Ammonium Perchlorate/hydroxyl-terminated Polybutadiene Research Articles

Assessing damage degree of solid rocket motor under various cook-off conditions

This research is focused to investigate the cook-off response of a solid rocket motor (SRM) utilizing ammonium perchlorate/hydroxyl‑terminated polybutadiene (AP/HTPB) propellants under various heating environments, also assessing the damage degree of the motor casing and exploring the formation mechanism of ignition site. We have developed a three-dimensional (3-D) numerical model that encapsulates a range of response behaviors, including cook-off and mechanical destruction of the SRM. This model integrates multi-step thermal decomposition and combustion reaction models for the propellant, as well as a damage response model for the SRM. The response mechanisms of the SRM under varying thermal conditions have been simulated using a validated model. The study reveals that under both fast cook-off condition (FCC) and slow cook-off condition (SCC), the motors experienced deformation and rupture. Under FCC, ignition initiated at the motor head, reaching a maximum pressure of approximately 800 MPa near the ignition point. This led to the separation of the motor head from the main body, producing numerous casing fragments with the fragment cloud's radius showing a linear correlation to time (R = 0.73289t + 28.24286). The ignition delay was 224.5 s at an ignition temperature of 551.83 K. In contrast, under SCC, ignition occurred at the propellant's midsection inner surface, with a significantly lower maximum pressure of about 700 MPa. The motor expanded at this site, with the expansion radius exhibiting a nonlinear relationship to time (R = 0.19102t2 + 0.0096t + 32.97464). The casing then ruptured, causing extensive damage to both ends and generating a debris cloud with a radius that correlated linearly to time (R = 0.82147t + 26.70826). More severe longitudinal tearing, exceeding the FCC scenario, appeared in the motor body. The ignition delay time was 18.82 h. with an ignition temperature of 557.49 K.

Estimation of surface geometry on combustion characteristics of AP/HTPB propellant under rapid depressurization

The 2D sandwich model serves as a potent tool in exploring the influence of surface geometry on the combustion attributes of Ammonium perchlorate/Hydroxyl-terminated polybutadiene (AP/HTPB) propellant under rapid pressure decay. The thickness of the sandwich propellant is derived from slicing the 3D random particle packing, an approach that enables a more effective examination of the micro-flame structure. Comparative analysis of the predicted burning characteristics has been performed with experimental studies. The findings demonstrate a reasonable agreement, thereby validating the precision and soundness of the model. Based on the typical rapid depressurization environment of solid rocket motor (initial combustion pressure is 3 MPa and the maximum depressurization rate is 1000 MPa/s). A-type (a flatter surface), B-type (AP recesses from the combustion surface), and C-type (AP protrudes from the combustion surface) propellant combustion processes are numerically simulated. Upon comparison of the evolution of gas-phase flame between 0.1 and 1 ms, it is discerned that the flame strength and form created by the three sandwich models differ significantly at the beginning stage of depressurization, with the flame structures gradually becoming harmonized over time. Conclusions are drawn by comparison extinction times: the surface geometry plays a pivotal role in the combustion process, with AP protrusion favoring combustion the most.

Unsteady heat and mass transfer mechanism in AP/HTPB propellant combustion under rapid depressurization

This work is to investigate the dynamic instability and unsteady heat and mass transfer mechanism of Ammonium perchlorate/Hydroxyl-terminated polybutadiene (AP/HTPB) propellant in the chamber of solid rocket motor (SRM). The environment of exponential depressurization is imposed on the heterogeneous propellant combustion to simulate the actual working condition of SRM. The microstructure of AP/HTPB propellant is generated by a code based on the sequential algorithm. The combustion model consists of gas reaction, heat conduction in the solid phase, and the non-planar moving surface. A quenched model under the rapid pressure decay is established. It is found that the rapid depressurization and the change of surface morphology are responsible for the combustion fluctuation. Finally, quenched times of different propellant packs are calculated. It is found that different AP particle positions in the propellant pack will also cause a change in the quenched time.

Efficient Construction of MXene/V2O5 Nanocomposites for Optimizing the Combustion of Ammonium Perchlorate-Based Composite Solid Propellants.

The combustion performance of composite solid propellants (CSPs) significantly affects their application in the field of military and civil aircraft. Ammonium perchlorate/hydroxyl-terminated polybutadiene (AP/HTPB) composite propellants are one of the common CSPs, and their combustion performance is mainly affected by AP thermal decomposition. In this work, a simple strategy was put forward to effectively construct MXene-supported vanadium pentoxide nanocomposites (MXene/V2O5, MXV). MXene provided a good loading interface for V2O5 nanoparticles, which made MXV obtain a large specific surface area and simultaneously improved the catalytic performance of MXV for AP thermal decomposition. The catalytic experiment results showed that the decomposition temperature of AP mixed with MXV (MXV-4, 2.0 wt %) was 83.4 °C lower than that of pure AP. Moreover, the ignition delay of the AP/HTPB propellant was significantly reduced by 80.4% after adding MXV-4. The burning rate of the propellant was also increased by 202% under the catalytic action of MXV-4. Based on the above results, MXV-4 was expected to be an additive for optimizing the burning process of AP-based composite solid propellants.

Investigation on micro-flame structure of ammonium perchlorate composite propellant under wide pressure range

This present work proposes a method to exactly evaluate the thermal process of ammonium perchlorate/hydroxyl-terminated polybutadiene propellant. The 3-D propellant pack is generated by the Monte-Carlo method to obtain a representative sandwich model on the base of the dual-slicing technique. The ammonium perchlorate monopropellant flame height H1, the primary diffusion flame height H2, and the final diffusion flame height H3 are jointly determined by the temperature and component distributions of the gas phase thermodynamic field, which is more accurate to capture this complicated combustion field distribution of the gas phase comparing with earlier reported studies. Peclet number and Damkohler number are also introduced to quantitatively investigate the influence mechanism of chemical kinetics and diffusion mixing process of components on this micro-flame structure under wide pressure range (0.69-20.7 MPa). Further, based on the criterion of whether the premixed flow above the burning surface can absorb heat flux continuously from the diffusion flame to approach adiabatic flame temperature, the diffusion flame is divided into two regions in detail: flame front and trailing diffusion flame.

Study on the interfacial interaction between ammonium perchlorate and hydroxyl-terminated polybutadiene in solid propellants by molecular dynamics simulation

Abstract The interfacial interaction between the main oxidant filler ammonium perchlorate (AP) and hydroxyl-terminated polybutadiene (HTPB) matrix in AP/HTPB propellants were studied via an all-atom molecular dynamics simulation. The results of the simulation showed the effects of the microscopic cross-linked structure of the matrix, stretching rate during uniaxial stretching, and contact area between the filler and matrix on the mechanical properties, such as the stress and strain of the composite solid propellant. Among the aforementioned factors, the stretching rate considerably affects the mechanical properties of the solid propellant, and the maximum stress of the solid propellant proportionally increases with the stretching rate. When defects were introduced on the surface of the AP filler, the contact area between the filler and matrix affected the strain type of the matrix molecules. Owing to the interaction between the molecules and atoms, the strain behaviour of the matrix molecule changed with the change in its microscopic cross-linked structure during uniaxial stretching. Molecular dynamics simulations were used to explore the characteristics at the AP–HTPB interface in AP/HTPB propellants. The aforementioned simulation results further revealed the interfacial interaction mechanism of the AP–HTPB matrix and provided a theoretical basis for the design of high-performance propellants.

Evaluation of graphene-ferrocene nanocomposite as multifunctional combustion catalyst in AP-HTPB propellant

A novel graphene-ferrocene nanocomposite (G-792-Fe) was designed, prepared and characterized systemically using SEM, EDS, FTIR, XPS and RAMAN methods. The as-synthesized G-792-Fe was used as a multifunctional combustion catalyst in ammonium perchlorate-hydroxyl terminated polybutadiene (AP-HTPB) propellant, and the combustion, anti-migration, safety and mechanical performances of AP-HTPB propellant were revealed. The results showed that anchoring on the surface of graphene can effectively inhibit the migration and volatilization of ferrocene compounds in AP-HTPB propellant. The excellent combustion catalytic performance of G-792-Fe on AP-HTPB propellant (burning rate of AP-HTPB propellant containing G-792-Fe increases from 13.87 mm·s−1 to 17.28 mm·s−1 at 15 MPa) is attributed to its positive effect on AP decomposition, which generates more gas-phase products beneficial to the combustion surface thermal feedback and the gaseous REDOX reaction. Additionally, the G-792-Fe presents positive effects on reducing the electrostatic sensitivity and improving the mechanical properties of AP-HTPB propellant. The ignition energy (E50) and the maximum tensile strength (σm) of AP-HTPB propellant containing G-792-Fe were increased by 24.4 mJ and 0.75 MPa respectively, compared with AP-HTPB propellant containing catocene (156.8 mJ and 1.26 MPa). The positive effects of G-792-Fe on safety and mechanical performances of AP-HTPB propellant can be due the outstanding electrical and mechanical properties of graphene-based material. Results of this study have implications concerning design and application of multifunctional combustion catalyst, which can improve the combustion performance, anti-migration performance, safety performance and mechanical properties of AP-HTPB propellant.

Burning and radiance properties of Si in AP/HTPB based compositions

Silicon (Si) has high caloric power and its combustion product silica (SiO2) shows high emissivity in the far infrared (8–14 μm). The study adds different contents of Si into the Ammonium Perchlorate/Hydroxyl-terminated polybutadiene (AP/HTPB) based compositions, aiming to investigare the influence of Si on the burning and radiance properties of AP/HTPB based compositions by using a FTIR remote-sensing spectrometer, a Fourier Transform Infrared Spectrometer, a far-infrared thermal imager and a synchronous thermal analysis instrument. The experimental results show that the combustion oxidation reaction of Si is restricted by the oxygen balance of the system. When the oxygen balance exceeds 25 (Si content is greater than 15%), the insufficient oxygen supply makes the combustion rate significantly reduced from 0.52 g/s to 0.36 g/s, and Si cannot be closely related to the infrared performance and combustion of the AP/HTPB agent. When the Si content is less than 15%, the infrared radiation energy of the entire infrared band of AP/HTPB can be increased, and as the Si content increases, the radiation temperature increases rapidly from 342 °C to 1070 °C, and the far-infrared radiation intensity increases from 1.56 W/Sr to 7.81 W/Sr. The smallest mid-to-far infrared radiation ratio of 3 can be obtained when the Si content is 15%. Therefore, Si has a positive effect on improving the infrared radiation performance of AP/HTPB.

Interactions between turbulent flow and diffusion flame of composite propellant in a solid rocket motor with sidewall mass injection

The numerical procedure for a diffusion flame based on Ammonium Perchlorate/Hydroxyl-Terminated Polybutadiene (AP/HTPB) composite propellant in a solid rocket motor (SRM) is presented. This article accounts for the turbulence effects in a combustion chamber with sidewall mass injection and the propellant diffusion flames at various axial locations are studied. The mutual coupling between the vortical and entropy modes is discussed comprehensively, with the inclusion of the influence of the chamber pressure. The effects of the AP particle size and the combustion environment are also analyzed. The flow evolution in the SRM is clearly observed in the results: laminar, transitional, and turbulent. Compared with the laminar area, the flames fluctuate drastically in the turbulent flow, and the chemical kinetics processes are altered greatly. The diffusion and mixing processes of the reactants are enhanced due to the contribution of the turbulence, and the heat release of the flames is elevated. The high chamber pressure and smaller AP size could lead to more intensive fluctuations of the flame heat release. This may drive thermoacoustic instability in the SRM.

Experimental study of material behavior of AP-HTPB base composite solid propellant

In the present study, to characterize material behavior of an ammonium perchlorate - hydroxyl terminated polybutadiene (AP-HTPB) base composite solid propellant, tension tests were performed with various temperatures and strain rates. Stress relaxation tests were also carried out at various temperatures to determine the tensile modulus of the material. In addition to these tests, three types of cyclic tests and dilatation test were carried out to identify the effect of internal damage evolution to the material response. From the results of three cyclic tests, Mullins effect was observed, and from the dilatation test, volumetric damage incurred by de-bonding between the filler and matrix material was observed as well. Based on such findings, a viscoelastic constitutive modeling with the Mullins effect and damage is proposed. Numerical simulations with the newly proposed damage model reproduced the experimental findings at various test conditions fairly well. It was found out that the proposed viscoelastic constitutive model could be used to efficiently characterize the material response of the solid propellant depending on various strain rates and temperatures according to the present study.

Flame Dynamics of AP/HTPB Composite Propellant in Unsteady Rocket Motor Based on Multilayer Sandwich Model

AbstractThe flame dynamics of ammonium perchlorate/hydroxyl-terminated polybutadiene (AP/HTPB) composite propellant subject to the acoustic oscillations in solid rocket motors (SRMs) is numerically...

Theoretical study of the flame describing function of AP/HTPB propellant flame based on sandwich model

The flame dynamics of ammonium perchlorate/hydroxyl-terminated polybutadiene (AP/HTPB) composite propellant subject to the acoustic oscillation in a solid rocket motor (SRM) is investigated numerically in this article. The sandwich model is suitably modified to carry out the simulations of propellant combustion under unsteady conditions and the flame response to the acoustic oscillations with different amplitudes and frequencies is investigated. Attention is focused on the heat flux to the propellant surface and the Flame Describing Function (FDF) is utilized to capture the characteristics of the heat feedback under different forcing amplitudes and frequencies. The results indicate the flame responses are non-linear for the first longitudinal mode of acoustic oscillation and the gain of the fluctuating heat feedback reaches the point saturation at the high forcing amplitude. For high forcing frequency (the third acoustic mode), the gain of the fluctuating heat feedback shows insensitivity to the acoustic oscillations with different amplitudes and varies little. The full FDF confirms the nonlinearity of the gain response and shows the trend of fluctuating heat feedback as a function of different forcing frequencies and amplitudes.

Numerical Simulation and Unsteady Combustion Model of AP/HTPB Propellant under Depressurization by Rotation

AbstractIn order to study micro‐scale combustion characteristics of base‐bleed propellant ammonium perchlorate/hydroxyl terminated polybutadiene (AP/HTPB) under rotation condition, a two‐dimensional unsteady rotating combustion model of periodic sandwich model of AP/HTPB was established. In the gas phase, a two‐step overall reaction was used to couple the gas‐solid boundary layer, the rotational momentum was fitted and the numerical simulation comparative analysis under the initial combustion pressure of 0.1 MPa–5 MPa and the rotation velocity of 0–10 800 r min−1 was experimented. The results show that under the conditions of an initial combustion pressure of 3.5 MPa, a depressurization rate of 1000 MPa s−1 and a rotation velocity of 10 200 r min−1, the formation of a narrow diffusion‐controlled chemical reaction zone appears in the initial stage of depressurization. When the combustion pressure drops to about 0.88 MPa, the flame shows dual characteristics: diffusion‐controlled and premixed combustion. When the pressure dropped to 0.1 MPa, the flame is premixed combustion. The steady combustion process at different combustion pressures and different rotational velocities was analyzed numerically. The results show that there is a linear positive correlation between the flame deflection angle and the pressure. When the rotation velocity is lower than 10 000 r min−1, the deflection angle of the gas flame increases linearly with the rotation velocity. However, when the rotation velocity is higher than 10 000 r min−1, the deflection angle and the rotation velocity of the gas flame increase exponentially. The trend of the average Reynolds number of the combustion surface is basically the same as that of the deflection angle of the gas flame. Therefore, the average Reynolds number of the combustion surface can be used to describe the influence of rotation and pressure on the deflection angle of the gas phase flame.

Oscillatory Pressure Effect on Mean Burning Rates of Solid Propellant Combustion at Low Frequency Conditions

AbstractThe investigation on acoustic pressure oscillation effect on the mean burning rate of solid propellant combustion under quasi and unsteady conditions has been carried out at elevated pressures typical of rocket operation. Experiments were conducted in the window bomb test facility to measure mean burning rate for the mean chamber pressure ranging from 1 to 7 MPa. This analysis facilitates current understanding regarding the mean burning rate variation due to the acoustic wave interaction with propellant flame complexes. The present work is focused on low frequency combustion instability and characterization of the long rocket motor configurations. The window bomb is connected with an exclusively designed rotating valve for the generation of acoustic pressure oscillations at selected mean chamber pressures and the corresponding frequency of operations. Two kinds of ammonium perchlorate/hydroxyl terminated poly butadiene (AP/HTPB) based composite propellants, one with aluminium, and the other without aluminium were considered. Experiments were conducted at the excited frequency of 140 Hz with different excited acoustic pressure amplitudes. This experimental setup facility has the provision of visual observation during propellant combustion using optical methods. The performance of the rotating valve was competent over the tested conditions. These studies reveal that the acoustic pressure amplitudes significantly enhance the mean burning rate of the solid propellants. The response of the aluminized propellants to the excited acoustic oscillations was relatively prominent as compared to that of its non‐aluminized counterparts. The results obtained are extremely useful in combustion instability characteristics of large scale rocket motors.

Numerical simulation of cook-off characteristics for AP/HTPB

Abstract The present study is devoted to researching the thermal security problems of large-scale solid rocket motor with Ammonium perchlorate/Hydroxyl-terminated polybutadiene (AP/HTPB). A two-dimensional axisymmetric model for the cook-off of solid rocket motor is established. The reaction kinetics for the cook-off process of AP/HTPB is described by the two-step global chemical mechanism. Numerical predictions of the cook-off behavior for the propellant are conducted at fast heating rate of 1.45–2.45 K/s, and slow heating rate of 0.001–0.003 K/s, respectively. The results show that in the fast cook-off condition, the initial ignition position of AP/HTPB occurs in the annular region of the outer wall of propellant without exception, and the center point in the region is (889.1, 149.5). For the region, the axial width is 1.8 mm and radial thickness is 0.8 mm. However, in the slow cook-off condition, the ignition center position is shifted along the axial direction toward the right end face of the propellant with the increase of heating rate. Therefore, the influence of heating rate on ignition temperature and ignition delay time is nonnegligible within a certain range.

Numerical Simulation of Quenched Combustion Model for AP/HTPB Propellant under Transient Depressurization

In order to study the out-of-phase blowing effect of ammonium perchlorate/hydroxyl terminated polybutadiene (AP/HTPB) under transient depressurization, a two-dimensional periodic sandwich model of AP/HTPB sandwich unsteady combustion was established. The gas-solid two-step total reaction was used to couple the gas-solid boundary layer, and the AP/HTPB step-down temperature criterion was applied. The numerical simulation comparative analysis under the initial combustion pressure 3.5 MPa∼10 MPa and the initial depressurization rate of 1000 MPa/s∼2000 MPa/s was experimented. The results show that under the condition of initial combustion pressure of 3.5 MPa and pressure reduction rate of 1000 MPa/s, the formation of narrow diffusion chemical reaction zone appears in the initial stage of depressurization; when the combustion pressure drops to about 1.7 MPa, the flame appears dual characteristics: diffusion and premixed combustion; when the pressure dropped to 0.1 MPa, the flame is premixed combustion. AP(g) is the main factor causing the out-of-phase blowing effect. During the initial stage of combustion, AP(g) decomposes rapidly and its decomposition product is close to the combustion surface; with the rapid decrease of pressure, the gas-phase heat feedback decreases, leading to the decrease of AP(g) decomposition rate and AP(g) diffusion to the gas phase, which is the initial phase of out-of-phase blowing effect; when the gas phase heat feedback reduced and the temperature is below 750 K, the decomposition of AP(g) is stopped and the undecomposed AP(g) is further diffused to the gas phase region, which exacerbates the out-of-phase blowing effect, leading to the extinguishment of AP/HTPB. The unsteady combustion flame extinction time increases with the initial combustion pressure, and decreases with the initial depressurization rate.

Effect of wall heat transfer characteristic on the micro solid thruster based on the AP/HTPB aerospace propellant

The combustion processes of Ammonium perchlorate/Hydroxyl-terminated polybutadiene (AP/HTPB) aerospace propellant are analytically studied to determine the effect of wall materials and thickness on the heat transfer characteristics within a solid micro thruster. In this study, the flow and combustion characteristic of the propellant are described in terms of a combustion rate equation associated with a chemical reaction kinetics model. The profiles describing changes in the combustion surface by the user-defined function (UDF) are used to obtain the distribution of temperature and content of propellant using a numerical method. The combustion time and thrust are simulated to validate the results by comparing them with the results from the literature. Moreover, we demonstrate that a smaller thermal conductivity and a larger heat capacity of the wall material can lead to improvement in the thruster performance. Finally, analyzing the significant effects of wall thickness the heat transfer characteristics within the thruster allowed us to determine an optimal wall thickness value.

Experimentally-validated mesoscale modeling of the coupled mechanical–thermal response of AP–HTPB energetic material under dynamic loading

This manuscript presents a combined computational–experimental study of the mesoscale thermo-mechanical behavior of the Hydroxyl-terminated polybutadiene (HTPB) bonded ammonium perchlorate (AP) composite energetic material subjected to dynamic loading conditions. The computational model considers the AP–HTPB interface debonding, post-debonding interface friction and temperature rise due to viscoelastic dissipation as well as dissipative interfacial processes. The interface is modeled using a cohesive zone model combined with a contact algorithm to account for the interface separation, particle/binder contact and heat generation. The HTPB binder is modeled as viscoelastic with adiabatic temperature rise. Three experiments are conducted to calibrate and validate the model. Raman spectroscopy and indentation experiment are employed to determine the interface properties, whereas Kolsky bar tension test along with in-situ synchrotron X-ray diffraction measurements are used to validate the model and understand the interface separation characteristics under dynamic loading.

Cook-off test and numerical simulation of AP/HTPB composite solid propellant

A small-scale cook-off experimental apparatus is designed to study the thermal security problem of ammonium perchlorate/hydroxyl-terminated polybutadiene (AP/HTPB) propellant. Both fast and slow cook-off tests are carried out. Then a cook-off model is established according to the apparatus, and compared with the experimental results. The theoretical calculation results are in good agreement with the experimental data. On this basis, the numerical predictions of cook-off properties of a solid rocket motor are performed at different heat fluxes. Results show that the initial cook-off response locations of AP/HTPB propellants all occur in the annular region on the outer wall of the propellant near the nozzle under the three kinds of heating flux conditions. The insulating effect of the insulating layer in the motor increases as the heat flux is increasing. When the heat flux is 103 W m−2, 104 W m−2 and 105 W m−2, the ignition delay time is 5071.5s, 348.5s and 41.0s, with the corresponding ignition temperature of 533.89 K, 554.97 K and 590.07 K, and the shell temperature of 521.38 K, 597.08 K and 837.92 K during the ignition near the nozzle. It can be concluded that the ignition delay time of the propellant is significantly shortened, and the ignition temperature as well as the shell temperature during the ignition near the nozzle increases gradually when the heat flux is increasing.

Experimental Study of Transient Combustion Characteristics of AP/HTPB Base Bleed Propellant Under Rapid Pressure Drop

The transient combustion characteristics of ammonium perchlorate/hydroxyl-terminated polybutadiene (AP/HTPB) base bleed propellant, associated with the transient instability by sudden depressurization in the moment that the base bleed projectile flies away from the muzzle, was investigated experimentally. The combustion pressure is between 3.15 MPa and 7.24 MPa and the depressurization rate is controlled in the range of 13.2–136.7 MPa/s. The gas phase flame blows away from the nozzle, leading to a thicker gas phase reaction layer and smaller heat feedback to the propellant under rapid depressurization. Two main traits were presented under depressurization, which is gas phase flame extinction and non-extinction. There is a smallest critical depressurization rate to extinguish AP/HTPB base bleed propellant, which increases as the maximum combustion pressure increases and has the empirical relationship of |dp/dt|cr = 58.29 p - 140.78 MPa/s. Although the gas phase flame is extinguished by rapid depressurization, the condensed phase decomposition may still exit and a stable re-ignition flame may occur if enough heat is accommodated. The gas phase flame extinction behaviors under transient depressurization can be divided into three regions, which are extinction with then re-ignition, extinction with maintaining decomposition, and absolutely extinction.