground control station, were designed, developed, and integrated as a system. The flight control system of DALBIT-1 was generated to enable the vehicle an autonomous flight and was validated using a computer simulation in a MATLAB/Simulink environment. Simulation results show that DALBIT-1 adequately follows the predetermined trajectory and softly lands on the surface. The developed test bed performed a series of static fire tests and water/propellant flow tests to enhance the predictability of the flight control system. Using the developed flight control system, finally, the preliminary flight demonstration was conducted and proved the capability of the test bed of performing a vertical liftoff, vertical landing flight.

Evaluation of the high-frequency oscillation parameters of a liquid-propellant rocket engine with an annular combustion chamber

The high-frequency instability (HF instability) of a liquid-propellant rocket engine (LPRE) during static firing tests is often accompanied by a significant increase in dynamic loads on the combustion chamber structure, often leading to the chamber destruction. This dynamic phenomenon can also be extremely dangerous for the dynamic strength of a liquid-propellant rocket engine with an annular combustion chamber. Computation of the parameters of acoustic combustion product oscillations is important in the design and static firing tests of such rocket engines. The main aim of this paper is to develop a numerical approach to determining the parameters of acoustic oscillations of combustion products in annular combustion chambers of liquid-propellant rocket engines taking into account the features of the configuration of the combustion space and the variability of the physical properties of the gaseous medium depending on the axial length of the chamber. A numerical approach is proposed. The approach is based on mathematical modeling of natural oscillations of a “shell structure of an annular chamber – gas” coupled dynamic system by using the finite element method. Based on the developed finite-element model of coupled spatial vibrations of the structure of the annular combustion chamber and the combustion product oscillations, the oscillation parameters of the system under consideration (frequencies, modes, and effective masses) for its dominant acoustic modes, the vibration amplitudes of the combustion chamber casing, and the amplitudes of its vibration accelerations can be determined. The operating parameters of the liquid-propellant rocket engine potentially dangerous for the development of thermoacoustic instability of the working process in the annular combustion chamber can be identified. For the numerical computation of the dynamic gains (in pressure) of the combustion chamber, a source of harmonic pressure excitation is introduced to the finite element model of the dynamic system “shell structure of an annular configuration – gas” (to the elements at the start of the chamber fire space). The developed approach testing and further analysis of the results were carried out for an engine with an annular combustion chamber (with a ratio of the outer and inner diameters of 1.5) using liquid oxygen – methane as a propellant pair. The system shapes and frequencies of longitudinal, tangential and radial modes are determined. It is shown that the frequency of the first acoustic mode in the case of a relatively low stiffness of the combustion chamber casing walls can be reduced by 40 percent in comparison with the frequency determined for a casing with rigid walls.

Conjugate Analysis of Silica-Phenolic Charring Ablation Coupled with Interior Ballistics

Because of its excellent insulation capability, the usage of a silica-phenolic charring ablator as a nozzle liner is a common practice in the solid rocket motor industry. During the design of a solid rocket motor employing a silica-phenolic nozzle liner, it is desired to conduct an accurate analysis yielding in-depth thermal response and recession characteristics. As the interior ballistics and nozzle recession rate mutually interact, the best practice is to perform a coupled solution to both. Commonly used one-dimensional analysis tools with empirical approaches for estimation of convective heat transfer rate and blowing effect generally lack sought accuracy and do not model the transient shape-change phenomenon, which affects the nozzle performance. This Paper considers governing equations for charring, including pyrolysis gas injection and surface energy balance for melting ablation, along with a boundary condition governed by interior ballistics, and demonstrates a framework in which these equations are solved with governing equations for the nozzle flowfield in a coupled manner. Development and validation of a one-dimensional material response solver based on the same governing equations is also demonstrated. Also, results from a static firing test conducted with a small-scale ballistic evaluation motor employing a silica-phenolic nozzle insert are provided. Results from both investigations are compared and discussed.

A Supervised Framework for Recognition of Liquid Rocket Engine Health State Under Steady-State Process Without Fault Samples

Data-driven intelligent models have received more and more attention in liquid rocket engine (LRE) state recognition. However, due to practical limitations, the fault data are rare for model training, which results in poor testing results. To better recognize the state of LRE, we proposed a supervised recognition framework without fault samples. First, we proposed a negative sample generation method. Then, fusion recurrent convolutional neural network (FRCNN) is constructed for data fusion and feature extraction. Finally, we evaluated the recognition results with several quantization indicators. The proposed method is verified with the real monitoring data, which is obtained from the static firing tests of a certain type of LRE. Results show that the FRCNN could recognize the state of LRE with accuracy over 93% under the confidence of 96%, which shows the effectiveness of the proposed framework. And the comparison with the literature method shows the superiority of the proposed framework. In addition, it could also perform real-time monitoring for LRE, and recognize the abnormal at the initial stage of the steady state, which has practical potential in monitoring task. At last, we further discuss and analyze the generated negative sample and the proposed framework.

Data-driven fault detection in a reusable rocket engine using bivariate time-series analysis

Data-driven fault detection using a bivariate time-series analysis was done for the maintenance of reusable liquid-propellant rocket engines. This method is based on the phase plane trajectory of feature vectors extracted from two sensor data by principal component analysis, and applied to all sensor pairs. System and sensor failures can be estimated by visualizing the state of all sensor pairs. The present method was applied to the static firing test results of a reusable rocket engine developed in Japan. Temperature sensor failure was successfully detected from 59 sensors in the 19 static firing tests. System failure caused by incorrect valve operation was also successfully detected. The ability to detect and estimate system and sensor failures was demonstrated, even if the engine's operational sequence changes dynamically due to ignition or cutoff or if there are deviations of the engine's operational sequence between tests.

Regression rate of paraffin-based fuels in hybrid rocket motor

Paraffin-based fuels are regarded as a promising solid fuel for hybrid rocket motor (HRM) owing to high regression rate, excellent mechanical performance, and high dropping point. The regression rate of paraffin-based fuels was characterized by static fire test with nitrous oxide (N2O) and gaseous oxygen (GOX) as oxidizer, and the effect of combustion catalyst, injector and chamber pressure on the regression rate was evaluated in this study. The results show that the average regression rates of paraffin-based fuel with GOX and N2O are similar under the optimal oxidizer/fuel (O/F) ratio, and the exponent n is larger with GOX as the oxidizer, indicating high sensibility to oxidizer mass flux. The average regression rate increased with the increase in swirling number Sg, and the tangential injector with Sg of 6.83 improved the regression rate significantly. The local regression rate unchanged with axial position when a jet injector was used, but a complex local regression rate appeared at the first half of fuel grain for tangential injector. Both copper chromite and cobalt stearate obviously increased the regression rate; however, only copper chromite could be used as a preferable regression rate improver owing to significant curing acceleration of fuels by cobalt stearate. The average regression rate slightly increased with the increase in chamber pressure with GOX as the oxidizer, but a noticeable positive relationship could be found for the cases of N2O.

Estimating model parameters of liquid rocket engine simulator using data assimilation

To generate training data for fault detection and diagnosis based on machine learning, a System-Level Simulation (SLS) is developed to model the global behavior of a reusable liquid rocket engine employing the expander-bleed cycle. Model parameters are estimated automatically by an Ensemble Kalman Filter. Parameters are estimated starting from the turbopump component model, followed by the whole rocket engine model, to increase accuracy. Sufficient convergence is obtained within one day by parallelizing the SLS computation for the ensemble members by two. Compared to the static-firing test results, reasonable agreement is obtained within error of 2.3% in the steady-state engine condition. Even for the startup and shutdown sequences, the present SLS reasonably agrees with the test results except for the temperature. Comparing the model parameters obtained for different static-firing test results, the effect of the overhaul inspection on the turbopump is captured as a difference in the model parameters.

Enhancement of the Fuel Regression Rate in Hybrid Propulsion by Expandable Graphite Additive

ABSTRACT Hybrid rockets can be an alternative to solid and liquid propellant systems, having improved safety, controllability, and high energetic performance. One of the features characterizing hybrid propulsion is a relatively low thrust due to low regression rate of the fuel. This may be a drawback for certain missions. Fuels with a higher regression rate while maintaining good mechanical properties may be of great advantage. This article investigates experimentally and theoretically a new direction to achieve this goal: the addition of expandable graphite (EG) to the fuel. EG exists as fine particles (typically, 100–400 µm). At elevated temperatures, they expand and form strings several times longer than the original particles. When mixed with a fuel matrix they can increase the heat transfer into the fuel via conduction through the EG strings, thus increasing the regression rate. Static firing tests employing gaseous oxygen as an oxidizer and either polymeric (polyester) or liquefying (paraffin wax) fuels (plain or with 5% EG additive) have been conducted. The test results revealed an increase of 30–200% in the fuel regression rate due to EG addition, with a larger effect on the polymeric fuel and a lesser effect on the paraffin wax. Theoretical model predictions demonstrated good agreement with the experimental results.

Investigation of Graphite Nozzle Erosion in Hybrid Rockets Using Oxygen/High-Density Polyethylene

A recently developed reconstruction technique is used to investigate graphite nozzle erosion in two scales of hybrid rocket motors, 30-N-thrust class and 2000-N-thrust class, using oxygen as the oxidizer and high-density polyethylene as the fuel. Thermocouple measurements taken from within the nozzles are used to estimate nozzle throat wall temperature. Forty-four static firing tests were conducted under varying experimental conditions to confirm the validity of the reconstruction technique results, to investigate the conditions at the onset of erosion, and to formulate an empirical predictive model of nozzle erosion rate. Results show that a single formula that treats the combustion gas as a single oxidizing agent for which heterogenous rate constants are functions of equivalence ratio can satisfactorily replicate the erosion rate of graphite by a combustion gas containing multiple oxidizing species. Furthermore, the chemical-kinetic-limited conditions of the onset of nozzle erosion are specified by a novel empirical correlation, which shows that erosion begins at lower temperature and pressure in oxidizer-rich combustion gas than in fuel-rich combustion gas.

Systems for assuring safety of static firing tests of oxygen/hydrogen propulsion units

Static firing tests of oxygen-hydrogen propulsion units with propellant tank load of up to 2700 kg of hydrogen are run with special safety precautions and using emergency test protection system covering up to 60% of contingencies. Test on propulsion units with higher hydrogen propellant loads (up to 7000 kg) require additional safety measures, based on improved effectiveness of the fire and explosion prevention systems, emergency protection, and the coverage ratio for emergency (off-nominal) situations. The key factors for improving off-nominal situations coverage ratio during tests are research into systems for earlier detection of hydrogen leaks and the use of inhibitors with nitrogen as the phlegmatizing agent to prevent explosion of the hydrogen/air (oxygen) mixtures. The paper presents results of studies into systems for early detection of hydrogen leaks and design of monitoring equipment based on fiber optics and microcavity structures that have higher speed of response, selectivity, immunity to noise and reliability. Key words: propulsion system, safety, emergency protection system, off-nominal situation, hydrogen leak, fiber-optic technology.

Integration validation of key components for small sounding rockets

To develop sounding rockets that can be used for various missions, a university, research center, and two enterprises collaborated to conduct a critical-system development project on a small-scale sounding rocket. The Korea Advanced Institute of Science and Technology (KAIST) Rocket Laboratory and Space Solution Co., Ltd. collaborated to develop a hybrid rocket and a propellant-feeding system using hydrogen peroxide as an oxidizer. The KAIST Satellite Technology Research Center has developed a non-pyrotechnic recovery system, and Naraspace Technology Co., Ltd. has developed rocket avionics. Using a 2.39-m technology demonstrator that weighs 16.4 kg, the developed system has been validated at low altitudes, and the applicability of the small-scale sounding rocket has been evaluated. The combustion stability, short ignition delay, and high performance of the H2O2 hybrid rocket were confirmed through static firing and flight tests. The reliability and capacity of the non-pyrotechnic recovery system (which uses a spring) were confirmed by ground and flight tests. Perfect operation of the developed rocket avionics, such as control of the main oxidizer valve and recovery system, emergency functions, data transmission, and data storage was confirmed through ground and flight tests.

Evaluation of Processing Parameters for Densification of Composite Propellants

The present work investigated the effects of varying two parameters of a pilot-scale composite propellant plant: the mechanical vibration and the vacuum level. The application of mechanical vibration to the casted mold after the end of mixing has improved the propellant density. On the other hand, the change of the vacuum level had no significant effect. Static firing tests showed an increase in the gas generation rate with the increase of the number of voids.

Ballistic Protection of Military Shelters from Mortar Fragmentation and Blast Effects using a Multi layer Structure

In addition to its usage as a top attack ammunition in the battle field against troops, 120 mm mortar bomblethas been recently used in the terror attacks against military shelters and civil constructions. This research studiesthe protection of vehicle and personnel military shelters against mortar fragmentation warhead projectile and it’sdestroying effects. The mortar warhead threat combines both blast load and ballistic fragment penetration effects. Composite structure layers are proposed herein to be integrated with concertina walls to achieve full protection against the mortar projectile destroying effects. The current paper investigates the ability of the proposed layers to stop the mortar’s fragments and to mitigate its blast load. The velocity and the mass distribution of the produced projectile fragments were estimated using Split-X software. Besides, the ability of the proposed protection layers to stop the fragments and mitigate the blast wave was evaluated using AUTODYN hydrocode. A static firing test was then performed to validate the theoretical results and verify the effectiveness of the proposed protection added layers. The current study showed that the proposed composite layers are sufficient to protect the military shelters from the mortar’s destroying ballistic effects.

Fault Detection of a Reusable Rocket Engine using Phase Plane Trajectory of Feature Vectors

A bivariate time-series analysis based on the phase plane trajectory of feature vectors extracted by principal component analysis is developed for fault detection in a reusable liquid-propellant rocket engine. Static-firing test results of the reusable rocket engine obtained at the Japan Aerospace Exploration Agency are employed for demonstration of the present method. The present method successfully detected temperature sensor failure from 19 firing tests of 62 sensors, even in the deviation of the engine operational sequence between the static-firing tests. The present method was also able to detect the system failure from 23 firing tests. Furthermore, the ability to distinguish the system and sensor failure was demonstrated.

Investigation of a Ducted Rocket with a Paraffin/Oxygen Hybrid Gas Generator

Analysis and testing of a ducted rocket employing a hybrid gas generator is presented. Over 150 successful static firing tests have been performed using a connected pipe facility supplying vitiated air to the ram combustor at stagnation temperature and pressure encountered in actual flights. The hybrid gas generator consisting of paraffin wax and gaseous oxygen operated unchoked, supplying hot, fuel-rich gases to the ram combustor for further burning with air. Empirical correlation of the fuel regression rate versus oxidizer mass flux in the gas generator has been displayed, and good overall combustion efficiency has been revealed. Thrust and fuel/air ratio control via the cold oxidizer flow rate into the gas generator is a major advantage of employing a hybrid gas generator. It was shown that such control could be done over broad ranges of flight altitude, flight Mach number, and thrust, while maintaining the ram pressure throughout the engine. This is in contrast to the thrust control of a ducted rocket employing a fuel-rich solid propellant. In that case, the process involves variation of the gas generator nozzle throat area by a mechanical device within the hot flow, implying a large increase of pressure in the gas generator combustor. Different scenarios of engine control and thrust modulation for variable flight conditions and requirements have been demonstrated.

Small-Scale Static Fire Tests of 3D Printing Hybrid Rocket Fuel Grains Produced from Different Materials

The last decade has seen an almost exponential increase in the number of rocket launches for sounding missions or for delivering payloads into low Earth orbits. The emergence of new technologies like rapid prototyping, including 3D printing, is changing the approach to rocket motor design. This project conducted a series of small-scale static fire tests of fused deposition manufacturing hybrid rocket motors that were designed to explore the performance of a variety of commonly available fused deposition manufacturing materials. These materials included acrylonitrile butadiene styrene, acrylonitrile styrene acrylate, polylactic acid (PLA), polypropylene, polyethylene terephthalate glycol, Nylon, and AL (PLA with aluminum particles). To test the performance of small-scale fuel grains, a modular apparatus with a range of sensors fitted to it was designed and manufactured. The small-scale testing performed static burns on two fuel grains of each material with initial dimensions of 100 mm long and 20 mm in diameter with a 6 mm straight circular combustion port. The focus of this study was mainly on the regression rates of each material of fuel grains. Acrylonitrile styrene acrylate and Nylon showed the highest regression rates, while the polyethylene terephthalate glycol regression rates were relatively poor. Also, the acrylonitrile butadiene styrene and acrylonitrile styrene acrylate demonstrating relatively high regression rates when compared to existing hybrid fuels like hydroxyl-terminated polybutadiene.

Comprehensive Data Reduction for N2O/HDPE Hybrid Rocket Motor Performance Evaluation

Static firing tests of a hybrid rocket motor using liquid nitrous oxide (N2O) as the oxidizer and high-density polyethylene (HPDE) as the fuel are analyzed using a novel approach to data reduction that allows histories for fuel mass consumption, nozzle throat erosion, characteristic exhaust velocity (c*) efficiency, and nozzle throat wall temperature to be determined experimentally. This is done by firing a motor under the same conditions six times, varying only the burn time. Results show that fuel mass consumption was nearly perfectly repeatable, whereas the magnitude and timing of nozzle throat erosion was not. Correlations of the fuel regression rate result in oxidizer port mass flux exponents of 0.62 and 0.76. There is a transient time in the c* efficiency histories of around 2.5 s, after which c* efficiency remains relatively constant, even in the case of excessive nozzle throat erosion. Although nozzle erosion was not repeatable, the erosion onset factors were similar between tests, and greater than values in previous research in which oxygen was used as the oxidizer. Lastly, nozzle erosion rates exceed 0.15 mm/s for chamber pressures of 4 to 5 MPa.

Mechanical integrity of gas turbine enclosure doors under fire test conditions for A0 fire rating

PurposeAs fire doors are passive fire protection parts, the doors have to be certified through standard fire tests. It is usual practice to perform the standard fire testing on the components which require the fire certification. However, some gas turbine enclosure doors are too large to test at the test facility and hence the fire resistance test is practically not possible. The purpose of this paper is to develop a reliable finite element model, validate the model using the specimen door test results and extend the method to actual gas turbine enclosure doors to support the fire certification.Design/methodology/approachFirst, the standard fire testing on enclosure door test specimen was carried out. Second, the finite element analysis model was built and tuned to match the standard fire test deflections, and finally, the same modelling technique was extended to model the actual gas turbine enclosure door to verify the results for fire certification process.FindingsGap analysis, a method of post processing is suggested for result analysis. It was found suitable to verify the gap openings which are required for A0 rated fire certification according to fire test procedure code and also to check the mechanical integrity of the enclosure door frame assembly.Originality/valueThe method presented in this work could be used as support information along with the test specimen results for A0 class fire rating certification of the doors according to International Maritime Organization Resolution MSC.307 (88) Annexure 1: Part 3.