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Article

Research on Ablation Device Suitable for Thermal Protection System of Solid Rocket Ramjet

by
Jiming Cheng
1,*,
Hang Yan
1,
Xiping Feng
1,
Guoqiang Zhu
1,
Jie Liu
2 and
Xintong Qi
2
1
National Key Laboratory of Solid Rocket Propulsion, Northwestern Polytechnical University, Xi’an 710072, China
2
Beijing Power Machinery Institute, Beijing 100074, China
*
Author to whom correspondence should be addressed.
Aerospace 2025, 12(9), 772; https://doi.org/10.3390/aerospace12090772
Submission received: 8 July 2025 / Revised: 18 August 2025 / Accepted: 26 August 2025 / Published: 27 August 2025
(This article belongs to the Section Astronautics & Space Science)
Browse Figures
Versions Notes

Abstract

In solid rocket propulsion systems, overload effects induced by aircraft maneuvers can lead to gas accumulation in the afterburning chamber, resulting in severe localized ablation of thermal insulation layers and significantly compromising overall operational stability. Traditional ablation experimental methods (e.g., oxyacetylene and plasma ablation) exhibit poor correlation with the actual thermal environments in solid rocket ramjets, thereby posing substantial challenges for simulating real operational conditions. To address this issue, an oxygen-kerosene engine-based ablation device was developed. Methodologically, the CEA-optimized oxygen-to-fuel ratio (3.5) enabled authentic combustion simulation, while 3D compressible flow modeling (Ansys Fluent 2020 R2) quantified critical parameters such as chamber pressure and achieved precise control of surface temperature. Ablation experiments were conducted on diverse ablative materials using this device, yielding a maximum error in mass ablation rate of only 5.67%. This demonstrates the high accuracy of the device, which meets the requirements for ablation experiments. This reliable simulator (with an error <6%) provides a validated platform for high-fidelity evaluation of ablation performance in maneuverable solid rocket ramjets.

1. Introduction

In aircraft, an excellent propulsion system is critical to maneuverability. Solid rocket ramjets, which use atmospheric oxygen as the main oxidant, have been widely utilized due to their simple structure and high specific impulse [1,2,3,4]. The afterburning chamber is the primary location for releasing energy stored in the propellant. However, the complex oxygen-enriched two-phase environment inside the afterburning chamber poses challenges for ensuring thermal protection, thereby slowing the development of solid rocket ramjets [5,6,7]. The ablation performance of the thermal protection system (TPS) materials in the afterburning chamber is a key factor in determining the efficacy of passive thermal protection [8,9,10,11,12] and an important component in the thermal structure design of solid rocket ramjets [13,14].
Initially, the solid rocket propulsion systems were primarily employed in vehicles cruising at specific altitudes. However, aircraft featuring beyond visual range air-to-air missiles have higher requirements for maneuverability. The propulsion systems must be sufficiently agile to accommodate operating conditions involving high angles of attack and specific sideslip angles [15,16,17]. The overload effects and non-uniformity of the internal flow field caused by such maneuvers can lead to gas accumulation inside the afterburning chamber, which deteriorates the operational environment of the TPS materials [18,19,20] and adversely affects their ablation performance.
Ablation experiments play a crucial role in investigating the performance of TPS materials. Currently, two main types of experimental methods are available. The first category utilizes high-energy beams to impinge on the surface of the materials, encompassing techniques such as oxyacetylene ablation, plasma ablation, laser ablation, arc ablation, as well as large-scale ablation wind tunnels. Hou et al. [21] prepared a new type of carbon fiber-reinforced polydisiloxane and characterized its ablative properties using oxyacetylene flames with different heat flux densities. Su et al. [22] employed plasma ablation to analyze the ablation behaviors and mechanisms of molybdenum aluminum boride (MoAlB) composite materials at 1943–2823 K, while comparing their heat release mechanisms under different temperature conditions. Geng et al. [23] investigated the interaction mechanisms between high-energy CO2 lasers and carbon/carbon (C/C) composites in an anaerobic environment, systematically analyzing the effects of material properties and laser parameters on ablation performance. Momozawa et al. [24] simulated re-entry conditions using an arc-heated plasma wind tunnel, assessed the dynamic oxidation behavior of silicon carbide (SiC) materials, and proposed a transition boundary between active and passive oxidation of SiC. The experimental methods involving high-energy laser beams can effectively simulate the high-temperature conditions required for investigating the ablation properties of TPS materials. Accordingly, such methods are commonly used in preliminary explorations of ablation mechanisms, facilitating the screening of new materials. However, such methods can only simulate limited environmental parameters (e.g., temperature), and their correlation with the actual thermal environments of ramjets remains poor.
Another method for investigating the performance of TPS materials involves ground testing of solid rocket ramjets or ablation experimental devices. Cohen et al. [25] compared the ablation performance of various TPS materials through experiments on solid rocket propulsion systems. Their ground experimental system shared the same structure as actual solid rocket propulsion systems, and the silicone TPS material DC93-104 exhibited the lowest ablation rate. Li and Liu et al. [26,27,28] designed a bend test motor to simulate overload conditions and evaluate the anti-overload ablation performance of various TPS materials. Lou et al. [29] constructed a large-scale oxygen-enriched ablation experimental system, and Wang et al. [30,31,32] designed an oxygen-kerosene ablation experimental device to investigate the anti-ablation performances of various TPS materials. However, these devices can only simulate oxygen-rich two-phase flows and have limitations regarding the simulation of internal pressure and gas velocity in solid rocket propulsion systems [33,34,35].
A relatively small number of studies have focused on the ablation performance of the insulation layer in the afterburning chamber of solid rocket ramjets. Consequently, in this study, an ablation experimental device was designed and constructed to evaluate the operational environment of solid rocket ramjets. This device was employed to evaluate the ablation behavior of TPS materials while precisely regulating critical ablation-influencing parameters, which include the internal pressure of the afterburning chamber, surface temperature of the specimens, gas flow velocity, incidence angle between gas and specimens, oxygen concentrations, particle phase concentration, and particle size. The characteristics of flow field parameters were determined through numerical simulations. The ablation tests on different TPS materials were conducted to verify the effectiveness and stability of the equipment.

2. Experimental Setup, Methods and Materials

2.1. Experimental Setup

2.1.1. Design of Ablation Experimental Device to Simulate Solid Rocket Ramjet

The experimental device comprises control, oxygen supply, kerosene supply, water cooling, and particle loading system, as well as a small-scale oxygen–kerosene liquid rocket engine (both its size and mass flow rate are smaller than those of full-scale rocket engines [36]), an outlet extension section, specimen holders, and other components, as illustrated in Figure 1. Water cooling system: Model DX-30AD (Zhengzhou Lijia Thermal Spray Machinery Co., Ltd., Zhengzhou, China), with a water pump pressure of 0.9 MPa, a maximum cooling water flow rate of 15 m3/h, a standard refrigeration capacity of 90 kW, R22 refrigerant, a finned condenser, and a shell-and-tube evaporator. Control system: Model HV-8000 (Zhengzhou Lijia Thermal Spray Machinery Co., Ltd., Zhengzhou, China), incorporating fully closed-loop PLC control. It enables closed-loop regulation of flow rates and pressures for kerosene, oxygen, and carrier gas, controls ignition startup and shutdown procedures, and facilitates real-time monitoring of flow rates and pressures. Key parameters include the following: kerosene pressure of 1.2 MPa (maximum flow rate 12 GPH [Gallons Per Hour]); oxygen pressure of 1.6 MPa (maximum flow rate 2300 SCFH [Standard Cubic Feet per Hour]); and carrier gas pressure of 1.2 MPa (flow rate 3–15 L/min). Particle loading system: Model C-1100 (Zhengzhou Lijia Thermal Spray Machinery Co., Ltd., Zhengzhou, China), featuring a powder feeding pressure range of 0.5–1.1 MPa, a carrier gas flow rate of 0–15 L/min (with nitrogen as the working medium), a maximum powder feeding rate of 350 g/min, a powder particle size range of 5–150 μm, and a feeding accuracy of 1%. Kerosene supply system: Model KM-80 (Zhengzhou Lijia Thermal Spray Machinery Co., Ltd., Zhengzhou, China), with a storage capacity of 200 L and a stable pressure of 1.3 MPa. Oxygen supply system: A standard 100 L liquid oxygen Dewar flask operating at 2 MPa. Liquid oxygen is vaporized into gaseous oxygen via an evaporator prior to entering the control system. The orange line connecting the control system and the liquid rocket engine represents the ignition power supply line, with ignition achieved via a spark plug.
The water cooling system is used to cool the oxygen–kerosene liquid rocket engine, enabling the liquid rocket engine to maintain long-duration operation (no less than 1000 s). The combustion chamber, nozzle, and extension section of the rocket engine all feature a jacketless water-cooled design. Cooling water flows through the combustion chamber-nozzle and its extension section, then returns to the water chiller. Within the chiller, the refrigerant absorbs heat through evaporation in the evaporator, thereby extracting heat from the water. It then releases heat through condensation in the condenser, discharging the heat to the external environment. This forms a continuous cycle: chilled water cooling → equipment cooling → heat discharge. During the test, the water flow rate was approximately 3.6 m3/h. Furthermore, since particles are added to the gas flow at the expansion section of the liquid rocket nozzle, a constant-diameter extension segment is employed at the nozzle outlet to promote more thorough mixing of particles with the gas flow, enhancing both particle velocity and uniformity in the gas flow.

2.1.2. Specimen Fixture and Back Wall Temperature Measurement System Design

The material specimen fixture primarily consists of a specimen clamp, a temperature sensor (Xi’an Xinmin Electronic Technology Co., Ltd., Xi’an, China) with a fixing plate, and a fixing support, as shown in Figure 2. The main assembly and connections between the components are shown in Figure 2 and Figure 3. When installing the specimen, the connecting bolts and pressure plates are removed; the specimen is placed in the groove; and the pressure plates and bolts are reinstalled. A DEWESOFT Sirius data acquisition system and K-type thermocouples (Xi’an Xinmin Electronic Technology Co., Ltd., Xi’an, China) were used to measure the back wall temperature of the specimens.
The specimen clamp consisted of silicate plates and 99% ceramic plates. Silicate plates are a type of material composed of magnesium silicate and additives. They are highly environmentally friendly as they do not release toxic gases or emit harmful radiation. Ceramic plates also exhibit excellent fire-resistant properties, demonstrated by the fact that large or rapid temperature changes do not significantly affect their characteristics. In the experiments, the specimens were placed on top and pressed using a pressure plate of 310S stainless steel [37]. Selecting the appropriate materials can significantly mitigate the influence of the front-end flames on the temperature measurement of the back wall of the specimen, as well as on the stability of the steel support and sensor fixing plate (310S stainless steel). This ensures the structural integrity of the specimen fixture during the ablation tests.
The support consisted of a steel bracket with three cutouts in the middle and a fixing foot, which were connected using M10 hexagonal socket-head cap screws. Different ablation angles can be achieved; in this study, the angle of the connection surface was set to 45° to ensure that the ablation angle was also 45°, and threaded holes positioned at the bottom of the fixing foot could be connected to an external workbench.
The temperature sensor and its fixing plate are located between the specimen clamp and fixed bracket, which are connected and fixed using M6 bolts. Owing to the ceramic partition plate, the specimen does not come into direct contact with the bracket and pressure plate. Moreover, the temperature sensor and its mounting plate were mostly located in the open region of the steel bracket, which reduces the contact between the fixing plate and the steel bracket. This further alleviates the influence of the heat conducted through the support on the temperature measurement results, thus ensuring the accuracy of the temperature measurement of the back wall of the specimen.
A K-type thermocouple was employed to measure the back wall temperature of the specimen. The nominal chemical composition of the positive pole (KP) was 90% Ni and 10% Cr, while that of the negative pole (KN) was 97% Ni and 3% Si, with a usable temperature range of 73 to 1573 K. The K-type thermocouples offer excellent linearity, high thermoelectric potential, high sensitivity, good stability, uniformity, strong oxidation resistance, and cost-effectiveness. They can be used in inert or oxidative atmospheres and are widely applied in medium- to low-temperature measurements.
To prevent unintended damage to the thermocouple and ensure reliable temperature measurements, the specimen tooling was designed such that at the center of the specimen, three temperature measurement points were evenly distributed along a 12 mm diameter circle. Customized K-type thermocouples were installed via an M6 mounting thread, and the thermocouple leads were constructed from braided high-temperature glass-fiber wires.
The Sirius acquisition system is a high-performance and compact industrial computer capable of rapidly and stably recording and processing real-time data. It employs dual 24-bit A/D technology, with a maximum sampling frequency of 250 kHz and a write speed exceeding 180 MB/s, thereby fully meeting the requirements for high-speed synchronous acquisition of video signals. This acquisition system is connected to the host computer via a USB data cable.

2.2. Experimental Methods

In solid rocket ramjets, the gas flow in the afterburning chamber is generally a high-temperature, high-speed two-phase flow, and the gas is typically oxygen-rich. Additionally, as ramjet engines usually operate for extended durations, the thermal insulation layer in the afterburning chamber is subjected to a severe ablation environment. Directly utilizing test engines to study the ablation of insulation layers is costly. The developed ablation device can simulate the thermal environment of the ramjet engines, enabling the investigation of ablation characteristics of thermal insulation materials at lower cost. The experimental simulation method for the thermal environment is introduced below.

2.2.1. Oxygen Concentration Simulation Methods

Since the experimental device utilizes an oxygen–kerosene engine to generate high-temperature gas flow, the oxygen-to-fuel ratio of the engine directly affects the gas species in the combustion chamber and the oxygen concentration of the gas flow at the specimen surface. This issue was addressed using the Chemical Equilibrium Application (CEA) software package, (CEA2) one of the most mature and widely used computational tools. The calculations in CEA are based on the principle of minimizing Gibbs free energy [38,39]. Based on CEA results of the combustion chamber under different oxygen-to-fuel ratio conditions, the variation of oxygen concentration in the combustion products with respect to the oxygen-to-fuel ratio can be determined, thereby enabling the determination of oxygen concentration in the ablation gas flow.

2.2.2. Temperature Simulation Methods

Upon determining the oxygen content of the gas flow, the oxygen-to-kerosene ratio of the liquid engine can be established. Subsequently, the gas flow temperature at the specimen surface can be derived from flow field simulation results. Since the gas flow temperature decreases with increasing distance between the specimen surface and the nozzle extension exit, the plume flow field distribution of the engine is obtained through iterative simulations. This provides the gas temperature distribution at varying distances from the nozzle exit, thereby enabling the determination of the distance from the center of the specimen surface to the extension exit.
Figure 4 presents the grid used in the present study. Simulations were conducted using the ANSYS Fluent (2020R2) software package [40]. To ensure accurate and efficient calculations, a three-dimensional model was employed, with the finite volume method and a Poly-Hexcore mesh utilized to calculate the flow field in the oxygen–kerosene engine nozzle and its outlet environment. No-slip and thermally insulating boundary conditions were applied at the walls. The gas flow was modeled as a three-dimensional steady flow, and the renormalization group (RNG) k–ε two-equation model was selected to characterize turbulence.

2.2.3. Two-Phase Flow Simulation Methods

A key characteristic of the ablation environment in the solid rocket ramjet afterburner is the high-speed two-phase flow, where particle erosion significantly affects material ablation behavior. Accordingly, a particle loading system was incorporated into the experimental device designed to simulate the two-phase flow environment. The particle loading system simulates a two-phase flow environment by injecting powder particles into the nozzle section of the oxygen–kerosene rocket, thereby enabling the investigation of particle erosion effects on TPS materials. Given the relatively low pressure in the straight-pipe section downstream of the oxygen–kerosene rocket nozzle outlet, a rotary-disc powder feeder was employed, as illustrated in Figure 5. Based on gas dynamics principles, this feeder operates as follows: powder falls by gravity from the hopper into the grooves of the rotating disc. Driven by an electric motor (Shenzhen Inovance Technology Co., Ltd., Shenzhen, China), the disc rotates until the powder-filled grooves are aligned with the suction tube. Carrier gas (nitrogen) is then injected into the sealed chamber through the inlet pipe. Under gas pressure, the powder is entrained and conveyed through the outlet tube to the outlet of the oxygen–kerosene rocket engine. The powder feed rate, and consequently the particle concentration in the two-phase flow, are regulated by adjusting the rotational speed of the disc.

2.2.4. Pressure Environment Simulation Methods

The previous experiments were conducted under atmospheric pressure. In ramjet environments, TPS materials need to withstand pressure conditions; accordingly, a pressure chamber was designed in this work. To simulate the pressure environment, a pressure chamber was installed downstream of the oxygen–kerosene rocket engine nozzle outlet. The pressure within the test chamber can be regulated by adjusting the throat diameter at the pressure chamber outlet, thereby enabling pressure environment simulation, as illustrated in Figure 6a,b. Two pressure taps are mounted on the sidewall for pressure measurement. A temperature measurement point is positioned in the rear section of the pressure chamber to monitor the gas flow temperature via a Type B platinum–rhodium thermocouple (Pt-30Rh/Pt-6Rh) with a measurement range of 273–2073 K. The gas flow temperature has decreased at this location, and this temperature sensor is used for non-critical parameter acquisition, serving solely for cross-verification with the flow field simulation results in the pressure chamber. The installation and removal of the thermal insulation layer specimens are accomplished by disassembling/reassembling the upper cover of the pressure chamber, as shown in Figure 6c. Figure 6c presents the top view of the pressure chamber, with its internal view corresponding to the specimen installation fixture in Figure 6d. The specimen fixture is equipped with thermocouples to measure the back wall temperature of the thermal insulation layer, as depicted in Figure 6d. The pressure chamber features a jacketed water-cooling structure, which is circulated and cooled via a water-cooling system (Figure 6e,f). It can simulate a maximum pressure of 0.7 MPa, adequately meeting the pressure parameters of typical flight conditions for subsonic solid rocket ramjet afterburning chambers.

2.3. Materials

We tested various TPS materials, including silicone rubber, PR, EPDM, and PIC (Purchased from the local chemical market), which are typical TPS materials employed in different components of ramjet engines such as combustion chamber of gas generators, afterburning chambers, and nozzles. To enhance the performance of TPS materials, various fillers (such as SiO2, fibers, and resins) were incorporated into the matrix materials. These fillers improve the ablation resistance of thermal protection materials by forming protective liquid films, strengthening the char layer skeleton, and increasing the char yield.
Moreover, following the reference standard GJB 323B-2018 [2], TPS materials with dimensions of 50 mm × 50 mm were selected. This dimension was chosen to cover the uniform heat flux zone in the core area of the ablation device. Considering both engineering requirements and experimental safety factors, a thickness of 17 mm was adopted.

3. Results and Discussion

3.1. Determination of Oxygen-to-Fuel Ratio

The combustion products of kerosene–oxygen combustion primarily comprise nine species, with the top five ranked by mole fraction being O2, CO2, H2O, CO, and OH (Table 1). This indicates that under this oxygen-to-fuel ratio, the kerosene–oxygen rocket engine operates under oxygen-rich conditions, with a substantial amount of oxygen remaining unreacted in kerosene combustion. The temperature, gas constant, oxygen content, and specific heat ratio of the gas in the combustion chamber were obtained via CEA calculations for nine different oxygen-to-fuel ratios: 2.5, 3, 3.5, 4, 4.5, 5, 5.5, 6, and 6.5.
First, the theoretical combustion temperature was analyzed. The theoretical combustion temperature is constrained by chemical reactions governed by the first law of thermodynamics. At an oxygen-to-fuel ratio of approximately 3, the combustion temperature reached a maximum of 3728.15 K. When the oxygen-to-fuel ratio was less than 3, the temperature of the gas in the combustion chamber increased gradually as the oxygen-to-fuel ratio increased. When the oxygen-to-fuel ratio was greater than 3, the gas temperature in the combustion chamber decreased gradually as the oxygen-to-fuel ratio increased. These results indicate that kerosene achieves complete combustion at an oxygen-to-fuel ratio of 3.
Figure 7 displays the variation in the oxygen content of the gas in the combustion chamber as a function of oxygen-to-fuel ratio. The oxygen content, expressed as a mole percentage (mol%), increases with increasing oxygen-to-fuel ratio. As the oxygen-to-fuel ratio increases, both the gas constant and the specific heat ratio gradually decrease, as shown in Figure 7b,c, while the temperature of the gas first increases and then decreases (Figure 7a). When the oxygen-to-fuel ratio exceeds the theoretical ratio and continues to increase, the proportion of unreacted oxidizer rises. The cooling effect of excess oxygen becomes dominant, causing the temperature to decrease with increasing oxygen-to-fuel ratio [41]. The overall thermodynamic calculations were based on the ideal gas assumption; therefore, the molar concentration was equivalent to the volume concentration. The data plotted in Figure 7d indicate that when the oxygen-to-fuel ratio increases from 1.0 to 6.5, the oxygen volume concentration increases from 0% to 39.7%. In addition, as the oxygen-to-fuel ratio was adjusted upward or downward, the oxygen concentration in the ablative environment changed accordingly, enabling a wider range of oxygen concentrations to meet the measurement requirements.
To ensure the experimental environment more closely reflect reality, an oxygen volume concentration of 10–15% was targeted, leading to the selection of an oxygen-to-fuel ratio of 3.5. The combustion chamber was custom designed with a maximum design flow rate of approximately 40 g/s. In practice, the throat diameter is determined based on varying flow rates, combustion chamber pressure, and temperature, and thus differs across tests. During testing, the gas flow rate is first estimated based on the surface heat flux of the specimen. Subsequently, the oxygen and kerosene flow rates are determined according to the oxygen-to-fuel ratio. The oxygen flow is controlled using an HFC-303 flow controller (TELEDYNE, Thousand Oaks, CA, USA), while kerosene flow control utilizes an NV-AT75Q pneumatic valve actuator (NANVA, Wuxi, China), a Burkert 3280 proportional control valve (Christian Bürkert GmbH & Co. KG, Ingelfingen, Germany), and an 8045 electromagnetic flow meter. The nozzle throat diameter is finally determined. Under the test conditions, the combustion chamber pressure is generally maintained at 0.9 MPa to ensure stable combustion. Based on the structure and characteristics of the oxygen–kerosene rocket engine, the kerosene flow was determined to be 0.0042 kg/s, with the oxygen flow rate set as 0.01479 kg/s, yielding a corresponding oxygen volume concentration of 12.82%. The combustion chamber pressure is 0.9 MPa, and the nozzle throat diameter is 6.5 mm.

3.2. Temperature Calibration

Due to the inclined surface of the external specimen (Figure 8a–d), the flame deflected upward after impacting the specimen, resulting in asymmetric flow. Consequently, the temperature in the upper region of the specimen was higher than that in the lower region. The flow field temperature upstream of the specimen was approximately 2200 K, with the maximum surface temperature of the specimen reaching around 2050 K. Furthermore, the temperature measurement technique (Figure 8e) used for flame temperature calibration indicated that the flame region temperature was approximately 2200 K (Figure 8f), while the surface radiation temperature of the specimen was approximately 1900 K [42]. The multispectral camera (model IMEC-25) features a maximum frame rate of 340 fps, a pixel array of 2045 × 1088 pixels, 25 spectral bands per spatial pixel, and a response range of 600–1000 nm. The zoom lens (Edmund Optics VIS NIR, Barrington, NJ, USA) is coated with a broadband anti-reflection coating optimized for the 425–1000 nm wavelength range. The spectral camera was mounted on a tripod and optical lifting platform and adjusted to align with the nozzle flame. Focus was calibrated prior to ablation testing to ensure clear imaging of the specimens. Spectral images were collected through a USB 3.0 connection to a computer, with acquisition triggered by software. During image acquisition, exposure time was adjusted in real time to obtain high signal-to-noise ratio spectral images of both the flame and specimen surface.

3.3. Repeatability and Discriminability of the Experimental System

To verify the reliability of the oxygen-enriched ablation experimental device, ablation experiments were conducted over 600 s on commonly used thermal insulation materials. In this work, six TPS materials were selected [43,44], of which the first three (silicone rubber 1, silicone rubber 2, PR) were used to verify the repeatability and discriminability of the experimental system, while the last three (EPDM 1, EPDM 2, PIC 1) were chosen to characterize the severity of the ablation environment. During the ablation process, the back wall temperature of the specimens was monitored, and the mass ablation rate was measured after the experiments were completed. To balance experimental efficiency and cost, two samples per group were tested for repeatability.
Figure 9 presents the morphology of three types of thermal insulation materials after the ablation. The silicone rubber #1 material exhibited surface cracks, which is detrimental to enhancing its ablation resistance. Table 2 summarizes the mass ablation rates and corresponding relative errors for the three types of thermal insulation materials. Employing multiple samples for repeated experiments helps mitigate experimental errors. The ablation resistance performance can be ranked in the following order: silicone rubber 2 > silicone rubber 1 > phenolic resin.
During the ablation experiments conducted on three types of thermal insulation materials, the back wall temperatures of the specimens were monitored. The back wall temperature data are presented in Figure 10, while Table 3 displays the back wall temperature values at 100 s intervals from 0 to 600 s. In-depth analysis of these data revealed that during ablation, the back wall temperature data of phenolic resin exhibited significantly greater dispersion compared to the other two materials. Throughout the entire ablation process, the maximum error in back wall temperature measurements for the other two TPS material types was only 4.4%. These results indicate that silicone rubber-based TPS materials demonstrate relatively high temperature stability during ablation. In the ablation experiments of PR material, the error in back wall temperature was slightly higher. Analysis suggests this may be associated with the distribution of thermocouples on the material’s back surface and the in-plane thermal conductivity properties of the material.
In addition, we further conducted ablation experiments on TPS materials such as EPDM TPS. After the conclusion of the ablation tests, phenomena including char layer fragmentation were observed, indicating that the ablation environment generated by this experimental device is relatively severe for EPDM TPS. During the ablation experiments on these TPS materials, Sample #1 was completely decomposed at 315 s (Figure 11a,d), whereas Sample #2 exhibited an abnormal increase in back wall temperature when the experiment proceeded to 281 s (Figure 11e). In the ablation experiments on polyphosphonitrile specimens, an abnormal increase in back wall temperature was observed at 365 s (Figure 11f). Based on these observations, the experiments were terminated. As shown in Figure 11b,c, the material exhibited obvious cracks after ablation. Such cracks can damage the structural integrity of TPS materials and weaken their heat transfer blocking capability, thereby adversely affecting their ablation resistance performance.
In summary, the results of the ablation experiments indicate that the errors in mass ablation rate measurements from the experimental device were small, with a maximum of only 5.67% (Table 2). This reflects the high accuracy of the experimental system used in this study, demonstrating that the design of the experimental system is rational and reliable. It meets the requirements for ablation experiments and lays a stable foundation for subsequent research. Additionally, in the ablation experiments on EPDM and polyphosphazene TPS materials, abnormal back wall temperature increases, and surface cracks occurred after a certain duration. This indicates that when such TPS materials are used in afterburning chambers, continuous monitoring of the oxygen-rich environment is necessary to detect potential issues.

3.4. Pressure Environment and Two-Phase Flow

Based on the ablation experimental device, the pressure environment and two-phase flow environment are simulated using the pressure chamber and particle loading system, enabling the investigation of the effects of pressure conditions and condensed-phase particles on the ablation behavior of TPS materials. The rationality and feasibility of the ablation experimental device in pressure simulation and two-phase flow environment simulation have been verified through verification tests.

3.4.1. Pressure Environment

To verify the pressure simulation capability of the experimental device and its stability, the ablation tests on TPS materials were conducted under different pressurized conditions (0.4 MPa, 0.5 MPa, and 0.6 MPa), with the measured pressure curves presented in Figure 12. The pressure from the two front and rear measurement points within the pressure chamber was consistent, indicating uniform pressure distribution and high simulation precision. The average stabilized pressures were 0.3986 MPa, 0.4990 MPa, and 0.6204 MPa, respectively, with a maximum relative error of 3.4%.

3.4.2. Two-Phase Flow Environment

To verify the capability of two-phase flow simulation, ablation tests were conducted on the silicone rubber TPS materials under two-phase flow environments in conjunction with the particle loading system. Al2O3 particles with a diameter of 5 μm were used, with a nitrogen carrier gas flow rate of 1 L/min. After mixing with the combustion gas, the particle concentration reached 0.16 kg/m3. The numerical simulation of Ansys Fluent (2020R2) demonstrates that the gas temperature upon reaching the specimen surface is 2367 K. Photographs of the specimens after ablation under particle-laden and particle-free conditions are presented in Figure 13.
This work focuses on the design and validation of the ablation experimental device. A horizontal comparison of ablation characteristics under different environments for the same type of material has not yet been performed. Subsequent research on the ablation performance of thermal protection materials will incorporate this work.

4. Conclusions

An oxygen–kerosene ablation experimental device, encompassing a main control system, oxygen supply system, kerosene supply system, water cooling system, pressure chamber, and particle loading system, was utilized to determine the suitable oxygen-to-fuel ratio. This ratio enables better simulation of the thermal ablation environment in the afterburning chamber under maneuvering flight conditions, bringing the simulated environment closer to actual operational conditions.
Additionally, numerical simulations were employed to obtain variation characteristics of flow field parameters within the device, thereby enabling effective control of parameters such as pressure in the afterburning chamber and surface temperature of the thermal insulation layer. To verify the reliability of this ablation test system, multiple ablative materials (both of the same type and different types) were selected for ablation experiments. The experimental results are as follows:
(1)
When the oxygen-to-fuel ratio exceeds the theoretical value and continues to increase, combustion gas parameters (including combustion temperature, gas constant, and specific heat ratio) exhibit a decreasing trend, whereas oxygen concentration increases gradually. Selecting an oxygen-to-fuel ratio of 3.5 results in conditions that approximate the actual operational scenarios of a typical ramjet.
(2)
Ablation experimental data indicate that for identical TPS materials, the maximum relative error in ablation rate is only 5.67%. This fully demonstrates that the ablation test system is reasonably and reliably designed, meeting the requirements for ablation experiments on thermal insulation materials.
The ablation experimental device designed in this work successfully simulates actual operating conditions in the afterburning chamber. It is of great significance for subsequent research on ablation performance of thermal insulation materials and enables in-depth investigation of material performance under real operational conditions.

Author Contributions

Conceptualization, X.F.; methodology, J.C.; writing—original draft preparation, J.C.; data curation, H.Y.; investigation, X.Q.; writing—review and editing, G.Z. and J.L.; project administration, X.F. All authors have read and agreed to the published version of the manuscript.

Funding

This work was supported by the National Natural Science Foundation of China [grant number 52202481], the China Postdoctoral Science Foundation [grant number 2022MD723835].

Data Availability Statement

The data presented in this study are available on request from the corresponding author.

Conflicts of Interest

The authors declare no conflict of interest.

Abbreviations

AbbreviationFull name
TPSThermal Protection System
C/CCarbon/Carbon Fiber
SiCSilicon Carbide
CEAChemical Equilibrium Application
EPDMEthylene Propylene Diene Monomer
PICPhenolic Impregnated Carbon
PRPhenolic Resin

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Figure 1. Oxygen-enriched ablation experimental device. 1, control system; 2, oxygen supply system; 3, kerosene supply system; 4, oxygen–kerosene liquid rocket engine; 5, outlet extension section; 6, water cooling system; 7, particle loading system; 8, specimen holders.
Figure 1. Oxygen-enriched ablation experimental device. 1, control system; 2, oxygen supply system; 3, kerosene supply system; 4, oxygen–kerosene liquid rocket engine; 5, outlet extension section; 6, water cooling system; 7, particle loading system; 8, specimen holders.
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Figure 2. TPS material specimen fixture along with assembly diagram.
Figure 2. TPS material specimen fixture along with assembly diagram.
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Figure 3. Sectional view of the front of the TPS material specimen fixture.
Figure 3. Sectional view of the front of the TPS material specimen fixture.
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Figure 4. The grid diagram of the flow field.
Figure 4. The grid diagram of the flow field.
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Figure 5. Particle loading system, (a) rotary-disc powder feeder, (b) controller.
Figure 5. Particle loading system, (a) rotary-disc powder feeder, (b) controller.
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Figure 6. Device for pressure environment simulation, (a) and (b) pressure environment simulation, (c) top view of the pressure chamber, (d) internal view of the pressure chamber, (e,f) jacketed water-cooling structure.
Figure 6. Device for pressure environment simulation, (a) and (b) pressure environment simulation, (c) top view of the pressure chamber, (d) internal view of the pressure chamber, (e,f) jacketed water-cooling structure.
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Figure 7. Plots of various combustion chamber parameters as a function of the oxygen-to-fuel ratio, (a) temperature, (b) gas constant, (c) gas specific heat ratio, (d) oxygen content.
Figure 7. Plots of various combustion chamber parameters as a function of the oxygen-to-fuel ratio, (a) temperature, (b) gas constant, (c) gas specific heat ratio, (d) oxygen content.
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Figure 8. Simulation and characterization of the rocket engine exhaust temperature, (a) ablation experimental device, (b) external specimen, (c) temperature results from simulation, (d) simulated temperature of the specimen surface, (e) temperature measurement technique, (f) measured temperature of flame region. (The dashed arrow indicates the position of the bevel).
Figure 8. Simulation and characterization of the rocket engine exhaust temperature, (a) ablation experimental device, (b) external specimen, (c) temperature results from simulation, (d) simulated temperature of the specimen surface, (e) temperature measurement technique, (f) measured temperature of flame region. (The dashed arrow indicates the position of the bevel).
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Figure 9. Morphologies of TPS materials after ablation experiments.
Figure 9. Morphologies of TPS materials after ablation experiments.
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Figure 10. Back wall temperature data of TPS materials after the ablation experiments (Curves 1, 2, and 3 correspond to the three thermocouples), (a) Back wall temperature of silicone rubber 1-#1, (b) Back wall temperature of silicone rubber 2-#1, (c) Back wall temperature of PR#1, (d) Back wall temperature of silicone rubber 1-#2, (e) Back wall temperature of silicone rubber 2-#2, (f) Back wall temperature of PR#2.
Figure 10. Back wall temperature data of TPS materials after the ablation experiments (Curves 1, 2, and 3 correspond to the three thermocouples), (a) Back wall temperature of silicone rubber 1-#1, (b) Back wall temperature of silicone rubber 2-#1, (c) Back wall temperature of PR#1, (d) Back wall temperature of silicone rubber 1-#2, (e) Back wall temperature of silicone rubber 2-#2, (f) Back wall temperature of PR#2.
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Figure 11. Morphologies and back wall temperatures of TPS materials after ablation experiments on EPDM and polyphosphazene (PIC represents Phenolic Impregnated Carbon), (a,b) morphology of EPDM materials, (c) morphology of PIC material, (d,e) back wall temperatures of EPDM materials, (f) back wall temperatures of PIC materials.
Figure 11. Morphologies and back wall temperatures of TPS materials after ablation experiments on EPDM and polyphosphazene (PIC represents Phenolic Impregnated Carbon), (a,b) morphology of EPDM materials, (c) morphology of PIC material, (d,e) back wall temperatures of EPDM materials, (f) back wall temperatures of PIC materials.
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Figure 12. Pressure curves from the two measurement points within the pressure chamber, (a) pressurized condition of 0.4 MPa, (b) pressurized condition of 0.5 MPa, (c) pressurized condition of 0.6 MPa.
Figure 12. Pressure curves from the two measurement points within the pressure chamber, (a) pressurized condition of 0.4 MPa, (b) pressurized condition of 0.5 MPa, (c) pressurized condition of 0.6 MPa.
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Figure 13. Photographs of silicone rubber TPS materials after ablation, (a) without particles, (b) with particles.
Figure 13. Photographs of silicone rubber TPS materials after ablation, (a) without particles, (b) with particles.
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Table 1. Combustion chamber parameter calculations (oxygen-to-fuel ratio is 6).
Table 1. Combustion chamber parameter calculations (oxygen-to-fuel ratio is 6).
ParameterValue
Pressure (kPa)900
Temperature (K)3206.57
Density (kg/m3)0.987
Specific heat ratio1.123
Molar fraction of O20.36645
Molar fraction of H20.00893
Molar fraction of OH0.07930
Molar fraction of H2O0.18900
Molar fraction of CO0.08151
Molar fraction of CO20.22125
Molar fraction of O0.04432
Molar fraction of HO20.00022
Molar fraction of H2O20.00001
Molar fraction of H0.00902
Table 2. Distribution of the mass ablation rates of TPS material specimens.
Table 2. Distribution of the mass ablation rates of TPS material specimens.
Test PieceQuality Before Experiment (g)Quality After Experiment (g)Mass Ablation Rate
(g/s)		Average ValueRelative Error
Silicone rubber 1#153.8440.540.0222
#253.7139.710.02330.022752.42%
Silicone rubber 2#155.4546.480.0149
#255.7947.830.01330.01415.67%
Phenolic resin#163.3430.150.0553
#263.8832.570.05220.053752.88%
Table 3. Temperatures (K) of TPS materials specimens at different times.
Table 3. Temperatures (K) of TPS materials specimens at different times.
Title of the Sample0 s100 s200 s300 s400 s500 s600 s
Silicone rubber 1#1311.55347.98421.09503.25574.42633.06679.82
#2315.61352.32426.34504.27576.72636.98686.19
Average value 313.58350.15423.72503.76575.57635.02683.01
Relative error 0.65%0.62%0.62%0.10%0.20%0.31%0.47%
Silicone rubber 2#1306.97329.19386.23524.60685.68849.051075.43
#2308.58335.12389.85480.14631.73803.18984.57
Average value 307.78332.16388.04502.37658.71826.121030
Relative error 0.26%0.90%0.47%4.4%4.1%2.8%4.4%
Phenolic resin#1309.42375.12502.44650.53932.711046.51080.8
#2311.09348.58393.57480.19600.881066.91104.7
Average value 310.26361.85448.01565.36766.801056.71092.8
Relative error 0.27%3.7%12.2%15.1%21.6%1.0%1.1%
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MDPI and ACS Style

Cheng, J.; Yan, H.; Feng, X.; Zhu, G.; Liu, J.; Qi, X. Research on Ablation Device Suitable for Thermal Protection System of Solid Rocket Ramjet. Aerospace 2025, 12, 772. https://doi.org/10.3390/aerospace12090772

AMA Style

Cheng J, Yan H, Feng X, Zhu G, Liu J, Qi X. Research on Ablation Device Suitable for Thermal Protection System of Solid Rocket Ramjet. Aerospace. 2025; 12(9):772. https://doi.org/10.3390/aerospace12090772

Chicago/Turabian Style

Cheng, Jiming, Hang Yan, Xiping Feng, Guoqiang Zhu, Jie Liu, and Xintong Qi. 2025. "Research on Ablation Device Suitable for Thermal Protection System of Solid Rocket Ramjet" Aerospace 12, no. 9: 772. https://doi.org/10.3390/aerospace12090772

APA Style

Cheng, J., Yan, H., Feng, X., Zhu, G., Liu, J., & Qi, X. (2025). Research on Ablation Device Suitable for Thermal Protection System of Solid Rocket Ramjet. Aerospace, 12(9), 772. https://doi.org/10.3390/aerospace12090772

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