Theoretical Specific Impulse Research Articles

Advanced Al Hydrogel Propellants: Preparation, Thermal Stability, and Self-Sustainable Combustion Characteristics.

A new group of Al-based hydrogel propellants has been prepared and evaluated, aiming to replace currently used Al-ice propellants. The theoretical specific impulse (Isp) of these formulations is around 235 s. To enhance the ignition and combustion performances of these propellants, a 2% metastable intermixed composite Al@PVDF-CuO and minor ammonium perchlorate (AP) were introduced. The measured maximum ignition temperature of these hydrogel propellants was 2277 °C, higher than that of the Al-ice propellant (1700 °C). This increase in the ignition temperature could be attributed to the enhanced reactivity and energy release provided by the additives. Additionally, the initial mass loss temperature for this kind of hydrogel propellant, ranging from 211 to 235 °C, indicates higher thermal stability compared to the evaporation temperature of moisture water, indicating the strong water-locking capability of the gelling agent triaminoguanidine-glyoxal polymer (TAGP). The two main reaction steps involved in the combustion process are the oxidation of Al by H2O with H2 release, which coincides with the decomposition of TAGP at 246 °C, and a subsequent oxidation step with remaining H2O and TAGP residues peaking at 320 °C. The burn rate of these propellants is stable and controllable, ranging from 2.5 to 5.9 mm·s-1, which is crucial for maintaining a consistent thrust during propulsion. The combustion condensed products are predominantly α-Al2O3, which could influence the overall combustion efficiency and thermodynamic properties of the system. The effect of AP on the Al-water reaction has been further studied using reactive force field (ReaxFF) molecular dynamic simulations. It has been shown that the presence of AP may lead to a slight decrease in H2 production, which was attributed to the increase in the H2O concentration as the major decomposition product of AP.

Vaporization in Liquid-Core Nuclear Thermal Rockets

This work provides an overview of vaporization phenomena in liquid-core nuclear thermal rockets to aid in unifying modeling assumptions by characterizing the impact of vaporization. Historical variations in vapor modeling have led to reported specific impulses ranging from 1000 to 2000s following model refinements in the 1960s and the omission of vaporization consideration post 1990. The only preexisting vapor transport study to not rely on the Reynold’s analogy was performed for the radiator design. This paper performs the first such analysis for the bubbler design, given its renewed interest. Consideration is given to potential centrifugal species separation, proposed fuel mixtures, interplay between chamber pressure and vaporization, impacts on propellant thermoproperties, evaporative cooling, power requirements for vapor separation, and the complexity of hydrocarbon interactions under the proposed chamber conditions. The bubbler design is shown to be well approximated as fully saturated, whereas the radiator and droplet designs may be configured to reduce vaporization to 20% saturation. Characterization of vaporization within liquid cores is essential to early design decisions, such as overall archetype, defining the operating point, and liquid media composition. For a threshold temperature corresponding to around 10% of the propellant being vapor by mass, further increases in temperature begin reducing the specific impulse. Below this threshold temperature, the impact of vaporization on thermal modeling with the chamber is likely negligible. Above this threshold temperature, higher-fidelity thermal modeling will be required, and the work required to separate the vapor begins exceeding 100 kW/(kgH/s). Nozzle kinetics are applied to liquid cores for the first time, showing that specific impulses increase with increasing chamber pressure contrary to historical findings. Historically, modeled uranium-zirconium-carbon mixtures are predicted to reach a specific impulse up to 1260 s. A uranium-tungsten-carbon fuel mixture yielded the best theoretical specific impulse of slightly under 1400 s; however, no nucleonic or thermal analysis has been conducted with this fuel composition.

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

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

Combustion of Date Stone and Jojoba Solid Waste in a Hybrid Rocket-like Combustion Chamber

The performance of two solid biomass wastes, date stone and jojoba solid waste, was experimentally examined for their potential application in combustion and propulsion systems. The fuels were tested in a hybrid rocket-like combustion environment, and the test result was analyzed with combustion and propulsion parameters. The performance of both fuels was comparatively evaluated and compared with a conventional hydrocarbon fuel in a hybrid rocket, with paraffin wax serving as a baseline. A compression device was introduced to compress the solid biomass wastes into a circular-shaped fuel grain compatible with a hybrid rocket combustion chamber with a hot surface ignitor. Thermogravimetric analysis (TGA) and chemical equilibrium analysis (CEA) results revealed that the performance of the biomass fuel can be comparable to conventionally used hydrocarbon paraffin-wax-based propellant within a certain range of oxidizer-to-fuel ratio, in terms of theoretical specific impulse performance. Through experimental performance tests, it was found that the compressed biomass fuel grains were successfully ignited and produced thrust. Both biomass fuels tested in a hybrid rocket combustion chamber are expected to pave the way for further developments in biomass fuels in the waste-to-energy field for their application in combustion and propulsion systems, potentially replacing fossil fuels with renewable resources.

Novel Solid Propellants Enabled Through In Situ Martian Perchlorates

With evidence for the native perchlorates existing in the Martian regolith, this paper examines the feasibility and performance of propellants formed from perchlorate salts reported to be present on Mars. Thermochemistry calculations indicate that the Martian perchlorate-based propellants provide less theoretical specific impulse than AP composite propellants but could still be viable propellants. Three propellants made from Martian perchlorates were manufactured and compared to a control propellant with AP as the oxidizer. Deflagration experiments were performed to obtain the burning rates as a function of pressure, with results comparable to AP baseline propellant. The propellant energy density was evaluated through bomb calorimetry. The propellant formulation with a similar perchlorate mixture to the distribution found in Martian soil was then subjected to thermal analysis, elemental analysis, and sensitivity testing to examine its combustion behavior and suitability for handling. Further characterization and development work would be needed to field these propellants, but initial conclusions indicate an in situ blend of calcium perchlorate and magnesium perchlorate could serve as a novel oxidizer for future safe, high-performing, and economical solid propellant rocket motors, offering an alternative to most current proposals for Martian ascent vehicle architectures.

A Solar Thermal Steam Propulsion System Using Disassociated Steam for Interplanetary Exploration

Sustainable space exploration will require using off-world resources for propellant generation. Using off-world-generated propellants significantly increases future missions’ range and payload capacity. Near Earth Objects (NEOs) contain a range of available resources, most notably water-ice and hydrated minerals. However, water-bearing regolith needs to be excavated and the water extracted. Water is a compelling choice for fuel as it is readily available in interplanetary space and easily stored. In this paper, we propose using solar concentrators, which can efficiently convert incident sunlight into heat without the need for moving parts. When water is heated up to 4000 K, a value consistent with high-performance refractive materials, it experiences significant disassociation into H2, O2, OH, H, and O components, providing a path for adding considerable additional chemical energy per degree of temperature increase, and producing theoretical specific impulse (Isp) values in the range of 643 s to 659 s.

Experimental study on the discharge of a xenon-assisted krypton Hall thruster

The theoretical specific impulse of Hall thrusters using a krypton propellant is much higher than that of those using xenon. However, the actual benefits of this improvement are limited by the difficulty in achieving krypton atom ionization, and its efficiency is much lower than that of xenon. In this study, the anode specific impulse and anode efficiency of Hall thrusters using krypton propellant were improved by adding a small quantity of xenon. A dual-channel supply of xenon and krypton was achieved by optimizing the discharge channel structure based on the design of the aft-loaded magnetic field. The effects of the xenon and krypton supply method and xenon fraction on the discharge characteristics of the Hall thruster were investigated by conducting discharge experiments using the mixed propellant. The results indicated that injecting a small quantity of xenon significantly improved the ionization of krypton, improved the ionization and acceleration efficiency of the thruster, and fully utilized the high specific impulse characteristics of krypton. Furthermore, the addition of a small quantity of xenon also significantly increased the upper limit of the discharge voltage for stable operation using krypton from 500 V. At the maximum tested xenon fraction, the maximum discharge voltage was loaded to 900 V. At a discharge voltage of 800 V, the optimal anode efficiency and anode specific impulse increased by 4.5 % and 534 s, respectively, compared with those for pure xenon propellant (with an anode efficiency of 54 % and anode specific impulse of 3034 s). The results of this study are of great significance for the design of high-efficiency, high-specific-impulse, and low-cost Hall thrusters.

Propulsion theoretical and experimental analysis of composite propellants motors

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

Effect of AlH3 on the energy performance and combustion agglomeration characteristics of solid propellants

Increasing the performance of an energetic material can be achieved by changing its composition or using additives, such as metal powders. Aluminum hydride (AlH3) exhibits attractive properties, such as high hydrogen/energy storage, relatively good stability, and low dehydrogenation temperature. AlH3 has excellent prospects for use as a component in solid propellants for increasing the specific impulse of rocket engines and effectively reducing the erosion of engine nozzles. In this work, the effect of AlH3 on the energy performance, ignition characteristics, and combustion characteristics of solid propellants, as well as the agglomeration characteristics of metal particles was investigated. It was found that with increasing AlH3 content, the theoretical mass heat of combustion and the theoretical volumetric heat of combustion of the propellants increased, and the combustion efficiency also gradually improved. When the content of AlH3 in the propellant increased from 0.00% to 17.00%, the temperature of the propellant combustion chamber was reduced from 3314.59 to 2957.49 K and the theoretical specific impulse increased from 2584.2 to 2641.7 N‧s/kg. The average molecular mass of the gas was reduced from 25.54 to 21.54 due to the large amount of hydrogen gas released by the combustion of AlH3. The maximum combustion temperature decreased and the burning rate increased with the gradual replacement of aluminum (Al) in the propellant with AlH3. The maximum combustion temperatures of the pure Al propellant and the pure AlH3 propellant were 2510.5 and 2232.5 K, respectively. The linear burning rates were 0.29 and 0.46 cm/s at 1.0 MPa, respectively. The agglomeration process of AlH3 included three main stages, namely accumulation, aggregation, and agglomeration. With the increase in the AlH3 content, the number of agglomerated particles in the propellant gas-phase flame region decreased continuously.

Energetic Characteristics Analysis of Propellants Containing ammonium salts of bis(3-dinitro-1H-1,2,4-triazolyl) borate

The energy characteristics of HTPB propellant and modified double-base propellants containing bis(3-nitro-1,2,4-triazolyl) ammonium borate are simulated and calculated by NASA’s computer program CEA in the standard state (Pc: P0 = 70:1). The results show that the theoretical specific impulse of the modified propellant is 2539.9 N·s/kg. When the mass fraction of ammonium borate is 80%, the maximum Isp is 2488 N·s/kg, which is 6.5% higher than that of the basic formula, so adding ammonium borate to the propellant is beneficial to improving the overall performance of the propellant. In addition, it is environmentally friendly since it does not contain halogen elements.

IMPACT SENSITIVITY OF NITROMETHANE-BASED GREEN-PROPELLANT PRECURSOR MIXTURES

Nitromethane is a promising green propellant candidate. However, one potential issue for industrial use may be its high sensitivity to mechanical impact. In this paper, we assess the effect of adding several phlegmatizing substances (or inhibitors as we call them), on the impact sensitivity and the potential propulsive performance of the resulting mixtures. The impact sensitivity of the nitromethane mixtures was characterized using a BAM Fallhammer impact tester. The theoretical rocket performance of the mixtures was calculated utilizing the NASA CEA program. This study resulted in the identification of several nitromethane-inhibitor mixtures, which lack the high impact sensitivity of pure nitromethane but still retain much of its propulsive performance. The most promising approach was the addition of 6 wt.% of dimethyl sulfoxide: the theoretical specific impulse was only decreased by 3%, while the impact sensitivity was significantly lowered (> 10 J vs. < 5 J, as in pure nitromethane).

Rotating Detonation Combustion for Advanced Liquid Propellant Space Engines

Rotating (also termed continuous spin) detonation technology is gaining interest in the global research and development community due to the potential for increased performance. Potential performance benefits, thrust chamber design, and thrust chamber cooling loads are analyzed for propellant applications using liquid oxygen or high-concentration hydrogen peroxide oxidizers with kerosene, hydrogen, and methane fuels. Performance results based on a lumped parameter treatment show that theoretical specific impulse gains of 3–14% are achievable with the highest benefit coming from hydrogen-fueled systems. Assessment of thrust chamber designs for notional space missions shows that both thrust chamber length and diameter benefits are achievable given the tiny annular chamber volume associated with the rotating detonation combustion. While the passing detonation front drastically increases local heat fluxes, global energy balances can be achieved if operating pressures are limited to be comparable to existing or prior space engines.

Thermolysis and sensitivities of solid propellants using characterized nano oxidizers involving energy performance evaluation

This investigation was devoted to exploring the thermal decomposition and sensitivity characteristics of solid propellants using nano-AN and nano-AP as oxidizers. The morphology, surface elements, and crystal structure of nano-AP and nano-AN were investigated by SEM, BET, EDS, and XRD. DSC and TG-MS were used to investigate the thermal decomposition mechanism of nano-AP and nano-AN solid propellants. The mechanical sensitivity of the different solid propellants containing nano-AN and nano-AP was tested and their energy performance was also evaluated. The results show that the microscopic morphology of nano-AN and nano-AP is network-like with the one-dimensional scale less than 100 nm, and the crystal phases of the two are consistent with the raw AN and raw AP, respectively. The thermal decomposition activation energy of AN/CMDB-4 propellant containing nano-AN is 717.46 kJ/mol, while that of AP/HTPB-6 propellant containing nano-AP is 422.33 kJ/mol. The main decomposition products of AN/CMDB-4 propellant are CO2, N2O, NO, CH2O, CO, N2, H2O, CH4 and H2. The products of CO2, N2O, NO, CH2O, CO, N2, H2O, CH4 and H2 are also generated during the decomposition of AP/HTPB-6 propellant. With the increase of nano-AN and nano-AP in propellants, the mechanical sensitivity of AN/CMDB solid propellant decreases, while that of AP/HTPB solid propellant increases. The theoretical standard specific impulse Isp of AN/CMDB and AP/HTPB propellants are 2439.5 N·s/kg and 2449.3 N·s/kg, respectively. The above conclusions show that nano-AP and nano-AN can improve the thermal decomposition performance and the safety of solid propellants, which are expected to be widely used in solid propellants.

Effects of aluminum and aluminum hydride additives on the performance of hybrid rocket motors based on 95% hydrogen peroxide

This paper aims to study experimentally and numerically the effects of aluminum (Al) and aluminum hydride (AlH3) additives on the performance of HRMs. A steady-state numerical model of HRM is established, which adopts the 95% hydrogen peroxide (HP) as the oxidizer. A firing test is completed for the certification of the numerical simulation model, the results illustrate that the regression rate of addition of 38% AlH3 and 20% Al increases by 55.6% compared with pure hydroxyl-terminated polybutadiene (HTPB). A good agreement exists between the simulation regression rate and the experimental data. The deviations of thrust and pressure are less than 3.3% and 3.7%, respectively. Specifically, the error of the regression rate is only 0.3%, indicating that the accuracy of adopting numerical simulation method to predict the characteristics of the HRM is acceptable. Several steady-state simulations based on the different content of AlH3 and Al are performed to investigate the effects on regression rate and combustion performance of the HRM. Due to the addition of H2, the thermal conductivity increases with the decrease of the mass fraction of AlH3, which leads to the increase of the regression rate. The regression rate of the fuel with 8% AlH3 and 50% Al increases by 35.5% than the fuel with 58% AlH3, demonstrating that the effect of Al on the regression rate enhancement is more than that of AlH3. The theoretical specific impulse increases with the mass fraction of AlH3. However, it is owing that the part of H2 does not participate in the reaction with O2 and the oxidizer to fuel ratio (O/F) is far from the optimal O/F, the specific impulse is lower than the theoretical specific impulse.

Reactivity of hypergolic hybrid solid fuel with industrial grade hydrogen peroxide

Hypergolic solid fuels and hydrogen peroxide oxidizers have attracted considerable attention for the development of safe hybrid rocket propulsion systems. In this study, a new approach was explored to develop a safe hypergolic hybrid propulsion system using industrial-grade hydrogen peroxide (∼70 wt%). Initially, the propulsion performance of different hypergolic solid fuels with an industrial-grade hydrogen peroxide was calculated using the NASA CEA code to find an appropriate propellant combination. The theoretical specific impulse of the hypergolic solid fuel was 27.8 % over hydrazine and 13.4 % over LMP-103S. Based on theoretical investigations, hypergolic solid fuels, named such as HSF-1 to HSF-10, were fabricated using molding and pressing methods. The ignition delay time was measured using the drop-test method. Interestingly, HSF-3 and HSF-4 exhibited very short ignition delay times of 4.92 and 8.75 ms, respectively, with industrial-grade hydrogen peroxide. In addition, the drop test results of hypergolic solid fuel with varying compression pressure in the pressing method confirmed that a pressure above the compressive strength of the binder shortens the ignition delay times deviation. Further, HSF was also ignited even at low concentration of IGHP (>40 wt%); opens the new area of research in combustion science. Overall, low-toxicity hypergolic solid fuels and industrial-grade hydrogen peroxide are promising alternatives to conventional toxic hypergolic propellants.

On the Use of Fluorine‐Containing Nano‐Aluminum Composite Particles to Tailor Composite Solid Rocket Propellants

AbstractThe burning rate of solid propellants is an important factor for optimizing rocket motors and improving performance. The enhanced burning rate can increase thrust and reduce a propulsion system‘s overall size and weight. In this study, a novel nano‐aluminum/THV composite additive was prepared and introduced into a solid ammonium perchlorate/polybutadiene composite solid rocket propellant to enhance its burning rate. The morphology of the composite particle additive and its effects on combustion were characterized. The use of small quantities (<15 wt.%) of the additive resulted in a burning rate enhancement of up to 2.1 times that of the conventional coarse aluminized propellant with a specific impulse loss of only 3 seconds, and as much as 4.7 times enhancement with a predicted loss of 22 seconds in theoretical specific impulse. Some of this loss may be recovered by the improved combustion efficiency in smaller rocket motors because the additive was shown to significantly reduce the aluminum agglomeration at the propellant burning surface and reduce the size of reaction products which may reduce two‐phase flow losses. The additive also provides wide burning rate tailorability, favorable for motor, grain, and thrust curve design. The burning rate enhancement mechanism is thought to be a physical cratering mechanism governed by the burning rate disparity between the binder/oxidizer system and the nano‐aluminum/fluoropolymer additive and not a chemical catalytic effect.

A novel green electrically controlled solid propellant with good electrical response and high energy performance

Electrically controlled solid propellants (ECSPs) are a new type of propellants with electrically controllable ignition and extinguishment properties. At present, ESCPs are mainly achieved using ammonium nitrate (AN) or hydroxylamine nitrate (HAN) as the oxidant, and thus usually show the disadvantage of low energy performance. In the present work, a novel green ECSP has been designed and prepared using the ammonium dinitramide (ADN) as the main oxidant. For comparison purpose, three propellants were formulated with AN (ECSP-1), ADN (ECSP-2) and the mixture of AN and ADN (ECSP-3) as the oxidant, respectively. The theoretical calculation suggests that the energy of ECSP-3 increases with the increase of the relative content of ADN and the theoretical specific impulse of ECSP-3 reaches 1841.13 N s/kg, 6.5% higher than that of ECSP-1. The combustion performance and circuit characteristics of the propellants were tested with an electric ignition device built in-house. The ignition energy threshold of ECSP-3 is measured to be 516 J, close to the that of ECSP-1. The thermal stability of ECSP-3 is between those of ECSP-1 and ECSP-2. Yet ECSP-3 exhibits the optimal electrical response and high energy performances simultaneously. The gaseous combustion product of ECSP-3 at 25–300 °C was analyzed by continuous infrared spectroscopy and mass spectrometry. The results show that the gaseous product is clean and no harmful gases are found. Finally, according to the experimental results, the combustion mechanism of the prepared propellant was deduced. Our work has provided a new strategy to develop new green high-energy ECSPs.

Room-temperature catalytically reactive ammonium dinitramide–H2O2 monopropellant for microsatellites

Room-temperature catalytically reactive ammonium dinitramide (ADN)-based green monopropellants are interesting for space applications, particularly for microthruster applications. In this study, a blend of ADN and 90-wt% hydrogen peroxide monopropellant was developed and its catalytic reactivity was investigated at different temperatures. An optimal ADN–hydrogen peroxide monopropellant (AHMP) was fabricated at a weight ratio of 40:60 wt:wt (ADN:90-wt% H2O2). The theoretical vacuum specific impulse of AHMP was 206 s, 20% higher than the specific impulse (172 s) of 90-wt% hydrogen peroxide. The thermal properties of AHMP were evaluated using thermogravimetric analysis–differential scanning calorimetry, which revealed that thermal behavior of AHMP containing ADN and 90-wt% hydrogen peroxide was similar to their pure compound. The storability of AHMP was estimated using Fourier-transform infrared spectroscopy and nuclear magnetic resonance spectroscopy. Notably, AHMP could be stored at low temperatures (2–3 °C) for 4 months and at room temperature (20–23 °C) for 1.4 months without chemical changes. Furthermore, drop test and microthruster test were performed to determine the reactivity of AHMP with a catalyst (Pt/La/Al2O3). The response time of the catalyst on the decomposition of AHMP was approximately 104.6, 24.8, and 7.0 ms at 23, 60, and 120 °C, respectively. This study on the AHMP green monopropellant paves the way for further studies on improved ADN–hydrogen-peroxide-based monopropellants for microthruster applications.

Hypergolic ignition of hydrogen peroxide with various solid fuels

An experimental investigation was conducted to determine the hypergolic ignition delay time and behavior of various solid fuels containing sodium borohydride with high-grade hydrogen peroxide. It is shown that the hypergolic reaction between sodium borohydride and hydrogen peroxide provides the heat necessary to cause the ignition and combustion of the fuels. Ignition delay times between 5.5 ms and 10 ms were measured for tests performed at ambient conditions. The hypergolic ignition has also been studied at elevated pressures, resulting in decreased ignition delay times for most cases. Ignition delay times as short as 2.93 ms were measured for tests at 10 bar. The theoretical specific impulse of the tested fuels was found to be comparable to other liquid hypergolic propellants as well as common hybrid rocket combinations. Additional tests were conducted using radial oxidizer injection over a hollow cylindrical surface to prove that the hypergolic ignition of the selected fuels is certain. Approximated ignition and combustion delay times averaged 23.1 ms. The results imply that in a hypergolic-hybrid rocket configuration, short ignition delay times can be achieved, thus eliminating the need for a separate ignition system.

Preparation of ammonium nitrate-based solid composite propellants supplemented with polyurethane/nitrocellulose blends binder and their thermal decomposition behavior

To improve the performance of solid composite propellants (SCPs) supplemented with ammonium nitrate (AN) as an oxidizer, the incorporation of energetic ingredients such as explosives, energetic binders or catalysts is a common effective approach. For this purpose, polyurethane (PU), a typical inert binder, was mixed with nitrocellulose (NC) as an energetic polymer. Numerous composite solid propellant compositions based on AN and NC-modified polyurethane binder with different NC ratios were prepared. The prepared formulations were characterized using Fourier transform infrared spectroscopy (FTIR), RAMAN spectroscopy, X-ray diffraction (XRD), electron densimetry, thermogravimetric (TG) analysis, and differential scanning calorimetry (DSC). A kinetic study was then performed using the iterative Kissinger-Akahira-Sunose (It-KAS), Flynn-Wall-Ozawa (It-FWO), and non-linear Vyazovkin integral with compensation effect (VYA/CE) methods. The theoretical performances, such as theoretical specific impulse, adiabatic flame temperature, and ideal exhaust gaseous species, were also determined using the NASA Lewis Code, Chemical Equilibrium with Application (CEA). Spectroscopic examinations revealed the existence of NC and full polymerization of PU in the prepared propellants. According to density tests, the density of the propellant increases as the nitrocellulose component increases. According to the thermal analysis and kinetics study, the increase in NC content catalyzed the thermal decomposition of the AN-based composite solid propellants. Based on the theoretical study, increasing the amount of NC in the propellant increased the specific impulse and, as a result, the overall performance.