Strand Burner Research Articles

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.

Survey of burning rate measurements in small solid rocket motors

Burning rate plays a crucial role in determining the performance of solid rocket motors (SRMs). In the traditional approach for solid propellant propulsion, technical activities regarding burning rates are developed at three different operational levels: (i) Strand burners or laboratory-scale devices in general; (ii) Small-scale motors (SSMs); And (iii) full-size or end-item rocket motors. While strand burners are extensively used for propellant development (formulation exploration, ingredients screening, performance verification, and production control) and relatively little is done experimentally at the full-size motor level (being large-scale experimentation too expensive and dangerous), a lot is usually carried out at SSM level to obtain burning rate information under motor operating conditions. In the introductory part of this work, burning rate fundamentals are recalled and burning rate measurement devices are quickly summarized. Then, a survey of subscale motors is conducted and several automated procedures to deduce burning rate from SSM testing are analyzed. Data reduction methods commonly used by leading European companies are based on the thickness-over-time (TOT) definition. Attention is dedicated to procedures used in Italy and France for quality control of the European space launchers (solid propellant boosters of Ariane family and core solid rocket motors of VEGA family). In addition, automated data reduction methods based on mass balance (MB) and often used in USA are investigated. Specific features and general trends of the tested industrial procedures are pointed out. Since for any tests the actual burning rates are unknown, results can only be compared based on the statistical quality of the deduced ballistic data. Mainly reproducibility, ease of application, and suitability for automated computer implementation are of interest to industrial users. The effects of test variability, input data quality, and data reduction methods on result reproducibility are discussed with reference to fire tests of the successful Ariane-5 solid boosters. The systematic analysis of industrial data conducted in this work suggests that improving the actual mix reproducibility and quality of experimental data is more important than perfecting the current data reduction methods. Moreover, the international trends suggest that the fundamental TOT procedures are being replaced by MB procedures or advanced TOT procedures with burning times evaluated using pressure integrals.

Extinguishing and Combustion Characteristics of Electrically Controllable Solid Propellants Under Elevated Pressures

An experimental study was conducted on electrically controllable solid propellants (ECSPs) created using a polyethylene oxide polymer binder, lithium perchlorate, and multiwalled carbon nanotubes. The propellants decompose and ignite shortly after the application of a voltage potential and extinguish when the voltage is removed under atmospheric conditions. The ignition delay as a function of the applied voltage magnitude was determined for a range of ECSP compositions. Pressurized experiments were conducted in an optically accessible strand burner to characterize the burning properties of the ECSPs as a function of pressure and electrical power. Additional experiments were conducted at elevated pressures where the voltage potential was removed and reapplied to extinguish and reignite the propellant and determine the self-extinction limits of the ECSPs. The results demonstrate that small compositional changes can drastically impact the ability to extinguish the ECSPs at elevated pressures.

Prediction of Burning Characteristics of Dihydroxyglyoxime Composite Propellant

An optimized engine start procedure is critical to the successful operation of a liquid rocket engine in launch vehicles. A solid propellant gas generator is widely adopted for the turbine starter during engine startup, and ammonium nitrate and ammonium perchlorate propellants are conventionally used for this purpose. However, these propellants have shortcomings such as high flame temperature, corrosive combustion residues, and low ignitability. In this study, a dihydroxyglyoxime (DHG)-based propellant was applied to turbine starters. The burning rate, characteristic velocity, and combustion temperature of the DHG propellant were evaluated using motor tests. The DHG-based propellant burned 3–11% slower in motor firing tests than that in strand burner tests, and an inversely proportional relationship was observed between the strand burn rate and the burning rate factor (ratio between motor burning rate measurement and strand burner prediction). The temperature sensitivity of the burning rate factor was found to be 0.23–0.24%/°C, and the pressure sensitivity of the characteristic velocity was 0.48–0.50%/MPa. These burning characteristics of the DHG-based propellant from static evaluations provide the evolution of the chamber pressure and the mass flow rate versus the time of the motor using internal ballistic analysis.

Lab-scale ballistic and safety property investigations of LMP-103S

LMP-103S is a storable, green, ammonium dinitramide (ADN)-based liquid monopropellant with significant potential owing to its improved specific impulse (≥ 6%) and density (∼24%) in comparison to hydrazine, and its superior safety and handling characteristics. LMP-103S already has a significant history of thruster testing and flight heritage. However, its baseline combustion behavior is a surprisingly understudied topic. Constant-volume strand burner experiments with LMP-103S are reported herein for pressures between 0.69 and 34.5 MPa (100 – 5,000 psia). Linear burning rates were measured from transient pressure data and high-speed video. Visible and near-IR emission spectra were recorded as well. Phenomenological combustion behaviors are reported for the entire pressure range as deduced from these diagnostics, and two clear pressure regimes are observed. Hydrodynamic instabilities and supercritical combustion processes dominate the combustion behavior in the low- and high-pressure regimes, respectively. Supporting experiments include thermal cook-off of LMP-103S and impact sensitivity testing of ADN crystals extracted from LMP-103S solution. The data presented herein significantly expand upon those available in the literature for LMP-103S, and the phenomenological combustion behavior deduced from transient, high-speed video data improve upon the state-of-the-art knowledge regarding ADN-based liquid monopropellants.

Experimental study of linear burning rates of liquid isopropyl nitrate

A high-pressure burner with continuous feed is used to investigate linear burning rates of liquid isopropyl nitrate as a function of chamber pressure. The present study investigates the pressure dependence in a subset of the low-pressure region, ranging from 5.5 MPa to 8 MPa, which has been previously neglected. The new measured data fall between the extrapolated fits from prior measures at lower and higher pressures. A fit to Saint Robert’s Law is found for 5.5 MPa to 7.05 MPa with a = 0.351 and n=1.03 with a slope break indicating a combustion regime shift identified around 7.1 MPa. To the best of our knowledge, these data represent the first measured burn rates of a liquid nitrate ester in a continuous feed strand burner. A numerical simulation with a detailed kinetic model reveals the elementary reactions that are most responsible for distinct heat release zones. The potential influence of multi-phase interactions is noted.

Experimental study of linear burning rates of liquid nitromethane using a novel high-pressure continuous feed liquid strand burner

A novel high-pressure strand burner with the ability to maintain a continuous feed of a liquid monopropellant is used to investigate the linear burning rates of liquid nitromethane as a function of chamber pressure. A new procedure for obtaining linear burning rates is presented. The present study investigates that pressure dependence in a subset of the low-pressure region, ranging from 3 to 8 MPa, which corresponds to a region with significant discrepancies in existing literature. The linear burning rate of liquid nitromethane is known to have a pressure dependence that follows Saint Robert’s Law in select pressure regions. The results fall between the upper and lower limits of existing literature with a pressure exponent of n = 1.106. To complement the experimental results, numerical simulations of a simple 1D burner-stabilized flame were conducted with three different nitromethane mechanisms. Similar to the experimental results, the simulations showed significant discrepancies for the predicted flame temperatures and heat release rate for nitromethane combustion in an inert environment. The combined experimental and modeling results highlight the need to further investigate the sensitivity of the reaction kinetics to different external factors.

Active Control of Fuel Position in Opposed-Flow Strand Burner Experiments

ABSTRACT This paper describes a new active method of controlling the position of the burning surface of a solid fuel strand in an opposed-flow burner experiment. This method is designed to avoid biases to the combustion process introduced by retaining wires and to avoid experimental delays caused by wire failures. Regression rates are presented for hydroxyl-terminated polybutadiene fuel strands burning in various oxidizer streams. The active control method uses a control loop composed of a diode laser, a photodiode, and a stepper motor. For comparison, tests were also performed using the commonly used passive control method of spring-loading the fuel against a retaining wire. This active control method can only be used at high values of oxidizer mass flux; at low mass flux, soot accumulation obstructs the laser beam. In passive control tests, the retaining wire conducts heat from the flame into the fuel and increases regression rate. The presence of the wire also modifies the extinction limits of the flame. A model was developed for regression rate as a function of oxidizer mass flux, oxidizer composition, and retaining wire diameter. This model can be used to correct future regression rate data for the bias introduced by retaining wires.

Effect of Nano‐Copper Chromite on the Thermal Decomposition and Combustion of AP‐Based Solid Propellants

AbstractThis investigation was devoted to exploring the application performance of nano‐combustion catalyst CuCr2O4 in solid propellants. In this paper, raw CuCr2O4 and nano‐CuCr2O4 prepared by the mechanical grinding method were applied in solid propellants containing ammonium perchlorate (AP), which are ammonium perchlorate/ hydroxy‐terminated polybutadiene propellant (AP/HTPB) and ammonium perchlorate/composite modified double‐base propellant (AP/CMDB). And the scanning electron microscope, displacement volume method, thermogravimetric analysis, differential scanning calorimetry, and strand burner method were employed to characterize the performance of AP‐based solid propellants. The results show that the AP/HTPB propellant has fewer defects on the surface of it due to the presence of nano‐CuCr2O4, but it has no significant effect on the AP/CMDB propellant. Nano‐CuCr2O4 can also significantly increase the density of AP/HTPB propellant and AP/CMDB propellant to 1.736 g/cm3 and 1.633 g/cm3, respectively. Nano‐CuCr2O4 obviously advances the decomposition temperature of AP/HTPB propellant and AP/CMDB propellant, which also increases the apparent heat of decomposition. In addition, the burning rate of nano‐CuCr2O4‐AP/HTPB propellant and nano‐CuCr2O4‐AP/CMDB propellant increased to 1.60 mm/s and 3.23 mm/s, with an increase of 28.3 % and 26.3 %, respectively. Therefore, nano‐CuCr2O4 is expected to be widely used in solid propellants due to its excellent properties.

Comparison of flame inception behavior of liquid nitromethane in inert and air environments

Nitromethane is a potential green monopropellant, but there are limited kinetic studies on its burning behavior in oxygen-free conditions, such as those required byunderwater and space propulsion applications. Mechanisms for nitromethane combustion have generally been validated against combustion experiments in oxidative environments. These studies provide valuable data on gas-phase kinetics,but they are less informative for elucidating and quantifying the complex phase interactions and initiation reactions that control the ignition and performance of liquid monopropellants at high pressures in chemically inert environments. This manuscript presents new experimental data from a novel continuous flow, liquid propellant strand burner to measure the ignition behavior of nitromethane.

Combustion of Gelled HAN/Methanol/Water Propellants

AbstractMonopropellants based on hydroxylammonium nitrate (HAN) are a low‐toxic, high‐performance alternative to hydrazine in space propulsion systems. Gelling such propellants may offer additional advantages such as the long‐term stability of compositions with added energetic and catalytic particles in suspension. However, the effect of gelling on the combustion characteristics and mechanisms of HAN‐based propellants is not well understood. In the present work, an aqueous HAN/methanol solution (70.1 wt% HAN and 14.9 wt% methanol) was gelled with 1 wt% polyacrylamide in an acoustic mixer. Combustion of the obtained gel was studied in a strand burner at nitrogen pressures up to 30 MPa. Thermogravimetric analysis, differential scanning calorimetry, and mass spectrometry were used to investigate the reaction mechanisms in the HAN/methanol/water solutions with and without polyacrylamide. The strand burner tests have shown that gelling suppressed the multiple pressure regimes observed in previous studies of the HAN‐based propellants, with a consistent pressure exponent of 1.15±0.05 over the tested pressure range 4–30 MPa. Thermoanalytical testing of the gelled mixture indicates there is no chemical interaction between the gellant and the HAN decomposition, and the effect of gelling on the combustion of HAN/methanol/water propellants is hydrodynamic in nature.

A novel multi-band plenoptic pyrometer for high-temperature applications

A snapshot, multi-band imaging pyrometer was developed for two-dimensional, high-temperature measurements. The multi-band pyrometer is composed of a plenoptic camera and a continuous wavelength filter (450 nm–850 nm), located at the aperture plane of the primary imaging lens. A microlens array forms micro-images of the color filter with a resulting spectral resolution of approximately 25 nm per pixel. A wavelength calibration method was developed to use eight narrow-band light-emitting diodes at discrete wavelengths with sub-pixel wavelength location determined using a weighted intensity centroid algorithm. Temperature calibration was conducted using a graphite block as a gray-body radiator placed in a box furnace with controlled temperatures of 600 ∘C–1100 ∘C. Temperature is estimated at each microlens location using the spectral pyrometry method, which reduces the dependency of the measurement on a priori knowledge of material emissivity. Simulated data with Gaussian noise demonstrates reduced uncertainty at higher temperatures and larger bit-depths. Experimental measurements with a heated graphite block demonstrated accuracy within the uncertainty bounds of a reference thermocouple (±0.75%). In addition, validation experiments were conducted in the more dynamic environments of a strand burner and a solidifying copper pool, which contain high-temperature, moving particles and phase dependent emissivity, respectively. The addition of alumina particles to solid fuel produced an expected increase of over 400 ∘C that was measured by both a multi-band imaging pyrometer and a thermocouple. The pyrometer demonstrated the ability to measure temperatures for liquid and solid phases of copper simultaneously with temperature accuracy better than 5%, despite larger differences in material emissivity.

Slow-Burn Ammonium Perchlorate Propellants with Oxamide: Burn Rate Model, Testing, and Applications

Low-thrust long-burn-time solid rocket motors may be useful as propulsion for small, fast, uncrewed aerial vehicles. These motors require a slow-burning propellant that can operate at unusually low chamber pressures (0.3–2 MPa). Slow-burn propellants were developed using ammonium perchlorate oxidizer and the burn rate suppressant oxamide. By varying the amount of oxamide (from 0 to 20%), burn rates from 4 to (at 1 MPa) were achieved. The adjustable burn rate allows a set of similar propellants to serve many aircraft and mission concepts. This work presents burn rate measurements (from both a strand burner and a research motor), minimum burn pressure measurements, and combustion chemical equilibrium simulations. A novel model of oxamide’s effect on burn rate is also presented, and it fits well to the experimental data. Finally, these propellant data and models are applied to select the propellant and chamber pressure for an example low-thrust solid rocket motor.

A continuous flow liquid propellant strand burner for high pressure monopropellant and bipropellant combustion studies

A combustion assembly capable of continuously burning monopropellant and bipropellant liquid fuels at pressures up to 80 bars (1145 psig) was designed and constructed. The assembly is based on a liquid propellant strand burner where a manifold maintains small positive differential pressures on the fuel to maintain a steady supply into the reaction vessel. Optical ports enable direct visualization of the flame and will allow for future spectroscopic and imaging studies of the flame. The strand burner design was tested using nitromethane with both air and inert environments in the reaction vessel. Continuous combustion was sustained for almost 8min in air (34 bars/500 psig) and more than 6min in N2 (70 bars/1000 psig). A unique outcome from the initial testing of this device is the ability to ignite liquid nitromethane in an inert environment without the use of a pilot flame started in air.

Effect of metal additives on neutralization and characteristics of AP/HTPB solid propellants

Conventional ammonium perchlorate (AP) based solid propellant system has been used in rocket boosters and most tactile missiles. But the toxic chloride emissions from the combustion of AP during the rocket launch increase the air pollution by 1%. For the conversion of conventional AP based propellant into Green propellant there approaches can be used. These are scavenging propellants, Neutralizing propellants and Non-chlorine propellants. In present study an attempt is made to investigate the characteristics of neutralizing propellant with the help of different metal additives such as Al, Mg, and combined Al–Mg. The investigations were theoretically and experimentally carried out and the results compared with base composite solid formulation (AP/HTPB). Theoretical investigation on that addition of metal additives clearly indicated helps in the reduction of toxic HCl without the compensation of performance. The burn rate study was carried out using Crawford strand burner setup in inert atmosphere under different pressure conditions ranges from 0.689 MPa to 4.136 MPa. A considerable enhancement in the burn rate of AP based solid propellant by 14.02% using pure metals as additives and by 54% using binary metal mixtures as additives was observed. Studies on HCl reduction of AP based propellant samples using gas bubbling setup shows that addition of Mg-15% to the base propellant enhances the HCl reduction by 47% but the propellant was porous in nature and addition of Al–Mg(15%) to the base propellant enhances the reduction by 37%. Thus, the binary metal mixtures in AP/HTPB based solid propellant resulted in better reduction of toxic HCl without reduction in performance than the sample without metal additives.

Combustion of aqueous HAN/methanol propellants at high pressures

“Green” monopropellants based on hydroxylammonium nitrate (HAN) are of interest for replacing highly toxic hydrazine in space propulsion systems and for other applications. However, their combustion mechanisms at high pressures are not well understood. In the present paper, decomposition of HAN at pressures up to 15 MPa was studied with differential scanning calorimetry, while combustion of an aqueous HAN/methanol solution (70.1 wt% HAN and 14.9 wt% methanol) was studied in a strand burner at pressures up to 30 MPa. The experiments in both setups were conducted in a nitrogen environment. It has been shown that the decomposition temperature of HAN significantly decreases with increasing pressure from atmospheric to 2 MPa and remains virtually constant with further increasing pressure to 15 MPa. The decrease with pressure is explained by suppressing evaporation of HNO3 and H2O formed during the decomposition process. The strand burner experiments have revealed an increase in the linear burning rate by over 50% at 12−14 MPa, explained by reaching the critical pressure of the liquid. Other observed pressure dependencies of the burning rate include a decrease at 9–12 MPa, a plateau at 14–19 MPa, a decrease at 20 MPa, and an increase at 22–30 MPa.

Generating Plateau Burning Segments for Composite Solid Rocket Propellants

AbstractPlateau burning is a highly desired phenomenon for composite solid rocket propellants. Three composite propellant compositions were prepared in an attempt to generate plateau burning segments. In these compositions the curative type was changed along with controlling the coarse to fine ratio of the oxidizer particles. The prepared compositions were extensively tested at one‐MPa segments from 1 to 14 MPa through strand burner technique. The results showed that depriving the combustion process from the thermal energy causes the plateau to exist. The thermal depriving was achieved through controlling the spacing of coarse and fine particle of the oxidizer, to reduce the intensity of their combustion flame, and through employing binders that possess high activation energy for the first decomposition stage. These findings were proven through the appearance of bi‐plateau segments for the first time for one of the employed isocyanates.

Performance of advanced composite solid rocket propellants based on novel oxidizers

Three novel high energy dense oxidizers (HEDO), Bis(2,2,2-trinitroethyl)oxalate (BTNEOx), 2,2,2-Trinitroethyl-nitrocarbamate (TNENC), 2,2,2-Trinitroethyl-formate (TNEF) have been prepared and studied as oxidizers in composite solid rocket propellants (CSRPs). For comparison, traditional CSRP containing ammonium perchlorate (AP) bonded by hydroxyl-terminated polybutadiene (HTPB) binder system was studied. The optimum oxidizer percentage with respect to the specific impulse was determined using the thermodynamic code (EXPLO5_V6.03). In addition, the optimum oxidizer mixture based on the novel oxidizers with AP was studied. The combustion properties and gaseous products of the optimum propellant compositions were calculated. A selected composition was prepared in the lab. scale and the burning rate was measured by the strand burner method. It was concluded that TNEF based propellant possess the maximum specific impulse of all the studied compositions. In addition, TNEF based propellant has a higher burning rate than the traditional CSRP composition.

Catalytic activity of nano-sized CuO on AP-CMDB propellant

ABSTRACTIn this work, nano-sized CuO prepared by mechanical grinding method was applied to the AP-CMDB propellant. The catalytic activity of nano-sized CuO on AP-CMDB propellant was investigated by thermogravimetric analysis/differential scanning calorimetric techniques. The combustion rates at atmospheric pressure of AP-CMDB propellant modified by nano-sized CuO were measured by strand burner method. The experimental results show enhancements on the thermal decomposition and combustion of AP-CMDB propellant in the presence of nano-sized CuO. Compared with blank propellant, the thermal decomposition temperature of AP-CMDB propellant shifts downward. Simultaneously, the combustion rate of AP-CMDB propellant at atmospheric pressure increases by about 22.2%. Encouragingly, the catalytic activity of nano-sized CuO on AP-CMDB propellant is superior to that of the bulk CuO, which leads to a potential application of nano-sized CuO in the AP-CMDB propellant.

Experimental investigation on combustion behaviors and reaction mechanisms for 5-aminotetrazole solid propellant with nanosized metal oxide additives under elevated pressure conditions

5-AT/Sr(NO3)2 based composite propellants have been applied in fire extinguishing fields as novel power source for several years due to its high thermal stability and large gas production, but the disadvantage of poor combustion stability still indeed to improve. In this study, the influential mechanism of two common burning modifier, nanosized Copper oxide (nano-CuO) and nanosized Zinc oxide (nano-ZnO), on the thermal behavior and combustion characteristics under high pressure of 5-AT based composite propellant was deeply investigated. Simultaneous thermogravimetric-differential thermal analysis (TG-DTA) have been used to identify the changes in the thermal and kinetic behavior of the catalyzed composite propellant. Thermogravimetric analyzer coupled with Fourier transform infrared spectrometry (TG-FTIR) analysis and pyrolysis-gas chromatography with mass spectrometry (Py-GC/MS) analysis were employed to characterize the composition of evolved gas in propellant decomposition process. Combined with results of the combustion characteristics measurement conducted on a strand burner (1–5 MPa), different kinds of influential mechanisms of these two catalysts on decomposition process and combustion process under elevated pressure conditions were proposed based on presumed physical combustion model. Outcomes of this study will contribute to the design of 5-aminotetrazole solid propellant with a superior combustion property under high pressure, and promote its application into fire-fighting area.