Static Firing Tests Research Articles

Comprehensive approach to static firing tests of micro gas turbine engines powered by liquid fuels

The paper presents test results on the combustion of liquid fuels in micro gas turbine engines (MGTE). A setup was designed and built to determine the following MGTE characteristics: thrust, inlet static pressure, compressor static pressure, compressor total pressure, combustor total pressure, turbine total pressure, turbine speed, and the air temperature at the inlet, inside the compressor, at the turbine outlet, and at the nozzle exit. Thermodynamic cycle analysis of MGTEs was performed in the operation mode providing the required thrust. Noise and vibration levels of the MGTE operation were measured using a noise and vibration analyzer. Concentrations of anthropogenic emissions (CO, CO2, NO, NO2, N2O, SO2, CH4, C3H8) were recorded using a set of sensors. The tests revealed consistent patterns in the variation of throttle characteristics, gas composition, noise, and vibration in different engine operation modes. The test results were processed to produce plots of the key MGTE characteristics. Recommendations were developed for testing new types of fuels in MGTEs.

Development of hydrogen peroxide based micro thruster utilizing biomass-based catalyst bed

Micro thrusters, also referred to as micro-newton thrusters, are thrusters having a wide range of force capabilities in the micro-newton range. Micro thrusters are used in CubeSats, tiny robots and drones, micro- and nanosatellites, spacecraft attitude control, etc. The aim of the present study is to develop a monopropellant thruster (MPT) for 90 % concentrated hydrogen peroxide (H2O2). A rotatory evaporator was used to obtain a 90 % H2O2 form the available concentration of 50 % H2O2. The next stage involved the designing of the monopropellant thruster utilizing the NASA CEA rocket performance algorithm and the derived mathematical model. The test facility was constructed which includes a thruster, a feed line system, a static test firing platform, a data collecting system and the measuring sensors. To measure the thrust of the thruster, a pendulum thrust mechanism was constructed and fabricated into an experimental test stand. Lastly, the MPT chamber, which uses stainless steel mesh, was prepared and packed with the biomass-based catalyst. The 14 bar injection pressure conditions were used for the fire tests. With a mass flow rate of 5 g/s and a decomposition pressure of 10 bar, the theoretical specific impulse (Isp) was calculated to be 1993.2 Ns/kg at an O/F ratio of 0.002. The test firing with 90 % concentrated hydrogen peroxide gave an experimental specific impulse of 794.5 Ns/kg, which gives efficiency of 39.9 %.

Evaluation of solid propellant burning rate by using strand burner and static firing tests

Abstract The main objective of this paper is to compare the different methods used to measure the burning rate of hydroxyl-terminated polybutadiene (HTPB) propellant. The burning rate measurement is carried out under varying crosslinking density, which is determined by the curing ratio, i.e., the equivalent ratio of diisocyanate to total hydroxyl (NCO/OH) ratio. Four batches with different curing ratios were produced using fixed production processes. To measure the burning rate, cured solid propellant strands were tested using the acoustic wave emission method at different pressures ranging from 4 to 10 MPa and also by using static firing test of subscale motors. The study showed that there was little difference between the results obtained from the acoustic wave emission method and subscale motors.

Experimental Investigation on Transmission of Sound in To Metallic Enclosures During Static Firing Tests Of Solid Rocket Motor

Satellite launch vehicle during lift-off generates intense dynamic environment mainly the acoustic noise, which in turn acts on the nearby launch pad systems. Reliable operation of these launch pad ground systems needs to be ensured by safeguarding against these dynamic loads with adequate protection measures. For future launch vehicle programs where the launch turnaround time is a crucial aspect, sensitive systems like ground check-out electronic packages are planned to be placed adjacent to the launch vehicle mounting pedestal on the launch pad to ensure minimal time for preparation. The enclosures housing these electronic packages needs to be configured to minimize the transmission of acoustic levels being imposed from the lift-off induced noise. In the present work, an experimental investigation is carried out on acoustic transmission characteristics of metallic enclosures of different material and thickness during the static firing tests of solid rocket motors. Acoustics generated from the supersonic jet noise has a wide frequency band from 100 to 10,000 Hz with sound power peaking near 1,000 Hz. The electronic components and packages, which are generally vulnerable to the high frequency noise, need to be protected from this lift-off induced acoustics. The enclosures are placed in the far field of this jet noise and the transmission loss measurements are obtained. Variation of transmission loss with material and thickness of the enclosures are studied with specific attention to the internal acoustic modes and the coincidence phenomena. Subsequently, the enclosures treated with acoustic blankets of different thickness are investigated to verify its influence on internal acoustic modes and also the additional reduction in transmission of sound in the dominant frequency band of supersonic jet noise.

Thermal Tests of Type B Radioactive Transport Package: Experimental Description and Heat Flux Analysis

ABSTRACTThe design of Type B radioactive transport package should be able to withstand the fire accident scenario in transportation accident conditions. For the simulation of the fire scenario, there are mainly three methods: pool fire test, furnace test and thermal radiation test. This paper tried to carry out the fire test with the furnace method that the conditions in the furnace can be more carefully controlled. The paper described the requirements of fire test, test conditions and measurement test methods and analysed the temperature change of fire site and the heat flux data of the surface of the package during the fire test. The test results seemed that the fire test method could provide satisfactory thermal environment for fire test of Type B radioactive transport package.

STUDY OF STRENGTH REDUCTION OF REINFORCED CONCRETE BEAM UNDER FIRE TESTS

The article presents the results of studies of the regularity of changes in the concrete strength of reinforced concrete beams depending on the heating temperature of its inner layers under fire conditions by interpreting the temperature and deformation measurement results. To obtain the data necessary to identify patterns of concrete strength reduction, we developed a method of fire tests that substantiates the experimental base, design, and number of samples, type, number, and location of measuring devices, as well as the method for statistical processing of measurement data. Based on the developed methodology, we conducted fire tests. As a result of the tests, we obtained a set of measurement results in the form of time dependences of temperature indicators at individual cross-section points and the dependence of the maximum beam deflection on time, which are necessary to identify the coefficients of reduction in concrete strength of reinforced concrete beams under fire test conditions. We then used the proposed method to identify regularities in the reduction of concrete strength of a reinforced concrete beam, which includes interpolation using the obtained temperature indicators at individual points of the cross-section of the beam, which allows us to set the temperature at any point of the cross-section at any control time of measurements based on the approximation of isotherms by parabolas with a variable indicator of their degree. At the next stage, to identify the dependence of the concrete strength of reinforced concrete beams, we used a method based on the compilation of systems of equilibrium equations of internal layers in the cross-section of a reinforced concrete beam, in which the unknown values ​​are the coefficients of reduction of concrete strength for a range of temperatures every 100 °С in the range from 20 °С up to 1200 °С, provided that the reduction factor for the specified heating temperature of the inner layer of concrete is determined by linear interpolation. Using the proposed approach, we obtained appropriate results for a reinforced concrete beam subjected to fire tests. The obtained dependences of the strength reduction coefficient were used to evaluate the fire resistance of reinforced concrete beams and crossbars by the calculation method based on the deformation model. By comparing the obtained data with the data from the literature, we established the adequacy of the calculated data based on the relative deviation and Fisher’s test. The totality of the obtained data allows for improving the calculation method for evaluating reinforced concrete crossbars and beams by increasing the accuracy of the data obtained. Keywords: reinforced concrete beam, fire tests, concrete strength reduction coefficient, fire resistance.

Improved performance of data-driven simulator of liquid rocket engine under varying operating conditions

Integrating a model-based and data-driven approach to generate sufficient number of training data that include time-varying operating conditions is needed to develop a prognostics and health management technique using machine learning for a reusable rocket engine. A data-driven approach was used to predict the engine behavior, which cannot be determined using a model-based approach. This paper proposes a clustered regression model to predict various operating conditions for a reusable rocket engine. The training data were divided in advance according to the operating conditions using the Gaussian mixture model, and multiple regression models were individually trained on each cluster. By switching the regression models according to the operating conditions, the time-series data comprising various operating conditions, such as the combustion and chill-down phases and transition between phases, were accurately predicted when compared with the prediction results of a single regression model. The prediction results of the present hybrid approach agree reasonably well with the static firing test results of a reusable rocket engine if the explanatory variables are adequately included. By integrating model-based and data-driven approaches, adequate training data can be generated for prognostics and health management of the reusable liquid rocket engine.

Propulsion theoretical and experimental analysis of composite propellants motors

Rockets have revolutionized space technology and human space exploration. Most rockets and missiles are both propelled by rocket motors that use composite solid propellants. The ICT code and the NSAS CEA code are two programs that can be used to forecast theoretical propulsion parameters for composite solid rocket propellant. Rocket propellant performance is governed by a specific impulse factor, which is calculated theoretical and experiment. In this paper, the theoretical specific impulse for different composite solid propellant formulations at 70 bar combustion pressure and an adapted nozzle (optimum expansion) were calculated by the NASA-CEA code and the ICT code. Meanwhile, a static firing test was performed on a small scale test motor to experimentally determine the actual specific impulse. The objective is to verify theoretical calculations from two codes with experimental data, via the determination of the specific impulse deviation co-efficient.

A novel in-flight thrust modulation technique for SPRM: proof of concept

Thrust modulation (termination or reduction) is one of the most sophisticated topics in rocket motors. Thrust termination is employed when combustion must be terminated to ensure proper stage separation, to avoid motor explosion, to attain certain range, or for research purposes. While liquid and hybrid propellant rocket motors have the ability to extinguish thrust, thrust termination in solid propellant rocket motors (SPRMs) need the use of specialized technology. In-flight thrust termination of (SPRMs) can be achieved by sudden and sufficient increase in throat area. Increasing throat area causes high depressurization rate which consequently leads to extinguishing the rocket motor. In contrast, in-flight thrust reduction can be theoretically attained either by reducing chamber pressure to a new equilibrium level or by reducing the divergent section length. The objective of the present paper is to prove the concept of using Linear Shaped Charge (LSC) for thrust modulation; a technique novel to SPRM applications. Thrust termination concept is proved via a set of static firing tests on a standard test motor. In addition, a mathematical model is developed to predict the drop in thrust due to cutting the nozzle at any arbitrary location. The model is validated by comparing with published experiments for thrust termination and reduction. A parametric study is conducted based on model results. It is confirmed that by cutting the nozzle along its divergent section, thrust reduction less than 20% can be attained. Further reduction can be achieved by cutting the nozzle along its convergent section.

A Review on the Influence of Test Bed Dynamic Characteristics on Thrust measured during Static Fire Testing of Solid Rocket Motor case

Abstract: Solid propelled rocket motors (SRMs) static firing tests are crucial tests in the aerospace industry while developing new motors as well as to ensure the quality of a motor batch. The operator can use this sort of test to determine whether the motor performance meets the project criteria by obtaining the measured "thrust versus time of burning" graphic the motor produces while burning. In this paper over all thrust measurement uncertaintyof a solid propellant rocket motor test bed is studied. During static test of solid motors requires the dynamic characteristics of the Static test configuration. To find out the magnitude of thrust oscillation, characterization of test setup and force transfer characteristics from motor case to load cell are required.

High Heat Resistance Can Be Deceiving: Dripping Behavior of Polyamide 4.6 in Fire

AbstractPolyamide 4.6 (PA46) is a high‐heat‐resistant polymer, but it has no dripping resistance under fire. Three commercial grades of PA46 are investigated under UL 94 vertical fire test conditions. Their performances are discussed based on the materials' structural, thermal, and rheological properties. PA46 presents flaming drops, whereas dripping is prevented in the flame‐retarded PA46. Friction‐modified PA46 has increased flaming dripping. Temperature profiles of the specimens under fire and the temperature of the drops are measured by thermocouples. A UL 94 vertical test configuration consisting of two flame applications is designed to assess the quantitative dripping behavior of the set of materials by the particle finite element method (PFEM). Polymer properties (activation energy and Arrhenius coefficient of decomposition, char yield, density, effective heat of combustion, heat of decomposition, specific heat capacity, and thermal conductivity) in addition to rheological responses in high temperatures are estimated and measured as input parameters for the simulations. The dripping behavior obtained by simulated materials corresponds with the experimental results in terms of time and drop size. A consistent picture of the interplay of the different phenomena controlling dripping under fire appears to deliver a better understanding of the role of different materials’ properties.

Measurement of alumina film induced ablation of internal insulator in solid rocket environment

This study investigates the ablation of internal insulation induced by the deposition of an alumina film with different lateral film speeds. A sub-scale test solid rocket motor (SRM) was designed in an impinging jet configuration to form an alumina film on the sample and to encourage the lateral movement of the film by a high-speed wall jet. Fifteen static fire tests of the test SRM were conducted with six different jet velocities (Vjet = 100 m/s, 150 m/s, 200 m/s, 268 m/s, 330 m/s, and 450 m/s) that indirectly affected the velocity of the wall jet and the deposition rate of alumina droplets. The ablation velocity was deduced from the difference in the sample thickness after a test using a coordinate measuring machine. The droplet deposition mass flux and wall jet velocity were obtained via two-phase flow simulation with the same jet velocity and effective pressure. As a result, the characteristics of alumina-induced ablation and the changes in ablation with jet velocities were obtained. The area within 0.8 × jet diameter was focused upon, where the ratio of ablation velocity to incoming alumina mass was constant for each jet velocity, and showed a similarity in jet structure. When the ablation velocity was increased from 2.05 to 9.98 mm/s with increasing jet velocity, the ratio of the ablation velocity and alumina mass flux decreased from 1.07×10−4 to 0.49×10−4 m3/kg as Al2O3–C reactions became less efficient with a reduced residence time of the film. Because the decrease in residence time by the wall jet is more pronounced for slow reactions involved in Al2O3–C reactions, fast reactions in Al2O3–C reactions are less affected and result in a convergence of the volumetric rate of ablation per unit mass of alumina.

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.

Novel Testing of a Water-Breathing Naval Ramjet at Underwater Cruise Conditions

This paper presents, for the first time, firing tests and results measured during a hybrid-propellant marine water-breathing ramjet operation at underwater cruise conditions. The firing test demonstrated ignition under water, stable combustion, and a significant increase in chamber pressure due to the introduction of water by the ram effect. The exclusive test facility used to accommodate this unique experiment is described along with a detailed explanation of the motor design. In addition, a theoretical model to predict the performance of the motor is shown along with the numeric process used to solve the model. Measurements taken during the dynamic firing test reveal good compatibility with the theoretical model. The theoretical model was also used to evaluate values that could not be measured during the experiment. The results found during this work fit previous results attained by static firing tests, demonstrating significant improvement in specific impulse due to addition of water. An improvement of 50% in specific impulse was demonstrated when comparing the results of an underwater dynamic firing test of a marine ramjet to those of a parallel rocket test conducted under identical conditions in the water basin.

Student firing tests and launches with commercial and self-made solid rocket motors

This paper presents work performed by students at Sapienza University of Rome regarding on-ground testing of solid rocket motors and launch of single-stage rocket models. The activities are part of a multidisciplinary project, in collaboration with the Italian Space Agency, aimed at the design, test, build, and launch of model rockets to be used as technological demonstrators and for educational purposes. Student activities included: (i) building of electronic components; (ii) designing and building of an experimental test bench; (iii) manufacturing of rocket parts and fuels; (iv) static firing tests of commercial and self-made solid rocket motors; (v) rocket launches. Self-made solid propellants were obtained as a mixture between potassium nitrate and sorbitol, yielding chamber pressures up to 20 bar and thrust up to 20 N, while commercial motors with thrust up to 90 N were tested. Additive manufacturing was employed for building specific rocket parts such as the nosecone, fins, and videocamera housing. The payload consisted in self-programmed board computers which recorded telemetry during flight. A total of four rocket launches up to 540 m have been performed.

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.

Static Firing Tests of Solid Propellant Rocket Motors: Uncertainty Levels of Thrust Measurements

Static firing tests of solid propelled rocket motors (SRMs) are important tests in aerospace industry during the development of new motors as well as to assure the quality of a motor batch. This type of test allows the operator to obtain the measured “thrust versus time of burning” graphic yielded by the motor during burning to verify whether the motor performance matches the project requirements. Laboratory Quality Certificates and good measurement practices make necessary the evaluation of the typical uncertainty of these test results, which demand a wide number of instruments, called measurement chain. This work aims to present the uncertainty levels expressed to the thrust values gathered during a SRM static firing test for two differently configured measurement chains. In order to estimate the uncertainty of a complete measurement chain, the theories of expanded and combined uncertainties are employed. It was possible to verify that in the analyzed case the use of signal amplifiers increased the measurement uncertainty of the chain.

Investigation of a Marine Water-Breathing Hybrid Ram-Rocket Motor

This work presents theoretical and experimental performance studies and a general scheme of an innovative concept for propelling marine high-speed underwater vehicles, a water-breathing hybrid ram rocket. The water-augmented hybrid ram rocket enables obtaining superior performance relatively to a standard rocket and improved performance relatively to a parallel water–augmented solid ram rocket. In addition, it demonstrates improved safety and controllability due to the hybrid motor characteristics. The theoretical model incorporated data from a thermochemical program. The experimental investigation was based on static firing tests of hybrid rocket motors augmented by the addition of water during combustion. The study focused on the use of a non-water-reactive propellant combination, cured polyester as fuel and gaseous oxygen as oxidizer. Wide ranges of oxidizer-to-fuel and water-to-propellant ratios were considered. The experimental results showed a significant improvement (up to 70%) in specific impulse, relatively to a non-water-augmented hybrid rocket, in good accordance with the theoretical results. It is concluded that for standard chamber pressure of 6.9 MPa, corresponding to the ram pressure for cruising at , a specific impulse greater than 400 s can be achieved by a water-augmented hybrid ram rocket using a non-hydro-reactive fuel.

Experimental investigation of the axial oxidizer injectors geometry on a 1-kN paraffin-fueled hybrid rocket motor

The showerhead is the most common type of injector used in hybrid rocket motors due to its design and manufacture simplicity. The main drawback of the showerheads (SH) is the relatively low performance in terms of combustion efficiency because of the poor atomization of the liquid oxidizer. This study investigates the problem to enrich the technical literature by presenting detailed experimental data on showerhead injectors by a series of static firing tests using a 1-kN lab-scale hybrid rocket, applying four kinds of showerhead injectors, named SH1 (benchmark), SH2, SH3, and SH4. They differ from each other by the number, distribution, and dimension of the orifice elements. Main performance parameters such as fuel regression rate, specific impulse, and combustion efficiency are experimentally obtained.Two different series of tests were carried out. At first, the influence of SH injector geometries was studied. The injectors SH2, SH3, and SH4 were used under the same initial conditions, and the results were compared with SH1. In the second set of tests, the SH4 injector was chosen, and the effects of the fuel grain initial port diameter were investigated. Using the data obtained in the second battery of tests was possible to determine the influence of the fuel port diameter on the motor efficiency and establish the regression rate law of nitrous oxidizer/paraffin-based fuel for this specific configuration.

A Computational Tool for the Design of Hybrid Rockets

A computational tool able to perform a fast analysis of hybrid rocket engines is presented, describing briefly the mathematical and physical models used. Validation of the code is also shown: 16 different static firing tests available in the open literature are used to compare measured operational parameters such as chamber pressure, thrust, and specific impulse with the code’s output. The purpose of the program is to perform rapid evaluation and assessment on a possible first design of hybrid rockets, without relying on computationally expensive simulations or onerous experimental tests. The validated program considers as benchmark and study case the design of a liquid-oxygen/paraffin hybrid rocket engine to be used as the upper stage of a small launcher derived from VEGA building blocks. A full-factorial parametric analysis is performed for both pressure-fed and pump-fed systems to find a configuration that delivers the equivalent total impulse of a VEGA-like launcher third and fourth stage as a first evaluation. This parametric analysis is also useful to highlight how the oxidizer injection system, the fuel grain design, and the nozzle features affect the performance of the rocket.