Launch Vehicle Research Articles

Development of a Deployable Reflector Antenna for the Synthetic Aperture Radar Satellite, Part 2: Manufacturing and Qualification of the Main Reflector Using a Honeycomb Sandwich Composite Structure

A deployable reflector antenna (DR-A) is a structure that can be stored in a large-diameter Synthetic Aperture Radar (SAR) antenna and be mounted onto a launch vehicle. Considering the performance of the launch vehicle, it is necessary to develop a lightweight, high-performance antenna structure. The solid-type deployable reflector antenna is composed of a number of unit main reflectors. To reduce the weight of the antenna, a lightweight main reflector must be developed. In this paper, following “Development of Deployable Reflector Antenna for the SAR Satellite (Part 1)”, the manufacturing and qualification of the main reflector using honeycomb sandwich composites are described. Four types of composite main reflectors were manufactured with variables in the manufacturing process. The manufacturing variables include the curing process of the structure, the application of an adhesive film between the sheet and the core, and the venting path inside of the sandwich core. After manufacturing the main reflector, we performed weight measurements, non-destructive testing (NDT), surface error measurement using a Coordinate Measurement Machine (CMM), and modal testing for each type of composite main reflector. Through the research and development process, we found that a perforated hole is necessary when excluding the adhesive film during bonding of an aramid core and a CFRP sheet, and a lightweight composite reflector could be developed through this process. We selected the main reflector with the best performance and developed a composite main reflector that can be applied to satellites.

Multiphysics Analysis of Convergent-Divergent nozzle using FGM

Abstract: The study focuses on the Convergent-Divergent nozzle for the different types of flow regions and Mach numbers such as subsonic region, sonic region and supersonic region. The Mach number of the nozzle changes across the nozzle, the Mach number varies from the subsonic region at the inlet and varies to the supersonic region at the exhaust of the nozzle. The fluid domain is made across the nozzle to capture the effects of the pressure and temperature and Multi physics analysis is also done to find the deformation depending on the material. The multi-physics analysis of the nozzle is done in COMSOL Multiphysics software by using the system coupling of the turbulent model, solid mechanics and heat transfer models. The results show the deformation of the nozzle under the action of chamber pressure and chamber temperature as the nozzle is the crucial part of any rocket or jet engine of an aircraft or rocket or launch vehicle. The Convergent-Divergent nozzle first converges the flow from the combustion chamber and then diverges the flow to get the maximum thrust at the convergent section the pressure of the fuel increases and then at the exhaust or convergent section suddenly increases enhances the thrust and velocity of the coming the exhaust fuel.

Predicting Wall Pressure Fluctuations on Aerospace Launchers Through Machine Learning Approaches

Artificial intelligence (AI) can be used to optimize the prediction of pressure fluctuations over the external surfaces of aerospace launchers and minimize the number of wind tunnel tests. In the present research, various machine learning (ML) techniques capable of predicting the acoustic load were tested and validated. The methods included decision trees, Gaussian Process Regression (GPR), Support Vector Machines (SVMs), artificial neural networks (ANNs), linear regression, and ensemble methods such as bagged and boosted trees. These algorithms were trained using experimental data from an extensive wind tunnel test campaign conducted to support the design of a VEGA (Advanced Generation European Vehicle) launcher vehicle and provide wall pressure fluctuations in many configurations. The main objective of this study was to identify, among several algorithms, the most suitable method able to process such complex databases efficiently and to provide reliable predictions. Different statistical indices, including the root mean square error (RMSE), the mean square error (MSE), and a correlation coefficient (R-squared), were employed to evaluate the performance of the ML methods. Among all the methods, the bagged tree algorithm outperformed the others, providing the most accurate predictions, with low RMSE and high R-squared values across all test cases. Other methods, such as the ANNs and GPR, exhibited higher errors, indicating their reduced suitability for this dataset. The results demonstrate that ensemble decision tree methods are highly effective in predicting acoustic loads, offering reliable predictions, even for configurations outside the training database. These findings support the application of ML-based models to optimize experimental campaigns and enhance the design of aerospace launch vehicles.

Propulsive landing of launchers’ first stages with Deep Reinforcement Learning

The planetary landing problem is gaining relevance in the space sector, spanning a wide range of applications from unmanned probes landing on other planetary bodies to reusable first and second stages of launcher vehicles. In the existing methodology there is a lack of flexibility in handling complex non-linear dynamics, in particular in the case of non-convexifiable constraints. It is therefore crucial to assess the performance of novel techniques and their advantages and disadvantages. The purpose of this work is the development of an integrated 6-DOF guidance and control approach based on reinforcement learning of deep neural network policies for fuel-optimal planetary landing control, specifically with application to a launcher first-stage terminal landing, and the assessment of its performance and robustness. 3-DOF and 6-DOF simulators are developed and encapsulated in MDP-like (Markov Decision Process) industry-standard compatible environments. Particular care is given in thoroughly shaping reward functions capable of achieving the landing both successfully and in a fuel-optimal manner. A cloud pipeline for effective training of an agent using a PPO reinforcement learning algorithm to successfully achieve the landing goal is developed.

A multizonal numerical combustion model of ammonium perchlorate

Solid rocket motors (SRM) are essential for national defense, satellite launch vehicles for placing spacecraft for communications, resource management, and space exploration. Numerical modeling of composite solid propellants is very helpful for optimizing performance, ensuring safety, and complementing experimental testing. It provides insights into combustion dynamics and allows for precise customization to meet specific mission needs, so modeling compositepropellants will stay important. AP (Ammonium Perchlorate), as a synthetic oxidizer, has been widely utilized in modern composite solid propellants. Therefore, it is crucial to understand the physicochemical processes such as condensed-phase heating and reaction kinetics, the interactions between the condensed and gas phases, and gas-phase combustion. A steady-state numerical simulation model is presented to study the combustion of AP. Zonal modeling is employed to treat the solid phase, melt layer, and gas phase separately with conservation of mass, energy, and species, and the solutions are coupled with appropriate boundary conditions. A simple global reaction is developed, validated, and used for the condensed phase with better surface species profiles than those available in the literature. A detailed reaction mechanism is used in the gas phase combustion. This model considers only liquid as a condensed phase and uses a newly condensed phase mechanism and a premixed AP/HTPB (hydroxyl-terminated polybutadiene) gas phase reaction mechanism instead of AP monopropellant gas phase mechanism. This modeling is a prerequisite for a more sophisticated multi-modal composite propellant model with AP grains and AP/HTPB binder. The predicted burn rate and initial temperature sensitivities for different motor operating pressures match well with experimental and other theoretical data. Also, the simulated melt layer thickness of the present model agrees well with experimental observations. Sensitivity analysis is performed for the melt temperature and activation energy for the condensed phase reaction. The simulation also predicts surface temperature and species profile with reasonable accuracy.

Coupling of Advanced Guidance and Robust Control for the Descent and Precise Landing of Reusable Launchers

This paper investigates the coupling of successive convex optimization guidance with robust structured H∞ control for the descent and precise landing of Reusable Launch Vehicles (RLVs). More particularly, this Guidance and Control (G&C) system is foreseen to be integrated into a nonlinear six-degree-of-freedom RLV controlled dynamics simulator which covers the aerodynamic and powered descent phase until vertical landing of a first-stage rocket equipped with a thrust vector control system and steerable planar fins. A cost function strategy analysis is performed to find out the most efficient one to be implemented in closed-loop with the robust control system and the vehicle flight mechanics involved. In addition, the controller synthesis via structured H∞ is thoroughly described. The latter are built at different points of the descent trajectory using Proportional-Integral-Derivative (PID)-like structures with feedback on the attitude angles, rates, and lateral body velocities. The architecture is verified through linear analyses as well as nonlinear cases with the aforementioned simulator, and the G&C approach is validated by comparing the performance and robustness with a baseline system in nominal conditions as well as in the presence of perturbations. The overall results show that the proposed G&C system represents a relevant candidate for realistic descent flight and precise landing phase for reusable launchers.

Design and application of deployable heat radiator

Abstract Limited by the envelope size of the launch vehicle, comparing with the body-mounted radiator of the traditional spacecraft, the thermal control scheme based on single-phase fluid loop is proposed in order to improve the system-level heat dissipation. The thermal analysis model of the deployable heat radiator, including the spacecraft, is established. Through the comparative analysis of the fin parameter, infrared radiation effect, and working fluid selection, the heat dissipation efficiency of the radiator is improved, and the first flight application of the deployable heat radiator is realized. The research results show that in view of the GEO spacecraft’s bad external heat flow condition, the heat dissipation of a single radiator can reach 250W/m2 above; the method of using multiple deployable heat radiators can effectively solve the problem of insufficient heat dissipation capacity on the spacecraft.

Modular Flat Structure with Miura Origami for Space Solar Power Station

To address the challenges associated with existing space solar power station (SSPS) concepts, including noncompact structural design, nonuniform solar energy flow density, and orbital deployment complexities, an integrated, highly modular, flat functional structure based on the Miura origami pattern is proposed. The flat functional structure consists of a flat quadrilateral Fresnel concentrator for solar energy collection, a photovoltaic array for photoelectric conversion, and a transmitting array for microwave power transmission. To optimize space utilization and enhance the deployment ratio, the Miura origami pattern is introduced, and the subarray composed of multiple flat functional structures can be packaged in a payload capsule before launch and unfold as a standard rectangle with uniform thickness and a flat surface in orbit, improving the transportability of the launch vehicle. Additionally, to solve the nonuniform solar energy flow density, a Fresnel concentrator was employed to achieve uniform irradiation, while the transmitting array was designed to achieve good electrical performance. Finally, the structural parameters and efficiency of the SSPS system composed of multiple subarrays are presented in detail.

Effect of Free-Stream Mach Number on the Base Thermal Environment of Launch Vehicle

Effect of Free-Stream Mach Number on the Base Thermal Environment of Launch Vehicle

Studies on multiple weld repairs of M250 maraging steel for solid motor casing used in satellite launch vehicle application

Studies on multiple weld repairs of M250 maraging steel for solid motor casing used in satellite launch vehicle application

Fatigue limit prediction of 7050 aluminium alloy based on experimental and shallow + deep hybrid neural network

7050 aluminium alloy, as an important lightweight high-strength structural material because of its excellent characteristics, has increasing applications in aviation, aerospace, and launch vehicle manufacturing. Fatigue failure accounts for approximately 80 % of the structural failures in the engineering field, which leads to numerous losses and is highly uncertain and sudden. Therefore, fatigue life prediction of 7050 aluminium alloy has become a prominent research field. In this study, considering the two variables of stress ratio and stress concentration coefficient, fatigue tests of 7050 aluminium alloy under five different working conditions were conducted to evaluate the effect of these factors on the durability of 7050 aluminium alloy. A novel hybrid neural network model was proposed to solve the time-consuming problem of obtaining a large amount of S-N curve data through traditional fatigue tests. The model uses relatively low cycle fatigue test data for training, realizes data derivation, accurately predicts the fatigue limit of materials, and realizes the purpose of quickly evaluating the fatigue properties of 7050 aluminium alloy materials under different working conditions.

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.

The Use of Laser Sensing for Solving Meteorological Problems Related to Researching and Ensuring the Safety of Space Flights.

This paper is devoted to the issue of using laser (lidar) sensing to determine wind speed and direction when solving practical problems during the analysis of meteorological conditions in the area around spaceports. This issue is relevant both for making decisions on the possibility of a safe launch of a launch vehicle and for conducting search and rescue operations using groups of manned and unmanned aerial vehicles. Based on numerical and experimental modeling, it is shown that lidars provide highly accurate measurements of wind speed profiles and allow for the determination of weak wind shear in vertical and horizontal directions. This paper proposes a method for determining the main parameters of lidar sensing (range, resolution, detectability, etc.), which allows for the capabilities of this technology in solving the practical problems of meteorological monitoring to be predicted. Of particular interest in this article are experimental modeling data verifying the proposed calculation methods and the experimental determination of the capabilities of lidar diagnostics. This paper summarizes the data from multi-month experiments measuring wind speed in clear weather conditions when other means of remote diagnostics are ineffective. As a result of the experiments, a statistical distribution of the maximum range of wind speed measurement in normal weather conditions with natural variation in the concentration of scattering particles in the atmosphere was obtained. This article also discusses the possibility of combining lidars and meteorological radars for the meteorological support of flights.

Optimal Selection of Active Jet Parameters for a Ducted Tail Wing Aimed at Improving Aerodynamic Performance

The foldable tail of the box-type launch vehicle poses a risk of mechanical jamming during the launch process, which is not conducive to the smooth completion of the flight mission. The integrated nonfolding ducted tail proposed in this article can solve the problem of storing the tail in the launch box. Moreover, traditional mechanical control surfaces have been eliminated, and active jet control has been adopted to control the pitch direction of the flight attitude, which can improve the structural reliability of the tail wing. By studying the effects of parameters such as momentum coefficient, jet hole position, jet hole height, and jet angle on improving the aerodynamic performance of ducted tail wing, relatively good jet parameters are selected. Research has found that compared with jet hole height and jet angle, momentum coefficient and jet hole position are more effective in improving the aerodynamic performance of ducted tail wings. Under a trailing edge jet, a relatively good jet condition occurs when the jet hole height is equal to0.25% of the aerodynamic chord length, and the jet angle is equal to 0°. At this time, with the increase of the jet momentum coefficient, the effect of increasing the lift of the ducted tail wing is the best. Finally, a comparative analysis is conducted on the lift and drag characteristics between the ducted tail wing and traditional tail wing, and it is found that the ducted tail wing can generate lift at a 0° attack angle and will not stall in the high attack angle range of 12°~22°, with broad application prospects.

Bundle effect on a helical coil in liquid nitrogen with pool boiling for liquid oxygen densification

The need to densify oxidizers using cryogenic fluids is increasing to enhance the performance of launch vehicles. One of the most practical methods for oxidizer densification is heat exchange cooling between liquid oxygen and liquid nitrogen, typically using a tube bundle heat exchanger. Due to the multiple tubes in the bundle, a bundle effect arises, which enhances convective heat transfer by inducing liquid agitation from bubble generation and rising. This paper presents experimental results and prediction models that account for bubble behavior in a tube bundle. The experiment is conducted with saturated liquid nitrogen in a pool at atmospheric pressure and helical coil-type heat exchangers instead of a traditional tube bundle stack heat exchanger. Liquid oxygen densification is achieved by varying mass flow rates and inlet temperatures. Single-passage helical coils made of copper are used to minimize uncertainty from maldistribution flow and reduce thermal resistance compared to convective heat transfer coefficients in the inner and outer tubes. The coils, with an outer diameter of 12.7 mm, were tested in both vertical and horizontal directions and with various coil pitches. The bundle effect was clearly observed under helical coil conditions, and the experiment confirmed that the convective heat transfer coefficient increased with increasing heat flux and bubble generation rate. The prediction models considering bubble behavior—rising and generation rate—were validated by comparison with experimental results. The forced convective Nusselt number, experimentally measured to range from 23 to 361 through its correlation with the Boiling Reynolds number, closely followed the predicted correlation curve of the bubble generation model. It demonstrated a mean absolute error of 83.3, a standard deviation of 65.6, and an average relative error of 64.8 %. These values show improved accuracy compared to the relative errors of two predicted curves in the bubble rising model: 216 % for the single circular tube correlation and 381 % for tube bank correlations. This improvement suggests that the increased bubble generation rate with heat flux is better reflected for liquid oxygen densification with a helical coil submerged in large-scale static pool condition.

Innovation in sustainable composite research: Investigating graphene-reinforced MMCs for liquid hydrogen storage tanks in aerospace and space exploration

The demand for lightweight materials in space exploration applications has seen a significant rise. This study presents a novel approach by integrating graphene with Aluminium alloy (AA) 2900 powder, synthesized through High energy ball milling technique. The resulting powder was used to reinforce Aluminium alloy 2195, creating metal matrix composites (MMCs) via squeeze casting. The findings demonstrate that incorporating 0.5 wt % 2D graphene nanoplatelets (2D-Grnp) in AA 2195 achieved density match, enabling homogeneous dispersion, addressing composite casting challenges. Consequently, the AA 2900 alloy embedded with 2D-Grnp facilitated the homogeneous dispersion of reinforcements during the squeeze casting process, yielding MMCs Ideal for potential use in lightweight liquid hydrogen fuel tank structures for launch vehicles. Following T8 heat treatment, the final casted composite plate exhibited significant enhancements in ultimate tensile strength, with a 4.5% increase over monolithic Al 2195-T8, exhibiting nominal reduction in elongation behavior and higher hardness. Additionally, scanning and transmission electron microscopy (SEM, TEM), Small Angle Neutron Scattering (SANS) and Electron Backscatter Diffraction (EBSD) revealed the squeeze casting technique's effectiveness by exhibiting enhanced bonding among homogeneously reinforced 2D-Grnp, T1/θ’ precipitates, and AA 2195.

Asteroid disruption and deflection simulations for multi-modal planetary defense

Planetary defense from asteroids via deflective means alone does not offer viable solutions in terminal scenarios where there is little warning time before impact. The PI method of planetary defense enables operation in terminal interdiction modes where there is little warning time prior to impact, but can also operate in the same extended time scale interdiction modes as made possible by traditional deflection techniques, which results in a versatile, multi-modal planetary defense capability. The method is also practical and cost-effective since it relies solely on launch vehicles and penetrator materials already available today, and thus presents itself as a logical and competitive option for planetary defense. As per the PI method, we investigate the effectiveness of rubble pile asteroid disruption and deflection via hypervelocity impacts with 10:1 aspect ratio cylindrical tungsten penetrators. We present the results of an ongoing simulation campaign dedicated to investigating the PI method, using the Lawrence Livermore National Laboratory (LLNL) arbitrary Lagrangian–Eulerian (ALE) hydrodynamics code ALE3D run with the High-End Computing Capability (HECC) at NASA Ames Research Center. We model heterogeneous rubble pile asteroids with a distribution of spherical boulders of varying initial yield strengths set within a weak binder material. We find that rubble pile asteroids of this type in the 20–100 meter-class can be effectively mitigated via 20 km/s impacts with 100–1000 kg penetrators via the coupling of the penetrator kinetic energy into the bulk material of the asteroid.

Microstructure characterization and formability assessment of electron beam welded Nb-10Hf-1Ti (wt.%) refractory alloy sheets

Formability study on welded blanks of Nb-based refractory alloy C-103 (Nb-10Hf-1Ti (wt.%)) has tremendous potential for utilizing smaller sheets for realization of large-size divergent portions of upper-stage liquid engine nozzle of satellite launch vehicles and thrusters of attitude orbital control systems. In this work, vacuum-annealed monolithic C-103 sheets were electron beam welded successfully to obtain defect-free welds and detailed microstructural and mechanical characterizations were investigated. Columnar epitaxial grains were found in fusion zone (FZ) instead of equiaxed grains in base metal (BM). Detailed SEM and HRTEM analyses revealed spherical shape of nano-sized HfO2 precipitates with monoclinic crystal structure in the C-103 sheet. Further, microhardness across the weld cross-section indicated that the FZ had a uniform and maximum hardness of ∼195 HV0.5 due to formation of finer nano-sized HfO2 precipitates with higher number density, followed by a lowering of hardness in a narrow heat-affected zone. A marginal reduction in tensile strength of ∼6% with a considerable decrease in ductility of ∼29% was noticed for the welded sample. However, fracture occurred in the BM region, indicating good weld integrity. Furthermore, formability of the welded blanks was evaluated in terms of limiting dome height (LDH), forming limit diagram, and deep drawn cup depth. The effect of the presence of WZ on the formability of the C-103 sheet was estimated under uniaxial, plane strain, and biaxial tensile strain paths for the first time, and the results were compared with the monolithic sample. It was noted that the failure limit of the welded specimens decreased compared to that of the monolithic sample, resulting in a lower LDH of approximately 69% and 62% when deformed along plane strain and biaxial tensile strain paths, respectively. However, a marginal variation in LDH was observed for the uniaxial strain path. Also, the welded blank had more resistance to material flow into the die cavity due to the harder weld region, resulting in a reduction of deep drawn cup depth by 52% than that of the monolithic blank. The overall results concluded that the presence of the WZ significantly affected the formability of the C-103 sheet, especially in plane strain and biaxial strain paths. However, the fracture was never found to be propagating along the weld line, indicating the ability of the weld to sustain large plastic strains. This study provides insightful information on the formability of electron beam welded C-103 sheets for the successful fabrication of space components.

An application of aircraft noise metrics to the Artemis-I launch

In aerospace and acoustical research, there has been significant focus on understanding the negative effects of noise from jet aircraft and flyover vehicles. However, there has been relatively little investigation into the specific noise impacts of launch vehicles. Despite a considerable rise in the number of launches from various spaceports globally in the past decade, there appears to be poor understanding of the potential harm these increasing launches could pose to nearby communities, including both humans and wildlife. This paper aims to apply established noise metrics such as overall sound pressure level, sound exposure level, and effective perceived noise level, commonly used in aircraft policy, to quantify the potential noise impact of launch vehicles, using measurements from the Artemis-I mission as a case study.

Statistical reliability estimation of space launch vehicles: 2000–2022

This research examines the reliability of space launch vehicles (SLVs) performing commercial, civil, and military space lift missions through trend analysis and a variety of statistical methods. Data sets obtained from the Seradata database are analyzed for trends by examining data subsets including launch date, sector (commercial, civil or military), launch country, intended mission orbit, SLV family, and failed subsystem. Evolving, time-dependent first-level Bayesian success rate analysis is performed to assess SLV reliability and performance trends on 1873 launches occurring during the period 1 January 2000 to 31 December 2022. It was determined that even with a near-exponential increase in launch events in the 2020s, the number of failures is less than 6% each year, with the average success rate being 95% annually. Overall, 56% of launches were sent into low Earth orbit (LEO) and 32% were sent into geosynchronous Earth orbit (GEO) or geostationary Earth orbit (GSO). Although SLVs servicing missions to these orbital regimes featured success rates above 92%, LEO launches accounted for approximately 73% of total launch failures, while GEO-GSO launches accounted for 18% of failures. Specifically looking at U.S. and Russian SLV families, the main subsystem responsible for failure was propulsion, which accounted for 72.3% of 47 reported launch failures. First-level and second-level (using method of moments) Bayesian techniques were performed on 37 launch vehicle families and it was found that vehicles with a high accrual of launches performed with a higher reliability than vehicles with a low number of launches, resulting in first-level success rates of greater than 90%.