Space Propulsion Research Articles

Three-Dimensional Rapid Orbit Transfer of Diffractive Sail with a Littrow Transmission Grating-Propelled Spacecraft

A diffractive solar sail is an elegant concept for a propellantless spacecraft propulsion system that uses a large, thin, lightweight surface covered with a metamaterial film to convert solar radiation pressure into a net propulsive acceleration. The latter can be used to perform a typical orbit transfer both in a heliocentric and in a planetocentric mission scenario. In this sense, the diffractive sail, proposed by Swartzlander a few years ago, can be considered a sort of evolution of the more conventional reflective solar sail, which generally uses a metallized film to reflect the incident photons, studied in the scientific literature starting from the pioneering works of Tsander and Tsiolkovsky in the first decades of the last century. In the context of a diffractive sail, the use of a metamaterial film with a Littrow transmission grating allows for the propulsive acceleration magnitude to be reduced to zero (and then, the spacecraft to be inserted in a coasting arc during the transfer) without resorting to a sail attitude that is almost edgewise to the Sun, as in the case of a classical reflective solar sail. The aim of this work is to study the optimal (i.e., the rapid) transfer performance of a spacecraft propelled by a diffractive sail with a Littrow transmission grating (DSLT) in a three-dimensional heliocentric mission scenario, in which the space vehicle transfers between two assigned Keplerian orbits. Accordingly, this paper extends and generalizes the results recently obtained by the author in the context of a simplified, two-dimensional, heliocentric mission scenario. In particular, this work illustrates an analytical model of the thrust vector that can be used to study the performance of a DSLT-based spacecraft in a three-dimensional optimization context. The simplified thrust model is employed to simulate the rapid transfer in a set of heliocentric mission scenarios as a typical interplanetary transfer toward a terrestrial planet and a rendezvous with a periodic comet.

Experimental study on droplet dynamics behavior and combustion characteristics of high performance green propellant in electrical ignition mode

High performance green propellant represented by ammonium dinitramide-based liquid propellant and its new ignition method are the research hotspots of space propulsion in the 21st century. Exploring the complex multi-scale physical properties of multi-component ammonium dinitramide-based liquid propellant droplets in the electrical ignition mode has wide application significance for spray, propulsion system design and combustion control. The droplet dynamics behavior and combustion characteristics of propellant droplets at different ignition voltages were studied experimentally. The droplet dynamics behavior during the evaporation process, including violent volume oscillation, approximate steady-state expansion, contraction, secondary expansion, puffing and micro-explosion, have been determined by the generation, growth, and discharge of vapor bubbles. In the initial evaporation process, the heterogeneous nucleation is dominant. As the droplet is continuously heated, homogenization nucleation gradually dominates. The main physical and chemical mechanisms of bubble evolution driven by temperature involve methanol boiling, water overheating, ammonium dinitramide decomposition and combustion reaction between vapor molecules. Increasing the ignition voltage increases the droplet dynamics behavior and the combustion, but promotes the combustion instability. Increasing the ignition voltage increases the ignition delay time, puffing delay time, droplet lifetime, maximum temperature of droplet, and reduces the ignition critical diameter. It is proposed that the method of suppressing the droplet breakup dynamics at decomposition area and enhancing the droplet breakup dynamics at the combustion area are conducive to the combustion control of the thruster in electrical ignition mode. This research provides novel insight into the study of the electrical ignition mechanism of liquid fuels.

Elucidation of spacecraft surface wear phenomena in the plasma environment emitted by electric propulsion vehicles based on spacecraft charging analysis

Professor Takanobu Muranaka, Chukyu University, Japan, is part of a collaborative research effort to investigate and elucidate surface wear phenomena in electric propulsion spacecraft vehicles. He and his team are exploring how injected plasma emitted by these vehicles interacts with the spacecraft, leading to potential electrical and mechanical issues. For example, the ions and electrons in the plasma can collide with the spacecraft, causing wear and tear. This can compromise the spacecraftâ–™s durability and lifespan. In order to analyse these interactions, Muranaka and the team are utilising numerical calculations. To overcome limitations associated with the analysis of spacecraft systems and plasma in space, the researchers perform analyses of objects under predetermined conditions without consideration of space or timescales. Although it isnâ–™t possible to fully replicate space conditions in ground-based vacuum chambers, in their analyses the researchers are able to obtain important raw measurement data on physical phenomena that have not been modelled, such as parameter measurements of electrically propelled plasma plumes. Through its research, the team hopes to contribute to spacecraft design by informing design parameters such as the spacecraft shape and thruster parameters. Ultimately, the researchers are expanding and verifying numerical models and developing advanced plasma interference analysis technologies that will prove invaluable to spacecraft design. The researchers’ investigations on the Hayabusa2 spacecraft’s side-cathode ion thruster system have yielded exciting results on the ion energy distribution function (IEDF) of backflowing ions at multiple points around this thruster, expanding their findings on the significance of thruster reference potential on ion energy distribution, which is key to understanding spacecraft surface erosion.

Selection of design parameters of a spacecraft electric propulsion engine ensuring the geostationary orbit raising mission

The problem of selecting suboptimal parameters of an electric propulsion system as part of a spacecraft, ensuring a geostationary orbit raising mission using a Soyuz-5 launch vehicle at the initial launch stage, is investigated. An expression is obtained for the main optimality criterion – the payload relative mass. An algorithm is obtained for solving the problem of parametric optimization of an electric propulsion system as part of a hybrid transportation system. The task of selecting the optimal electric propulsion engine for launching a spacecraft with an electric propulsion system to a geostationary orbit using the launch vehicle under consideration is solved. A design ballistic calculation and a comparative analysis of the obtained data are performed.

Extreme electromagnetic rotation, chemical reaction, Hall and ion slip effects on weakly ionized fluid in a Riga channel

The utilization of external magnetic or electric fields, particularly through a Riga setup, markedly enhances flow dynamics by mitigating frictional forces and turbulent fluctuations, thereby facilitating superior flow management. Such improvements are especially beneficial in optimizing the operational efficiency of machinery and turbines. Our research focuses on the behaviour of a weakly ionized fluid within a porous, infinitely extended Riga channel (or electromagnetic channel) set in a rotational frame work affected by Hall and ion-slip electric fields. This model integrates the cumulative repulsions of an abruptly apply pressure gradients, electro-magnetic force, electromagnetic radiation, as well as chemical reactions. The physical configuration of the model features a stationary RHS wall as well as the LHS wall subjected to transversal vibration, establishing a complex flow environment. This circumstance is analytically modeled utilizing time dependent PDE, with the Laplace transform method applied to achieve a closed-format resolution for the course controlling equations. Through detailed graphical and tabular data, the investigations explored the impacts of assorted pivotal parameters on the model's flow traits and quantities. Our results indicate that an upswing in the modified Hartmann number significantly enhances fluids flow in the conduit, with the most important resultant flow showing marked improvement as Hall and ion-slip parameters amplify, and secondly the resultant flow diminishing chemical reaction parameter. Additionally, species concentration lowers with higher Schmidt numbers and chemical reaction rates, while an expanded modified Hartmann number correlate with enhanced shear stress near the conduit wall. Moreover, an elevation in the radiating parameter reduces the rate of temperature transport at the vibrating wall, whereas this at the stationary wall improves. This study has profound implications across several fields, notably in fusion energy research, spacecraft propulsion systems, satellite operations, aerospace engineering, and advanced manufacturing technologies.

Framework for Data‒Driven Fault Diagnosis of Numerical Spacecraft Propulsion Systems

The increasing complexity of deep space missions introduces significant challenges in maintaining spacecraft health, particularly in the propulsion systems, due to the inherent communication delays with Earth. This research proposes a novel framework for the autonomous, data-driven fault diagnosis of spacecraft propulsion systems. Leveraging data generated from a spacecraft propulsion system simulation model. The study addresses the limitations imposed by computational resources and sensor installation constraints through Sequential Forward Selection (SFS) for optimized sensor placement and feature selection. The framework's effectiveness is demonstrated through implementation on a microcomputer, showing promising results in terms of diagnostic accuracy and processing speed, thus highlighting its potential for onboard spacecraft application. This study not only advances the autonomous capabilities of spacecraft in deep space but also contributes to the broader field of Prognostics and Health Management (PHM) by providing a scalable, efficient approach to fault diagnosis in critical spacecraft systems. The proposed framework demonstrates a promising approach to optimizing diagnostic tasks for spacecraft systems. However, the trade-offs observed necessitate a careful consideration of task-specific requirements and the potential need for adjustments to maintain a high level of accuracy alongside computational efficiency.

Influence of constant and alternating electric fields on the deformation and destruction of a liquid droplet

Examining the impact of electromagnetic field, which provides thrust in contemporary rocket engines, is turning into a significant issue. Its resolution is required to guarantee the safety of space flight by enhancing engine performance and lowering fuel consumption. Spacecraft propulsion is achieved by low-thrust rocket engines. The principle of operation for ion colloid engines is the electrostatic acceleration of charged droplets. A static electric field accelerates charged liquid droplets created by electrospraying them in a low-thrust colloid engine. A promising technology in many industries is the active control of droplet motion and deformation with electric field. An electromagnetic field impact on droplet deformation and destruction in a viscous liquid is examined. The effect of electromagnetic field on individual droplets and emulsions is examined in terms of their physical mechanisms. Constant and alternating electric field effects on liquid droplet are represented numerically. The effects of droplet electric capillary number and viscosity ratio on droplet unsteady deformations is explored. The parameters that correspond to the droplet destruction are found. The results obtained have potential applications in enhancing the efficiency of current industrial electric dehydrators and in advancing the development of new electromagnetic field-based demulsification technologies.

Numerical Studies of Solid Rocket Propellant PMMA-PBAN-ALF3-Nitrocellulose-Difluoramine

Abstract The aim of the new investigation is the evaluation of solid rocket Propellant. In which performing numerical analysis of solid rocket propellant and calculating the constraints involved as derived from thrust i.e. velocity involved in the solid rocket propulsion motor usage. This paper aims to show the viability of solid rocket propellant for modern and future applications. Solid rocket motors are simple in design and fabricating. The need for a propellant that produces the same specific impulse as another chemical rocket engine. So it is needed to have a solid rocket motor with an enhanced combustion characteristic at a reasonable O/F ratio. For a healthy ecology, an economically reliable, highly effective, and environmentally concerned propellant is constantly suitable. As a superior alternative among the already used propellants in the sector, the paper suggested fuel PolyMethyl MethAcrylate (PMMA), PolyButadiene AcryloNitrile (PBAN) CoPolymer, and Aluminum hydride (AlH3), and oxidizers such as Nitrocellulose and Difluoramine. The desired O/F ratio for the technical criteria was found to result in a high combustion characteristic. With the features discovered through numerical and analytical research, we have suggested a design for the SRM that includes post-combustion, which enhances the reaction and creates the innovative elements of the existing SRM.

Operating Characteristics of a Wave-Driven Plasma Thruster for Cutting-Edge Low Earth Orbit Constellations

This paper outlines the development phases of a wave-driven Helicon Plasma Thruster for cutting-edge Low Earth Orbit (LEO) constellations. The two-stage ambipolar electric propulsion (EP) system combines the efficient ionization of an ultra-compact helicon reactor with plasma acceleration based on an ambipolar electric field provided by a magnetic nozzle. This paper reveals maturation challenges associated with an emerging EP system in the hundreds-watt class, followed by outlook strategies. A 3 cm diameter helicon reactor was operated using argon gas under a time-modulated RF power envelope ranging from 250 W to 500 W with a fixed magnetic field strength of 400 G. Magnetically enhanced inductively coupled plasma reactor characteristics based on half-wavelength right helical and Nagoya Type III antennas under capacitive (E-mode), inductive (W-mode), and wave coupling (W-mode) were systematically investigated based on Optical Emission Spectroscopy. The operation characteristics of a wave-heated reactor based on helicon configuration were investigated as a function of different operating parameters. This work demonstrates the ability of two-stage HPT using a compact helicon reactor and a cusped magnetic field to outperform today’s LEO spacecraft propulsion.

Symmetric engineered central cross-shaped broadband metamaterial absorber with high absorption and stability for solar sailing and solar energy applications

We propose a theoretical design and analysis of a broadband metamaterial absorber (MMA) with significant potential for solar sailing applications which is a method of spacecraft propulsion that usages the momentum of sunlight to propel a spacecraft through space. The absorber features a metal-dielectric-metal configuration with a tungsten (W) based resonator and ground plane, and a Silicon-dioxide (SiO₂) substrate. Addressing the critical need for materials that can efficiently harness solar radiation for propulsion in space, our design achieves an average absorption of 99.15 % over a broad spectrum from 250 nm to 1200 nm, covering the UV–Visible–NIR regions, with near-unity absorption peaks at 362 nm and 915.8 nm. It maintains high absorptions of 84.9 % and 86 % under transvers electric and transvers magnetic modes respectively, demonstrating excellent wide incident angle stability and polarization insensitivity due to its symmetric design. PCR values close to zero confirm its functionality as an absorber rather than a polarizer. The MMA shows minimal deformation across temperatures from 500 K to 1750 K and remains stable under various mechanical stresses, proving its durability and efficiency in space. Additionally, in solar thermophotovoltaic (STPV) systems, the MMA demonstrates high photothermal conversion efficiency (PTCE) over a wide temperature range (500 °C to 1500 °C) and different concentration factors. This dual functionality highlights its potential for both efficient space exploration and terrestrial solar energy harvesting, making it a versatile tool for future technological applications.

A Review of Recent Developments in Hybrid Rocket Propulsion and Its Applications

This paper extensively reviews hybrid rocket propulsion-related activities from combustion engine designs to launch tests. Starting with a brief review of rocket propulsion development history, a comparison among the three bi-propellant rocket propulsion approaches, and hybrid rocket engine design guidelines, a very thorough review related to hybrid rocket propulsion and its applications is presented in this paper. In addition to propellant choice, engine design also affects the hybrid rocket performance and, therefore, a variety of engine designs, considering, e.g., fuel geometry, swirl injection, ignition designs, and some innovative flow-channel designs are also explored. Furthermore, many fundamental studies on increasing hybrid rocket engine performances, such as regression rate enhancement, mixing enhancement, and combustion optimization, are also reviewed. Many problems that will be encountered for practical applications are also reviewed and discussed, including the O/F ratio shift, low-frequency instability, and scale-up methods. For hybrid rocket engine applications in the future, advanced capabilities and lightweight design of the hybrid rocket engine, such as throttling capability, thrust vectoring control concept, insulation materials, 3D-printing manufacturing technologies, and flight demonstrations, are also included. Finally, some active hybrid rocket research teams and their plans for flight activities are briefly introduced.

Catalytic effects of transition metal oxides on HTPB-based fuel polymer matrices

Low regression rate due to difficult pyrolysis is a major challenge in the practical application of terminated hydroxyl‑terminated polybutadiene (HTPB)-based fuels in hybrid rocket propulsion. Accelerating the decomposition of the polymer matrix is an effective method to improve the regression rate of HTPB fuels. To determine the difference in combustion performance between different transition metal oxides loaded HTPB-based fuels, nickel oxide (NiO), ferric oxide (Fe2O3), copper oxide (CuO) and manganese dioxide (MnO2) were introduced into the HTPB matrix at 5% mass fraction. The experimental results showed that the metal oxides could significantly catalyze the pyrolysis of HTPB-based fuels and enhance the fuel regression rate, and the catalytic effect was mainly concentrated in the middle and late stages of the thermal decomposition process of polybutadiene components. Among the four transition metal oxides, CuO and MnO2 showed better catalytic effects on the combustion performance of HTPB-based fuels in the high oxygen mass flux region, while NiO showed better catalytic effects in the low oxygen mass flux region. The present study compares the regression rate of fuel grains modified with different transition metal oxides, which provides a basis for the selection of future catalysts, verifies the catalytic effect of the transition metal oxides on the combustion of HTPB-based fuels and further analyzes the combustion reaction mechanism of the fuels.

EVALUATION OF SOLID ROCKET PROPELLANTS FOR LOW EARTH ORBIT

In contemporary aerospace technology, solid rocket propellants are commonly employed in the thriller and boost of space cargoes to LEO. These are reliable, uncomplicated, and high in thrust; and thus, can be used in both military as well as civil space ventures. This paper provides a comprehensive evaluation of five prominent solid rocket propellants: Five categories namely Ammonium Perchlorate Composite Propellant (APCP), Double-Base Propellant, Composite Modified Double Base (CMDB) Propellant, Hydroxyl-terminated polybutadiene (HTPB)-Based Propellant, and Ammonium Nitrate Composite Propellant (ANCP) have been identified. This also falls under the evaluation and includes aspects such as historical bases to see how the formulation of the propellant has changed over time regarding the development of propellant technology. In the case of each propellant, the chemical properties of the mix and reactions that its fundamental are analyzed. It also shares details regarding the changes and advancements made in the little while to enhance capability, safety, and environmental compatibility of these fuels, such as synthesizing fuels for liquid bipropellants. Some of the modern applications of technologies within the active spectrum of aerospace operations are explained in more detail to provide clients with the most versatile and essential information. Among the essentials of this work, the complex calculation of thrusts developed by the specific propellant with corresponding graphics has significant importance. These estimations afford chances of comparing the efficiency of the propellant in terms of producing thrusts as well as achieving the indicated optimal performance. Thus, the main objective of this study is to identify which of the solid rocket propellants to use in LEO missions and if any, what modifications could be made to improve them. Thus, the purpose of this paper is to contribute to the development of the constantly evolving branch of rocket propulsion by discussing the benefits and shortcomings of each propellant. These findings and recommendations could be useful in perfecting the recipes for the propellants and improving the efficiency of the subsequent missions to planets. Keywords: Aerospace, Solid rocket propellants, Thrust, Low earth orbit.

Experimental study and performance analysis of a dual-mode space propulsion system pressurized by electric pump

This study presents a new dual-mode propulsion system that utilizes an electrically pumped hydrogen peroxide and kerosene supply system, specifically designed for advanced spacecraft orbit and attitude control. The system incorporates two distinct electric pumps, one for hydrogen peroxide and another for kerosene, each engineered to optimize the propellant feed for the orbit control and attitude control engines. This configuration enables precise propellant management and thrust control, which are critical for the stringent demands of modern space missions. The dual-mode capability of the propulsion system allows for the use of hydrogen peroxide both as a monopropellant in the attitude control engine and as an oxidizer in combination with kerosene in the orbit control engine. This approach simplifies the propulsion architecture while maximizing the efficiency and utility of hydrogen peroxide as a propellant. The electric pumps are essential in providing high-pressure propellant delivery, which is necessary to achieve the desired combustion characteristics and thrust parameters. Performance evaluations of the propulsion system demonstrated significant results. The orbit control engine achieved a combustion chamber pressure of 5.9 MPa and a combustion efficiency of 98.42 %. The attitude control engine, operating at a peak pressure of 4.9 MPa, exhibited a stable pulsative mode, maintaining consistent pressure levels throughout its operational cycle. These performance metrics highlight the effectiveness of the electric pump technology in enhancing propulsion system responsiveness and efficiency. This propulsion system is expected to improve spacecraft maneuverability and reduce operational costs, making it suitable for long-duration space missions.

Data-driven sparse modeling of oscillations in plasma space propulsion

Data-driven sparse modeling of oscillations in plasma space propulsion

Oxygen production by solar vapor-phase pyrolysis of lunar regolith simulant

The oxide-rich lunar surface regolith can be used to extract the oxygen needed for the future of lunar exploration efforts as a consumable for life-support systems and spacecraft propulsion. Various techniques for the extraction of oxygen have been developed already, with solar vapor-phase pyrolysis shown to be a promising yet understudied approach. In contrast to other techniques, it requires only locally available resources, such as unbeneficiated regolith, sunlight, and vacuum in order to liberate oxygen and oxygen-bearing molecules. This study presents experimental work conducted in a purpose-built solar-vacuum furnace showing the evaporation of sodium and iron from a regolith simulant sample and their deposition on the crucible surface. This is matched by the thermochemical equilibrium modeling done in FactSage, which analyzes the process at varying pressures down to ultra-high vacuum. It highlights the need for precise temperature and pressure control, as well as the impact of regolith composition on oxygen dissociation for an efficient extraction of molecular oxygen.

Optimizing heat transfer with electro-osmotic ciliated propulsion in a non-uniform canals with Cu-blood characteristics considering thermal casson nano fluid

Electroosmosis peristaltic flows find promising applications in electro-fluid thrusters in space propulsion, polymer injection systems, ion exchange membrane designs, microbe fuel cells, biochip fabrication and fossil fuel energy process. In light of this, the current work aims to investigate blood based electroosmotic peristaltic flow in a ciliated, non-uniform propagation channel when copper nanoparticles are present. For that purpose, mathematical modeling is done in two dimensional frame and then governing partial differential equations are simplified and converted in ordinary differential equations using dimensionless quantities and long wave length and low Reynolds number approximations. In result we got coupled nonlinear boundary value problem that is solved numerically using three-stage Lobatto IIIa formula known as bvp4c invoking shooting technique. Peristaltic ciliated waves are thought to pass along the channel wall. The properties of both fluid phase and particle phase flow are modeled and investigated. Plotting against various parameters, streamline contours, temperature contours, concentration and temperature profiles, and pressure rise graphs are examined in-depth from a physical perspective. The obtained results revealed that concentration profile α upsurges for rising values of Casson fluid parameter β, nanoparticle volume fraction ϕ and flow rate Q while it drops with an increase in particulate volumetric fraction C, Eckert number Ec, Soret number Sr, Schmidt number Sc and viscosity constant. A decrease in temperature is also caused by an increase in the volumetric percentage of nanoparticles. In the pumping portion, there is a decrease in pressure rise; however, this decrease changes to an increase in the increased pumping zone.

Performance Characterization of a Radial Rotating Detonation Combustor with Axial Exhaust

Abstract This experimental study characterizes the performance of a radial rotating detonation combustor (RDC). An aerospike nozzle for rocket propulsion has been integrated into the center of the combustor, although the same combustor could also be coupled with turbomachinery. The radial RDC utilized a rapid to gradual (RTG) area change in the flow direction to effectively confine the detonation region close to the inlet plane and to improve the uniformity of the flow exiting the RDC. Three test cases were analyzed, (a) a baseline case at a total reactant mass flow rate, m˙ = 0.136 kg/s and equivalence ratio, ϕ = 0.6, (b) a higher reactant flow rate, m˙ = 0.318 kg/s and ϕ = 0.6, and (c) a higher ϕ = 0.8 at m˙ = 0.318 kg/s. All tests were conducted using methane and a 67% oxygen and 33% nitrogen (by mole) oxidizer mixture. Measurements were acquired using CTAP probes inside the combustion channel and along the aerospike to characterize the performance, PCB and ion probes near the detonation region to identify wave modes and their variations during the test, and thrust measurements using a six-axis force sensor. Results show highly complex wave modes with multiple co-rotating and/or counter-rotating waves depending upon the reactant flow rate. The pressure and thrust measurements are consistent with the wave mode analysis. In general, a positive (combustor only) pressure gain was inferred when losses associated with the injection system were excluded.

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.

Simulation and experimental analysis of aerogel's attenuation for high-energy alpha particles in fission-fusion fragment rocket applications

Emerging studies are geared toward exploring new methods of nuclear rocket propulsion to provide more efficient space transit beyond Earth's orbit. One method is to employ a Fission Fragment Rocket Engine utilizing fissionable layers embedded in a low-density aerogel. A quantitative understanding of particle attenuation is essential for developing a functional prototype that permits fission fragments to escape the layers and contribute to specific impulse rather than being attenuated and generating waste heat. In this investigation, the MCNP code was used to theoretically analyze the attenuation of alpha particles from 241Am sources within aerogel materials. Simulations were conducted on aerogels with various densities and compositions. These simulations aimed to predict the expected intensity of alpha particles reaching a detector. CR-39 was employed as a Plastic Nuclear Track Detector to assess particle attenuation by the aerogels. The experimental and simulation results show that the threshold areal density of atoms was found to be high 1020 atoms/cm2 for the three materials studied in this project.