Chemical Propulsion Systems Research Articles

Effect on the performance of a hybrid rocket using AP in the PVC and HTPB-based hybrid fuel

Hybrid rocket is a type of chemical rocket propulsion system where the fuel and the oxidizer are stored separated and in different state of aggregation. Typically, the fuel is in the solid state and the oxidizer is either in the liquid or gaseous state. It offers various advantages over other type of chemical rockets such as throattability, controllability and reduced complexity. It also has some drawbacks such as low regression rate, poor combustion efficiency, higher sliver losses, etc. Various methods have been adopted to overcome these drawbacks and improve the performance of hybrid rocket. The present study focus on the improvement of regression rate with the addition of oxidizer particles in fuel content. Ammonium Perchlorate is used as an additive in Hydroxyl Terminated Polybutadiene (HTPB) and Poly Vinyl Chloride (PVC) based fuel grain. From the experimentations, it was observed that the thrust, pressure and regression rate of the HTPB fuel grain with additive have higher improvement than with the PVC based fuel grain. The sliver losses are also reduced due to uniform burning after addition of oxidizer in fuel grain. The overall study strongly suggests that addition of oxidizer in limited quantity can be used to overcome the drawbacks of hybrid propellant rocket.

Electrical propulsion systems for satellites: a review of current technologies and future prospects

Electrical propulsion systems have revolutionized satellite technology by offering greater efficiency, longer mission durations, and increased maneuverability compared to traditional chemical propulsion systems. This review explores current technologies and future prospects in the field of electrical propulsion systems for satellites. The discussion begins with an overview of existing technologies, including ion propulsion systems, Hall effect thrusters, and pulsed plasma thrusters. Each technology's principles of operation, advantages, limitations, and notable applications are examined. The review delves into the future prospects of electrical propulsion systems, exploring advanced concepts such as magnetoplasmadynamic thrusters, variable specific impulse magnetoplasma rockets, and electrospray propulsion systems. Additionally, miniaturization and efficiency improvements, as well as sustainable and green propulsion alternatives, are discussed. Challenges and opportunities facing the field are addressed, including technical hurdles like power generation and management, thruster lifespan, and regulatory and economic considerations such as policy frameworks and market dynamics. In conclusion, the review underscores the critical role of continued research and development in electrical propulsion systems for satellites. As the demand for more capable and sustainable satellite missions grows, advancements in propulsion technology will be essential in meeting these evolving needs and pushing the boundaries of space exploration.

A Methodology for the Preliminary Design of a High-Efficiency Multistage Plasma Thruster

AbstractSpace electric propulsion represents a class of power-limited systems that utilize the interaction of electromagnetic fields with ionized inert gas propellants to generate thrust. This technology has emerged as a highly fuel-efficient and sustainable alternative to chemical propulsion systems, particularly for satellite constellations. However, the miniaturization potential of EP systems is impeded by certain limitations, necessitating the exploration of novel architectures. The high-efficiency multistage plasma thruster (HEMP-T) stands as a promising contender for stand-alone missions due to its employment of a cusped magnetic-field topology, which effectively mitigates plasma-wall interactions and enhances overall efficiency even at low thrust levels. Despite the growing interest in HEMP-Ts, there is a dearth of comprehensive and streamlined preliminary design procedures for these systems. Prior research has predominantly focused on extensive numerical analyses, neglecting the development of efficient and accessible design tools. To bridge this gap, this paper presents a novel preliminary design tool derived from integrating established analytical models available in the literature. The proposed design tool also incorporates an iterative procedure that refines geometric properties using a 2D magnetostatic solver. Through the application of this tool, a 4 mN HEMP thruster was analyzed. This finally exhibited a specific impulse of approximately 2000s and a good efficiency level of 23%. Also, the results obtained for a 10 mN application align closely with those achieved by other types of EP thrusters.

Optimizing Propellant Distribution for Interorbital Transfers

The advent of space exploration missions, especially those aimed at establishing a sustainable presence on the Moon and beyond, necessitates the development of efficient propulsion and mission planning techniques. This study presents a comprehensive analysis of chemical and electric propulsion systems for spacecraft, focusing on optimizing propellant distribution for missions involving transfers from Low-Earth Orbit (LEO) to Geostationary Orbit (GEO) and the Lunar surface. Using mathematical modeling and optimization algorithms, we calculate the delta-v requirements for key mission segments and determine the propellant mass required for each propulsion method. The results highlight the trade-offs between the high thrust of chemical propulsion and the high specific impulse of electric propulsion. An optimization model is developed to minimize the total propellant mass, considering a hybrid approach that leverages the advantages of both propulsion types. This research contributes to the field of aerospace engineering by providing insights into propulsion system selection and mission planning for future exploration missions to the Moon, Mars, and Venus.

Optimizing a Hall thruster with aft-loaded magnetic field by aft-loading design of gas flow

Electric propulsion systems with high-specific impulses have replaced chemical propulsion systems in many space missions. The aft-loaded magnetic field technology of Hall thrusters moves the ionization and acceleration processes downstream of the channel by varying the magnetic field distribution. Consequently, ion bombardment of the channel walls is inhibited, and the service life of the Hall thrusters is substantially increased. However, neutral gas leaks because of the rapid diffusion near the channel outlet, where strong ionization occurs. Thus, the propellant atoms are ionized insufficiently, resulting in a decline in the performance. Therefore, an aft-loading design for the neutral gas flow field was developed to optimization distribution of the neutral gas in this study. A 1.5 kW Hall thruster was verified experimentally. The experimental results show that the aft-loaded neutral gas flow field significantly improved the ionization and acceleration processes. The propellant utilization efficiency increased by 4–8% within the discharge voltage range of 300–900 V. The ion energy distribution moved toward the high-energy direction significantly, promoting an overall improvement in the performance, with a thrust increase of 7.3–11.5 %. The anode efficiency increased from 49–57 % to 58–61 %. The results of this study are significant for the design of high-performance, long-life Hall thrusters.

Numerical Simulation of Chemical Propulsion Systems: Survey and Fundamental Mathematical Modeling Approach

This study deals with the mathematical modeling and numerical simulation of chemical propulsion systems (CPSs). For this, we investigate and summarize a comprehensive collection of the simulation modeling developments of CPSs in academic works, applications, and industrial fields. Then, we organize and analyze the simulation modeling approaches in several ways. After that, we organize differential-algebraic Equations (DAEs) for fundamental mathematical modeling consisting of the governing Equations (ordinary differential equations, ODEs) for the components and other equations derived from several physical rules or characteristics (algebraic equations or phenomenological equations, AEs) and then synthesize and summarize the fundamental structures of analytic mathematical modeling by types (liquid-propellant rocket engines, solid-propellant rocket motors, and hybrid-propellant rocket motors) of CPSs.

Trajectory design and dispersion analysis of nano moon lander OMOTENASHI

OMOTENASHI is a CubeSat that was launched by a NASA SLS rocket. Its mission was to demonstrate that a CubeSat could make a semi-hard landing on the Moon. A chemical propulsion system which provides high thrust and can be re-ignited is large and difficult to install on a CubeSat. Therefore, a small cold-gas jet makes orbit corrections with only a small ΔV. Moreover, a rocket motor with a high Isp and structural mass ratio is used for a lunar landing deceleration. The spacecraft consists of an orbiting module, a rocket motor, and a surface probe. This paper analyzes the trajectory design of the Moon landing and a method that increases the landing success rate under the tight constraints of the CubeSat. To reduce fuel and power consumption, the maneuver conditions during the lunar transition to reach the target landing site are evaluated based on several error factors. This study considers the trade-off between vertical and horizontal landings. Because active control is not possible for decelerating a rocket motor for the lunar landing, the optimal altitude at the starting maneuver is determined by considering such errors as attitude and timing. Finally, the paper discusses mission feasibility and the landing success rate.

Optimization of the Maneuver to Ensure a High Velocity of the Spacecraft Entry into the Atmosphere

The transfer scheme that provides a parabolic entry of the spacecraft into the Earth’s atmosphere has been optimized. Such a maneuver can be of interest in experimental testing of the spacecraft reentry from the Moon or after interplanetary missions. It is assumed that the spacecraft is inserted into a low Earth orbit and is equipped with a chemical propulsion system and a limited-thrust engine, which should provide a maneuver to bring the spacecraft into the Earth’s atmosphere. The optimization criterion takes into account the characteristic velocity of the maneuver. The developed method of optimizing the transfer scheme and the spacecraft trajectory itself is based on the maximum principle. Single-revolution and multi-revolution transfer trajectories are analyzed. It is shown that for single-revolution trajectories there is an optimal time and an optimal angular distance of flight. Their values and the minimum characteristic velocity of the maneuver are evaluated. Unlike single-revolution trajectories, the characteristic velocity for multi-revolution trajectories monotonically decreases with increasing transfer duration. The dependence of the characteristic velocity on the transfer duration for single-, two-, three- and four-revolution trajectories is given. The transfer duration ranges in which it is advisable to use each type of trajectory are analyzed.

A predictive model of Lunar Gateway molecular outgassing and plume-induced contamination

The Gateway space station will operate both as a lunar outpost and a hub for space exploration missions in support of a long-term human presence on the Moon, and is under development by NASA in collaboration with ESA, CSA, JAXA, and international commercial partners. Gateway will experience induced molecular contamination from sources including materials outgassing, venting, and the nominal operation of its chemical and electric propulsion systems. The degradation effects of molecular contamination on spacecraft are well-documented and understood, and predictive modeling tools have been developed to aid in the mitigation of such contamination effects in the mission design phase. Therefore, the Gateway Induced Environments and Thermal teams have worked with JPL Contamination Control to develop modular simulation tools for the prediction of molecular contaminant depositions across the mission life, and for use in requirements derivation and sensitivity assessments. Free-molecular transport and deposition of outgassing products onto sensitive Gateway receiver surfaces is calculated using a viewfactor matrix approach, while International Space Station (ISS)-heritage bipropellant plume models are used to evaluate contaminant fluxes generated by the operation of thrusters on Gateway modules and the Orion spacecraft. The alteration of optical properties, e.g. the solar absorptance, of contaminated surfaces is likewise calculated using an ISS-heritage semi-empirical model premised on flight and laboratory testing data. This framework is intended to support early identification of potential integration issues and establish a baseline for incorporating improved analysis and test data as the Gateway design matures.

On the selection of propellants for cold/warm gas propulsion systems

Cold/warm gas spacecraft propulsion systems generate thrust by accelerating a cold or moderately heated gas through a nozzle. Despite their simplicity compared with typical chemical or electric propulsion systems, the choice of propellant nonetheless represents an important consideration that affects the design, operation, and performance of such thrusters. Here we review different propellants that have been used or proposed for cold/warm gas propulsion systems, and we perform a study investigating almost 5000 other possible alternatives. Propellants include organic and inorganic substances that can be stored in solid, liquid, gaseous, and multi-phase states at ambient conditions, and which undergo a phase change to a pure gaseous state (where needed) by heating before being ejected from the thruster. The different propellants are assessed by considering important evaluation factors, such as propulsive performance, storage requirements and storage density, mission suitability, safety, and cost. While conventional thruster performance metrics like the specific impulse are low for some propellants, the design and operational advantages they offer, together with a higher total impulse mass density (i.e. the impulse per propulsion system wet mass), make them superior choices for many missions.

Enabling Interplanetary Exploration for CubeSats with a Fully Chemical Propulsion System

Interplanetary CubeSat missions are currently becoming more popular, with a significant number of recently planned missions. The context of this paper is a Mars mission, starting from a parking orbit around Earth: the adoption of a chemical propulsion system for the Earth-Mars transfer phase is investigated, considering the recent technological developments for CubeSats. A trade-off of propulsion system type and propellant results in the choice of a mono-propellant system adopting the HAN-based propellant AF-M315E (ASCENT). The main challenge for the propulsion system is to fit inside a CubeSat standardised volume, which can range up to 24 U, for which the implementation of a suitable COTS micro-pump is considered. Finally, the complete architecture and design of the propulsion system is presented. This work demonstrates the feasibility of adopting full chemical propulsion for an interplanetary CubeSat mission, with consequent advantages in terms of transfer time and required power, at the cost of relatively small mass and volume left for the other subsystems. Even better results can be expected for interplanetary missions requiring slightly lower ΔV budgets, such as Near Earth Objects exploration or asteroid fly-by missions. Keywords: CubeSat, Chemical Propulsion, Green Mono-propellant, Interplanetary Mission, Mars Exploration

Development and tests of a thrust stand with an in-situ null position adjustment and calibration method for low power plasma thrusters

A growing number of space missions use electric propulsion systems due to their higher specific impulse and greater efficiency compared to chemical propulsion systems. Because they are able to generate relatively low thrust values, conventional thrust measurement mechanisms used for chemical propulsion systems can not be used for electric propulsion systems. Thus, millinewton level thrust measurement systems are needed to be developed for measuring the low-level thrust of these types of thrusters. At the Bogazici University Space Technologies Laboratory (BUSTLab), a double inverted pendulum type thrust stand with a changeable thrust measurement range between 6 mN – 300 mN was designed, built and tested. The measurement resolution of this thrust stand can be as low as 100 μN level and it can be adjusted by changing the flexure stiffness and the mass of a counterweight. This study investigates the BUSTLab thrust measurement system and its key design points, with a particular focus on its in-situ null position adjustment and calibration process, and methods for mitigating external perturbations during thrust measurements.

Heuristic Technique for the Search of Interception Trajectories to Asteroids with the Use of Solar Sails

With the advances of space-exploration technologies, a long-lasting concern is finally being addressed: the deflection of potentially hazardous objects (PHOs). Most recently, the first mission of this kind was launched by NASA—the Double Asteroid Redirection Test (DART). Nevertheless, it is estimated that a great number of these PHOs are unattainable by means of current chemical propulsion systems. With that in mind, this study proposes the development of a heuristic technique for the search of interception trajectories with the use of solar sails and its application in determining a set of possible trajectories to intercept hazardous asteroids. As a case study, a hybrid mission inspired by the DART mission is proposed by using a solar sail as means of propulsion after the initial chemical combustion. The dynamics consider a model of the solar radiation pressure acceleration as a function of the orientation of the sail. In turn, the orientation is defined by the application of the developed heuristic technique with the goal of defining alternative trajectories compared to the original mission. These trajectories result in different impact conditions and mission durations. Although the use of solar sails breaks the symmetry in the solutions, the results obtained offer the possibility of fuel economy or even better deflection results by the achievement of greater impact energy with the hazardous objects.

Propulsion Alternatives for Mars Transportation Architectures

Recently, NASA has pushed for returning humans to the moon, with in-situ resource utilization being the key capability to provide sustainability. One of the potential future developments could be a propellant depot in lunar distant retrograde orbit. Using Aerojet Rocketdyne Mars mission architectures and University of Alabama in Huntsville Nuclear Thermal Propulsion (NTP) engine models, this research analyzed the impacts of using chemical and engines as well as the liquid-oxygen (LOX) Augmented Nuclear Thermal Rocket (LANTR) engines for these missions and compared their performances to the reference hydrogen-based NTP (H-NTP) engines all the while assuming a propellant depot at lunar distant retrograde orbit. For a human mission to Mars originating in the lunar distant retrograde parking orbit, the LANTR engines will offer better overall performance than H-NTP engines with a predicted 55.6% decrease in propellant volume, 39% decrease in vehicle dry mass, and 50% decrease in the number of aggregation launches. This is due to LANTR’s 22% higher specific impulse than conventional chemical propulsion systems, three times higher density than pure hydrogen, and 440% higher thrust than the baseline H-NTP engines. However, these benefits come at the cost of the propellant mass, which is 32.4% higher for the conjunction class mission and 106.7% higher for the opposition class mission than the baseline H-NTP system.

Hugoniot Relation for a Bow-Shaped Detonation Wave Generated in RP Laser Propulsion

Repetitive-pulsed (RP) laser propulsion is expected to replace chemical propulsion systems because it can reduce launch costs. A laser-supported detonation wave (LSD) plays an important role in the thrust-generation process of RP laser propulsion. The LSD propagation mechanism has been studied. Nevertheless, the LSD propagation velocity measured in an earlier study was lower than the Chapman–Jouguet (CJ) velocity, which meant that Hugoniot analysis produced no solution. The findings suggest that the radial flow from the central axis of LSD exerts some effects, but it has not been evaluated quantitatively. Two-dimensional axisymmetric computational fluid dynamics (CFD) analysis using the measured propagation velocity was performed for this study to evaluate effects of the radial flow of a bow-shaped LSD. Results show that the ratios of the radial flow of mass, momentum, and enthalpy from the central axis can be calculated, respectively, as 0.82, 0.13, and 0.17. Additionally, the measured propagation velocity of a bow-shaped LSD was shown to be higher than the CJ velocity calculated using the two-dimensional axisymmetric CFD reproducing the experiment conditions.

Catalytic performance of nano-sized cobalt copper ferrite on thermal decomposition of ammonium nitrate

Ammonium nitrate (AN) is being explored as a high energetic oxidizer in propellants. The utilization of pure AN in propellants suffers from a low burning rate. To improve the burning rate, it is important to improve the thermal decomposition characteristics of AN to fasten its thermal decomposition. The effect of CoCuFe2O4 additive was explored to improve the thermos-kinetic parameters of AN's thermal decomposition. Coprecipitation was used to synthesize nanocrystalline CoCuFe2O4. The thermodynamic and kinetic parameters of decomposition of AN: CoCuFe2O4 were investigated using TG-DSC-DTA analysis at 5, 10, and 15 °C min−1. The thermodynamic and kinetic studies suggest that the thermal decomposition of AN+CoCuFe2O4 was more favorable than AN. The decomposition of AN+CoCuFe2O4 occurs at a lower 10 °C lower temperature with ∼44 % decrement in the activation energy and decreased pre-exponential factor. The faster decomposition at a lower temperature was obtained for AN+CoCuFe2O4 composition. Nano-sized particles with efficient optical and polarizable properties made this ferrite material easily separable and eco-friendly for the preparation of a chemical propulsion system based on AN. The AN+CoCuFe2O4 propulsion mixture can be utilized in place of pure AN for better thermal performance in a propellant system.•CoCuFe2O4 was able to influence the thermal decomposition parameters of ammonium nitrate (AN).•AN+CoCuFe2O4 decomposes at lower activation energy with a faster decomposition than pure AN.•Thermodynamics suggest favorable decomposition of AN+CoCuFe2O4 than pure AN.

Recent concept study for cryogenic fluid management to support opposition class crewed missions to Mars

NASA recently completed a mission concept study to evaluate the feasibility and propulsion technology development requirements for reduced travel duration crewed missions to Mars. A high-level goal of the study was to minimize the health impact on the crew caused by the space environment. This was implemented in the study by limiting the crew to a total of approximately-two years of in-space operations and travel time. For the initial mission, the crew would stay about 30 days on the Martian surface. The propulsive demands of such a mission are immense, and the study identified two advanced propulsion options with the potential to meet the mission requirements—both options rely on nuclear fission to provide efficient propulsive energy. One propulsion option was a nuclear electric propulsion (NEP)/Chem Hybrid, with a reactor and energy conversion system powering xenon propellant ion thrusters to provide an efficient, but lower-thrust, push for most of the mission duration. This concept also relied on a liquid oxygen/liquid methane (LO2/LCH4) chemical propulsion stage to provide high thrust for maneuvers while near the Earth and Mars. The second propulsion option was nuclear thermal propulsion (NTP), in which the reactor heats liquid hydrogen (LH2) propellant to expand through a nozzle for thrust at about twice the efficiency of the best chemical propulsion systems. Both vehicle concepts rely on storing large amounts of cryogenic propellant (either LO2/LCH4 or LH2) for multiple years in space without loss, far exceeding state-of-the-art capability. To enable this new capability, the team assumed the use of several advanced cryogenic fluid management (CFM) technologies and analyzed the integrated system performance. This included considering the vehicle-level effects of the size, mass, and power requirements of these CFM elements. Further, the team evaluated the development required to enable such a mission in the mid-2030 s and determined that it was feasible. The paper elaborates on the assumed CFM technologies, provides key analysis results, and illustrates the feasibility of technology development for the proposed solutions to the CFM challenges for each propulsion concept.

Experimental characterization of the electrospray propulsive performance for ionic liquid propellants [EMIm][DCA] and [BMIm][DCA]

A chemical-electrospray dual-mode propulsion system is significant in improving mission capability and flexibility of a spacecraft. Hypergolic ionic liquids are potential propellants of chemical and electrospray propulsion systems. In this work, the electrical propulsive performance of two hypergolic ionic liquids 1-ethyl-3-methylimidazolium dicyanamide ([EMIm][DCA]) and 1-butyl-3-methylimidazolium dicyanamide ([BMIm][DCA]) are experimentally characterized using two array-emitter electrospray thrusters. Compared with a commonly used ionic liquid propellant 1-ethyl-3-methylimidazolium tetrafluoroborate ([EMIm][BF4]), the current magnitude and current steadiness of the electrospray thruster with [EMIm][DCA] and [BMIm][DCA] are efficiently improved, especially with [EMIm][DCA] as propellant. The results show that the low viscosity is beneficial for steady and sufficient flow supply and thus favors a long-time steady operation of the thruster. For the propellant [EMIm][DCA], the maximum specific impulse is 1791 s, i.e., 27.8% higher than that of [EMIm][BF4], while the maximum thrust 32.5 μN is 128.9% greater than that of [EMIm][BF4]. For the propellant [BMIm][DCA], the thrust-to-power ratio of the thruster is approximately 82.8% higher than that of [EMIm][BF4] at a similar specific impulse of about 1120 s resulting from a much greater thruster efficiency. These results demonstrate that [EMIm][DCA] has a better performance than [EMIm][BF4] in both maximum thrust and specific impulse, and [BMIm][DCA] performs best in the aspect of thrust-to-power ratios. This work highlights the prospect of hypergolic ionic liquids and is valuable to promote their application in chemical and electrical propulsion communities.

An Optimized Plane-Change Solution for Microsatellite Formation Flying

The cross-track baseline requirement in a satellite formation flying mission requires an orbital plane-change maneuver during the acquisition phase. While the optimization of orbital plane-change maneuver has been well studied, the constraints of the propulsion system, such as the power limitation, battery charging cycle, and finite thrust durations, are often neglected. In addition, the self-pressurized chemical propulsion system used in this satellite mission experiences a performance degradation effect over prolonged use in terms of specific impulse and thrust force. To overcome these constraints on the orbital plane-change maneuver process, this article proposes an optimized algorithm by minimizing Edelbaum’s equation using a selective constrained ensemble Kalman filter that includes propulsion’s performance variation and firing duration constraints. Monte Carlo simulations have been conducted to benchmark the proposed method against the theoretical impulsive thrusts in terms of <inline-formula xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink"> <tex-math notation="LaTeX">$\Delta V$ </tex-math></inline-formula> , and the commonly adopted natural precession-based plane-change method in terms of the total time taken. Results from the simulation have shown that the required <inline-formula xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink"> <tex-math notation="LaTeX">$\Delta V$ </tex-math></inline-formula> of this proposed method is similar to the theoretical impulsive burn method. In addition, it requires a 75% shorter time than the natural precession method in achieving the desired plane change and is 36% more fuel efficient than the direct firing method. Furthermore, this study has verified that this algorithm can be directly integrated with the existing formation control algorithm.

Full-Core Coupled Neutronic, Thermal-Hydraulic, and Thermo-Mechanical Analysis of Low-Enriched Uranium Nuclear Thermal Propulsion Reactors

Nuclear thermal propulsion is an enabling technology for future space missions, such as crew-operated Mars missions. Nuclear thermal propulsion technology provides a performance benefit over chemical propulsion systems by operating with light propellants (e.g., hydrogen) at elevated engine chamber conditions. Therefore, nuclear thermal propulsion reactor cores exhibit high propellant velocities and elevated propellant and fuel temperatures, subsequently leading to relatively high thermal stresses and geometrical deformation. This paper details the numerical approach to solve the thermo-elastic equations, which was implemented into the recently developed ntpThermo code. In addition, this paper demonstrates the extension of the Basilisk multiphysics framework to perform full-core coupled neutronic, thermal-hydraulic, and thermo-mechanical analysis of nuclear thermal propulsion reactors. The analyses demonstrate and quantify thermo-mechanical feedback, which for the investigated cases, acted to reduce maximum fuel temperatures and pressure drop across the fuel element channels. Thermo-mechanical feedback had a significant impact on the mass flow distribution within the reactor core and, thus, a substantial impact on solid-material temperatures and stresses, but only a minor impact on reactivity and local power distributions. Sensitivity studies revealed that the friction factor correlation applied to perform the analysis has a significant impact on the pressure drop across the fuel element channels. The most important observation of this research is the importance of incorporating the thermo-mechanical feedback within an integrated multiphysics solution sequence to enable the consistent design of future nuclear thermal propulsion systems.