Launch Vehicle Research Articles

Experimental Investigation of a H2O2 Hybrid Rocket with Different Swirl Injections and Fuels

Hybrid rockets have very interesting characteristics like simplicity, reliability, safety, thrust modulation, environmental friendliness and lower costs, which make them very attractive for several applications like sounding rockets, small launch vehicles, upper stages, hypersonic test-beds and planetary landers. In recent years, advancements have been made to increase hybrid motor performance, and two of the most promising solutions are vortex injection and paraffin-based fuels. Moreover, both technologies can be also used to tailor the fuel regression rate, in the first case varying the swirl intensity, and in the second case with the amount and type of additives. In this way, it is possible not only to design high-performing hybrid motors, but also to adjust their grain and chamber geometries to different mission requirements, particularly regarding thrust and burning time. In this paper, the knowledge about these two technical solutions and their coupling is extended. Three sets of experimental campaigns were performed in the frame of the Italian Space Agency-sponsored PHAEDRA program. The first one investigated a reference paraffin fuel with axial and standard vortex injection. The second campaign tested vortex injection with low values of swirl numbers down to 0.5 with a conventional plastic fuel, namely polyethylene. Finally, the last campaign tested another, lower regressing, paraffin-based fuel with the same low swirl numbers as the second campaign.

Design and Rate Control of Large Titanium Alloy Springs for Aerospace Applications

During the separation between satellite and launch vehicles, large steel springs are often used as compression separation spring sets in a catapult separation system. Replacing the steel springs with titanium alloy springs could reduce weight by about 50%. Although titanium alloy springs have been widely used in the aerospace field due to their excellent performance, there are few reports on the design of high-precision titanium alloy springs. The current spring design standards mainly focus on steel springs with helix angles between 5° and 9°, which are not applicable to titanium springs. Moreover, the change in spring rate with ambient temperature should also be considered. In this paper, β-C titanium alloy was used to design and prepare large compression separation springs, replacing steel springs in the catapult separation system. The design of titanium alloy springs took into account the big helix angle. The relationship between helix angle and the number of active coils was calculated. The parameters of titanium alloy springs were determined by the shear stress of the spring at working length. The effects of aging temperature and aging duration on the mechanical properties and modulus of β-C alloy were studied. By adjusting the aging process, the β-C alloy spring rate was controlled to meet the design requirements. The effect of ambient temperature on the mechanical properties and modulus of β-C titanium alloy were also investigated. It was found that as the ambient temperature increased, the rate of the β-C alloy spring gradually decreased.

Flow and Aeroacoustic Characteristics of Underexpanded Supersonic Jets Exhausting from a Conical Converging Nozzle

Ensuring the safety of space flights and solving the problems of reducing acoustic loads during the launch of space vehicles requires not only the development of new technical systems for launch complexes, but also methods for the numerical simulation of fluid and aeroacoustic fields generated by supersonic jets. The growing regulations for space vehicle noise also explain the interest in developing models and techniques that anticipate flow and the aeroacoustic characteristics of supersonic jets. Together with integral techniques for computing far-field noise, development of relevant mathematical models and implementation of numerical tools, the concepts of computational fluid dynamics (CFD) and computational aeroacoustics (CAA) are covered. The noise generated by a supersonic underexpanded jet is used to illustrate the capabilities of current numerical modelling and simulation tools. The jet structure, flow properties, and aeroacoustic quantities are affected by the nozzle pressure ratio. The outcomes of numerical simulation are contrasted with existing experimental and computational data. The available numerical modelling and simulation tools facilitate the development of novel computational methods and methodologies for challenges in CFD and CAA, in addition to solving research and engineering problems.

Unsteady Surface Pressure Characterization on Launch Vehicle in Liftoff Configuration

Acquisition of unsteady pressure using remote scanners can provide practical advantages over making the same measurements directly using flush-mounted pressure transducers. However, pneumatic distortion of the signal is introduced by the tubing between the sensing module and surface port. This typically limits the usefulness of such measurements to averaged, steady data only, restricting the application of pressure scanners for unsteady pressure measurements in wind-tunnel and flight-testing applications. In this work, unsteady pressure on the surface of a 1.75%-scale model of the NASA Space Launch System Block 1B Cargo vehicle was remotely measured using an electronic scanning pressure module at the NASA Langley Research Center 14-by 22-Foot Subsonic Tunnel. Subsequent reconstruction using an empirically tuned, Wiener-filtered inverse transfer function produced excellent agreement compared to nearby surface-mounted transducers within a usable bandwidth of approximately 500 Hz. Interactions between the launch vehicle and the support tower produced strong oscillations that were absent in the vehicle-alone configuration, and it is likely that the wake interaction leads to a resonant flow response when the launch tower is immediately downstream of the vehicle.

Multi-disciplinary optimization of single-stage hybrid rockets for lunar ascent

This study aims to determine the optimal design and ascent trajectory for a single-stage hybrid rocket in the context of a lunar ascent mission, with the objective of reaching a 100 km circular orbit on the Moon equatorial plane. A multi-disciplinary optimization problem is formulated, integrating flight design variables and launch vehicle design variables to achieve the best mission scenario. A zero-dimensional internal ballistics model is employed to estimate key parameters such as thrust, mass flow rate, and gross mass of the hybrid rocket, while a three-degree of freedom dynamical model is utilized to simulate the ascent trajectory. The resulting integrated optimization problem is solved using an in-house heuristic algorithm that combines particle swarm and genetic algorithms. Two propellant combinations, liquid oxygen/paraffin-wax and hydrogen peroxide (90%)/high-density polyethylene, are evaluated, considering both their performance in the integrated system and the feasibility of the proposed mission. After conducting a sensitivity analysis to investigate the influence of particle count in the optimization process, results for both propellant combinations are presented. While paraffin and oxygen provide the highest payload ratio, hydrogen peroxide and polyethylene entail a more feasible and flexible design for a lunar mission.

Exploring the Feasibility of Kuiper Belt Missions Supported by Centrifugal Nuclear Thermal Propulsion

Missions to the Kuiper belt have previously been carried out only as flybys and with very small payloads. Investigating launch windows for Kuiper belt missions supported by centrifugal nuclear thermal propulsion (CNTP) contributes to defining its operational use case. Results indicate that CNTP enables rendezvous missions to the Kuiper belt, both with direct transfer trajectories or planetary gravity assist trajectories, although there are many challenges to making these mission architectures feasible. The direct trajectories have transfer times of roughly 14 to 16 years while combining CNTP with gravity assists from Jupiter could lower transfer time to as low as 10 to 12 years to Kuiper belt objects such as Pluto and Quaoar. These missions are then shown to inform the architecture of the CNTP injection stage vehicle, which can be supported by heavy and super-heavy commercial launch vehicles with a single launch. Last, drawbacks of the mission and vehicle architectures are given that impose limits on the use case for CNTP on these missions.

Experimental Investigation on Transmission of Sound in To Metallic Enclosures During Static Firing Tests Of Solid Rocket Motor

Satellite launch vehicle during lift-off generates intense dynamic environment mainly the acoustic noise, which in turn acts on the nearby launch pad systems. Reliable operation of these launch pad ground systems needs to be ensured by safeguarding against these dynamic loads with adequate protection measures. For future launch vehicle programs where the launch turnaround time is a crucial aspect, sensitive systems like ground check-out electronic packages are planned to be placed adjacent to the launch vehicle mounting pedestal on the launch pad to ensure minimal time for preparation. The enclosures housing these electronic packages needs to be configured to minimize the transmission of acoustic levels being imposed from the lift-off induced noise. In the present work, an experimental investigation is carried out on acoustic transmission characteristics of metallic enclosures of different material and thickness during the static firing tests of solid rocket motors. Acoustics generated from the supersonic jet noise has a wide frequency band from 100 to 10,000 Hz with sound power peaking near 1,000 Hz. The electronic components and packages, which are generally vulnerable to the high frequency noise, need to be protected from this lift-off induced acoustics. The enclosures are placed in the far field of this jet noise and the transmission loss measurements are obtained. Variation of transmission loss with material and thickness of the enclosures are studied with specific attention to the internal acoustic modes and the coincidence phenomena. Subsequently, the enclosures treated with acoustic blankets of different thickness are investigated to verify its influence on internal acoustic modes and also the additional reduction in transmission of sound in the dominant frequency band of supersonic jet noise.

COMPARATIVE STUDY OF THE MATHEMATICAL MODEL OF THE PROPELLANT TANK PRESSURIZATION SYSTEM

Abstract. The pressurization system of launch vehicle (LV) fuel tanks serves to maintain a specified excess pressure in the tank's gas cushion, based on calculated requirements for fuel component supply and tank design. This necessitates the coordination of all elements within the pressurization system. This paper examines a pressurization system consisting of single high-pressure cylinder, the fuel tank and the connecting line with the metering element. To maintain the required excess pressure in the gas cushion, pressurization gas is supplied with the mass flow rate dependent on the parameters within the high-pressure cylinder. This excess pressure ensures a consistent flow rate of fuel components to the pumps or combustion chamber. System elements such as the gas reducer and gas pipes are treated as a single line with an equivalent flow rate coefficient. The objective of this work was to develop a mathematical model of the system described above. The suggested model includes a system of ordinary differential equations for both the high-pressure tank and the fuel tank, which is derived from the conservation laws of mass and energy, as well as the equation of state for the pressurization gas. The model takes into account the gas reducer by allowing for changes in the area of the gas outflow section over time. The model's operation was simulated, revealing that with a constant cross-sectional area of the gas line, the deviation in the mass flow rate of the fuel component during tank emptying is up to 4% of the nominal value. Implementing a linear law of area change reduced this deviation to 1%. Overall, the results confirm the accuracy of the proposed method for calculating tank parameters, making it suitable for designing rocket pressurization systems.

Towards an extensible model-based digital twin framework for space launch vehicles

The concept of Digital Twin (DT) is increasingly applied to systems on different levels of abstraction across domains, to support monitoring, analysis, diagnosis, decision making and automated control. Whilst the interest in applying DT is growing, the definition of DT is unclear, neither is there a clear pathway to develop DT to fully realise its capacities. In this paper, we revise the concept of DT and its categorisation. We propose a DT maturity matrix, based on which we propose a model-based DT development methodology. We also discuss how model-based tools can be used to support the methodology and present our own supporting tool. We report our preliminary findings with a discussion on a case study, in which we use our proposed methodology and our supporting tool to develop an extensible DT platform for the assurance of Electrical and Electronics systems of space launch vehicles.

Intelligent design exploration method for complex engineered system architecture generation

ABSTRACT Owing to the discontinuous specificity and complexity of architecture design space, the issue of selecting and combining components comprising the engineered system, numerous constraints and associations need to be accounted for, adding up to a complex and substantial cognitive load on the system architects, which makes it challenging to tackle the current demand of adaptive improvements or innovative upgrading of the existing mature architectural solutions. To this end, this paper proposes an intelligent design exploration method for complex system architecture generation with reinforcement learning. The architectural design space (ADS) is identified by defining the dimensions of ADS, including model, quantity, and design chain, as well as the mathematical boundaries and representation to facilitate computable intelligent design exploration. On this basis, by adopting AI techniques primarily based on reinforcement learning, a massive and reliable architectural scheme is rapidly generated, and a more satisfying and robust architectural solution is selected by accessing the fuzzy Pareto frontier. Validation of the method is demonstrated through a case study of a launch vehicle's first and second-stage separation system. This research contributes valuable insights to overcoming the limitations of traditional techniques and enhancing the efficiency of the generative design and decision-making for complex engineered system architecture.

Ellipsoidal estimation of motion parameters of a non-cooperative space vehicle from visual information

The use of near-Earth space is currently complicated by presence of space debris objects in Earth’s orbit, which include spent stages of launch vehicles, inoperative satellites, and other large and small artificial objects. One approach to solving the problem of space debris involves docking and capturing a non-cooperative space object or spacecraft (target) with a so-called service spacecraft (chaser) for further actions to repair, refuel or change its orbit. Rendezvous and docking are complicated by the rotation of uncontrolled space objects caused by various factors. To perform this task, it is necessary to know the parameters of the orbital, rotational and relative motion of the target. The parameters of the orbital motion of such objects are usually known quite accurately from measurements from the Earth. This paper examines the case of a tumbling non-cooperative target located in an elliptical orbit. It is assumed that the target relative position and orientation are measured by the computer vision system (CVS) of the chaser. In this case, the position and orientation of the graphical reference frame (GRF) associated with the known 3-D graphical model of the target are determined relative to the reference frame associated with the chaser. The specific type of CVS is not considered. It is assumed that the chaser can carry out some maneuvers near the target and all parameters of the chaser angular motion are known. Thus, the attitude of the GRF relative to inertial reference frame (IRF) can be determined. The measured parameters are not enough to ensure safe rendezvous and docking with the target. To complete this task, it is necessary to determine all kinematic and dynamic parameters of the relative motion between the spacecraft. The rest of the required parameters are estimated. Orientation and rotation parameterization is done using quaternions. The angular motion equation of the target is considered in the GRF. This makes the angular velocity estimation faster and the inertia tensor estimation more stable. Stochastic characteristics of measurement errors are considered to be unknown and are not used. The only information about the errors is the bounds of their values. To determine the relative motion parameters, we use a new dynamic set-membership filter with ellipsoidal estimates. The filter can be successfully implemented on low-power on-board processors. The properties of the proposed algorithm are demonstrated using numerical simulation. The results obtained are expected to be used in the development of a navigation system for the rendezvous and docking, developed by a group of Ukrainian space industry enterprises under the leadership of the LLC “Kurs-orbital” (https://kursorbital.com/).

Development of Jets rig dedicated for an Active Launch Escape Abort System Wind Tunnel Model

The development of space launch systems requires rigorous testing and validation of safety mechanisms to ensure the protection of human life and mission-critical assets. One such safety mechanism is the Active Launch Escape/Abort System (ALEAS), designed to swiftly extract crew and spacecraft from a malfunctioning launch vehicle. To evaluate the performance of ALEAS, wind tunnel testing is indispensable. This paper presents the development of a specialized Jets Rig tailored for wind tunnel testing of an ALEAS model. The primary objectives of this research activity include the design, construction, and validation of a Jets Rig that can accurately simulate the propulsion dynamics of an ALEAS system within a wind tunnel environment. The Jets Rig incorporates a special instrumentation and control systems to replicate the complex operational conditions experienced during a launch abort scenario. By achieving this, it enables a comprehensive assessment of ALEAS performance, including thrust duration and plume interaction effects, among others. Key aspects of this study encompass the aerodynamic and structural considerations involved in designing the Jets Rig, the integration of high-fidelity sensors and data acquisition systems, and the development of advanced computational models for predictive analysis. Additionally, the research explores the challenges and solutions associated with the scalability of the Jets Rig to accommodate varying scales of ALEAS models. The findings from this project hold significant implications for advancing the safety and reliability of crewed space missions. A comprehensive understanding of ALEAS performance in a wind tunnel setting allows for the refinement of design parameters, algorithms, and the enhancement of abort system efficiency. Ultimately, the successful development of this dedicated Jets Rig contributes to the broader mission of ensuring the safe exploration of space and the protection of human life in the challenging and dynamic environment of space launch.

DETERMINATION OF THE OPTIMUM DESIGN OF THE MEANS TO ENSURE FUEL CONTINUITY USING THE METHOD OF NUMERICAL SIMULATION

Residues of liquid propellant components in the power system of the propulsion system at the end of the operation of the launch vehicle stage significantly affect its energy characteristics. The fuel continuity features found in modern launch vehicle tanks ensure a continuous supply of propellant components from the tank to the propulsion system without disrupting flow and mini- mizing residues. In the fuel tank of launch vehicles, the presence of a tunnel pipeline complicates, and in certain cases excludes the possibility of taking fuel from the pole of the tank. Solving the problem of fuel intake, in this case, allows the use of means to ensure the continuity of fuel of the side type. However, the use of such a device, due to the displacement relative to the pole of the bottom of the tank by a certain angle of the sampling point, can lead to an uneven level of breakthrough and immersion of the boost gas at the inlet to the flow line and, as a result, a significant increase in the remainder of the fuel components. The authors searched for and substantiated the optimal design of the means for ensuring fuel continuity on the example of the fuel tank of the first stage of the Cyclone-4 launch vehicle, which is equipped with a side-type fuel continuity means in the form of a profiled plate (plate). The designs of siphon and annular fuel continuity means are considered. An analytical calculation was carried out using empirical dependences, a physical experiment, and a numerical simulation of their main parameters. Based on the work performed, a comparative analysis of the change in the energy characteristics of the launch vehicle depending on the chosen design was carried out. The efficiency of the operation of the considered means of ensuring the continuity of the fuel is determined in terms of the mass parameter of the remainder of the propellant components under static conditions. The result of the experimental and computational-analytical work was to obtain the most optimal variant of the fuel continuity means, which turned out to be a siphon according to several parameters. The introduction of a siphon means of ensuring continuity, as more optimal, into the design of the fuel tank of the Cyclone-4 launch vehicle will improve its energy characteristics by increasing the mass of the output payload by 5.4 kg. Also, using the numerical modeling methods (CFD methods) for design work on devel- opment, optimization, and improvement, instead of empirical and semi-empirical dependencies, allows you to: obtain more accurate results; reduce the number of experimental designs; reduce the required number of tests; reduce the duration of exami- nations; reduce the total duration of computational and design and experimental work; reduce the overall material and technical costs and, as a result, reduce the final cost of the development product, which will increase its competitiveness.

EVALUATION OF THE SCATTER OF LIQUID LAUNCH VEHICLE POGO OSCILLATION AMPLITUDES DUE TO THE INFLUENCE OF THE SCATTER OF INTERNAL FACTORS

Almost all liquid launch vehicle developers faced the problem of ensuring stability in relation to POGO oscillations. The level of POGO amplitudes oscillations of the launch vehicle can be significantly affected by the scatter of internal factors. The study aims to create a mathematical model that can determine the range of POGO amplitudes in liquid launch vehicles. This will be demonstrated through the example of the Dnipro launch vehicle, which is affected by a variety of internal factors that cause its POGO amplitudes to vary. We developed the non-linear non-stationary mathematical model of POGO oscillations of the prototype of the Dnipro space launch vehicle. The model is built by taking into account the two lower vibration modes of the LV structure, two lower oscillation modes of the oxidizer feedline, and the first oscillation mode of the fuel feedline of the propulsion system. Modeling of dynamic processes was conducted in a combination of four liquid rocket engines based on the schematic of the staged rocket engine. The simulation takes into account cavitation phenomena in the engine pumps and delay times in the gas generators’ chambers. We have developed a method for determining the scatter of the POGO oscillations caused by the action of internal factors, which is based on the use of the LP uniformly distributed sequences. As internal factors, the frequencies, decrements, and shapes of LV structural oscillation modes, the values of pressurization of the propellant tanks, and the engines’ specific thrust impulses were considered. Based on the results of the calculations, the dependence of the POGO amplitudes in two regions of LV instability was determined, and the lower and upper enveloping curves for the POGO amplitudes were constructed. It is shown that the maximum POGO amplitudes oscillations in the first region of instability lie in the range from 0.23 g to 0.72 g and in the second region of instability — from 0 g to 0.60 g. Variants of combinations of internal factors, which provided the largest and smallest values of POGO amplitudes, were analyzed. This made it possible to determine the internal factors, the scatter of which has the greatest effect on the POGO amplitudes scatter: frequency, decrement, shape coefficients of oscillations of the oxidizer feedlines and the LV 1st mode structural longitudinal oscillations in the payload cross-section.

Alleviation of SWBLI over the payload of a launch vehicle by change of nose shape

Wind tunnel tests on a typical heavy-lift launch vehicle configuration featuring a core vehicle flanked by two strap-on boosters showed very high levels of unsteady pressures over the payload region at transonic Mach numbers when the nose had a cone-cylinder shape. This was due to very high levels of Shock Wave Boundary Layer Interactions (SWBLI) associated with the conical shape, which caused alternating flow featuring the formation of a unsteady λ-shock system over the payload region and a pair of counter-rotating vortices. The phenomenon was found to strongly depend on the proximity of the noses of the strap-on boosters to the boat tail downstream of the payload region. When the conical nose was replaced by an ogive-shaped nose, significant reduction in SWBLI occurred, with dramatic reduction of shock oscillations, pressure fluctuations, replacement of alternating flow by steady flow and disappearance of the vortex-pair. Time-averaged pressure fluctuations indicate reduction in pressure gradients due to a more gradual expansion of flow over the ogive nose compared to conical nose, an observation also supported by CFD simulations. The alleviation of SWBLI with ogive nose was so effective that the presence or absence of strap-on boosters had practically no effect on pressure fluctuations over the payload region. Unlike the conical shape, the ogive nose cone shape seems to meet both the shape and pressure parameter-criteria of NASA to ensure buffet-free environment.

Computational Techniques to Generate Space Launch System Aerodynamic Databases

This document describes the reasoning and trade studies used to evaluate tools for constructing the aerodynamic lineload databases for the liftoff and transition phases of flight for the Space Launch System. Three computational fluid dynamics codes (USM3D, FUN3D, and Kestrel) were investigated with various turbulence models as well as detached eddy simulation variants for the launch vehicle in free air and in proximity to the tower. Decisions were made mostly based on results from brief developmental studies performed in response to specific, unforeseen challenges that were encountered in the analysis of a given configuration. The need to develop databases in a timely manner, as well as to accurately capture the expected leeward-wake flowfield characteristics, led to the selection of the Kestrel flow solver with its delayed detached eddy simulation method, the Spalart–Allmaras turbulence model, and the adaptive mesh refinement capability in the off-body Cartesian grid region.

Investigation of different strategies for access to space of small satellites on a defined LEO orbit

In recent years small satellites have shifted from being secondary items to dominate the space market thanks to, but not only, the development of the large LEO constellations. For small satellites to be highly effective, particularly when arranged in such constellations, each satellite has to be placed in a specific orbital plane and orbital position.Currently, three main options are available to reach a specific orbit. The first, more straightforward solution is to employ a dedicated launch with a small launch vehicle. Plenty of small launchers are in development around the world (with few already operational); however, their cost per kilogram is predicted to be much higher than for large launchers. The second possibility is the exploitation of a rideshare option, where the satellite is transported by a large launcher together with other payloads on a general predetermined orbit. Afterwards the satellite needs to be transferred to its designed orbit. This can be done in two ways: through the use of a satellite carrier (also called self-propelled dispenser) or with the satellite own propulsion system. In both last cases, the mass transported by the launcher into the release orbit is higher than the final one necessary for the nominal mission, impacting total costs.In this paper these three possibilities are compared considering the need to reach various specific orbits, starting from a different release one in the case of the rideshare options. First of all, the change in velocity for different orbital parameters (altitude, eccentricity, phase, argument of perigee, inclination, RAAN) is computed. Afterwards the propulsion mass budget is calculated.Everything else being equal, it is demonstrated mathematically that a dispenser is inherently less efficient than a group of autonomous satellites, particularly for the RAAN change and a large number of carried satellites. However, it is not always possible or convenient to provide small satellites, particularly the smallest ones (nanosats/cubesats/microsats) with comparable propulsion capabilities of a larger dispenser, making the latter still an attractive option in several situations. A cost analysis also shows that, particularly for sophisticated small satellites, when the final orbit is far from the release one, a dedicated small launch vehicle can be cost competitive with the nominally much cheaper large launcher.

A dual‐mode integration framework and application to agile feedback design of launch vehicles

AbstractIn recent decades, the design of complex systems like launch vehicles in the aerospace industry has presented engineers with challenges that go beyond system complexity. Issues such as time‐to‐market pressures and intricate industrial processes have underscored the increasing significance of agile design methodologies. Agile design is derived from the simplification of the design process and enhancing cross‐domain data transmission and feedback. While methods based on model‐based system engineering have improved iteration times in system architecture design, challenges persist in cross‐domain data transmission. Due to the diversity of complex system models and data, a single‐mode integration method is difficult to realize the data link construction of all tools used. To address this challenge, this paper proposes a dual‐mode data integration framework with expansibility, universality, and cost‐efficiency which leverages the benefits of Remote Procedure Call and Intermediate Exchange Module, addressing the challenge of constructing cross‐domain data links under single‐mode integration. In this study, two critical requirements of the first‐ and second‐stage separation systems, namely, weight and minimum separation gap, are selected for data feedback. A Modelica‐based multiphysics simulation model is developed in MWorks; visualization and computation of the minimum gap are carried out in CoppeliaSim. To bridge the gap between domain‐specific tools, Matlab and Functional Mock‐up Unit modules are introduced as middleware, facilitating data feedback linkage. The entire simulation process is orchestrated using activity diagrams in the MagicDraw tool. The study delves into the influence of critical design parameters, such as the initial angular velocity of separation and the thrust of the retro rocket, on the minimum separation gap. It provides an analysis of minimum separation gap variations under uncertain operating conditions and examines design margins. Significantly, the paper highlights the significance of controlling the initial angular velocity during separation and the reliability of the retro rocket, providing essential decision supports and valuable insights to agile the process of system design.

SunbYte: an autonomous pointing framework for low-cost robotic solar telescopes on high altitude balloons

The design and usability of a fully autonomous robotic control system (SunbYte - Sheffield University Balloon “lYfted” TElescope) for solar tracking and observational applications onboard high-altitude balloons are addressed here. The design is based on a six-step development plan balancing scientific objectives and practical engineering requirements. The high-altitude solar observational system includes low-cost components such as a Cassegrain-type telescope, stepper motors, harmonic drives, USB cameras and microprocessors. OpenCV installed from ROS (Robotic Operating System), python and C facilitated the collection, compression, and processing of housekeeping and scientific data. This processed data was then transmitted to the ground station through the launch vehicle’s telecommunication link. The SunbYte system allows the brightest spot in the sky, the sun, to be identified, and a telescope pointed towards it with high enough accuracy that a scientific camera can capture images. This paper gathers and presents the results from primarily two missions with the High-Altitude Student Platform (HASP, NASA Balloon Program office and LaSpace). Additionally, a discussion will be made comparing these with an earlier iteration flown with the German-Swedish “REXUS/BEXUS” programme coordinated by the European Space Agency. By capturing and analysing a series of tracking images with the location of the Sun at the calibrated centre, the system demonstrated the tracking capabilities on an unstable balloon during ascent. Housekeeping sensor data was collected to further analyse the thermal and mechanical performance. The low temperature increased friction in the drive train and reduced the responsiveness of the harmonic drive actuation system. This caused some issues which require further work in future missions, for example, with SunbYte 4 and its work when flying with the HEMERA ZPB (Zero Pressure Balloon) program.

Оценка риска здоровью населения, проживающего на территории влияния космодрома «Восточный», в условиях многофакторной многосредовой экспозиции имических веществ

Rocket and space activities occupy an important place among modern sources of impact on the environment. Its results can manifest themselves in an area where a cosmodrome is located, adjacent territories, in areas where separated parts of a launch vehicle fall. This impact has certain specific features that create population health risks. Health risks for population of Tsiolkovsky settlement were assessed upon exposure to chemicals that pollute the environment in accordance with the methodical approaches stipulated in the valid Guide (R 2.1.10.3968-23). Initial data were represented by results of environmental surveillance accomplished within social-hygienic monitoring and in field conditions as well as by data collected for a project aimed at substantiating a sanitary protection zone for the Vostochny Cosmodrome launch sites. Average annual concentrations and their 95 % confidence intervals, which were established based on average daily concentrations, were used to calculate chronic chemical exposures. The 95 % percentile of maximum single concentrations was employed to calculate acute exposures. Ambient air was established to make the greatest contribution (88.99 %) into non-carcinogenic health risks for people residing in Tsiolkovsky settlement. Particulate matter made the greatest contribution (42.4 %) to the risk of chronic non-carcinogenic effects. Alerting chronic non-carcinogenic risks of respiratory diseases (up to 5.9 HI) and the blood system (up to 4.1 HI) were established under combined inhalation chemical exposures. Major contributions to non-carcinogenic risks are made by particulate matter (up to 42.4 %), lead (up to 25.9 %), nitrogen oxide (up to 20.2 %), nitrogen dioxide and carbon oxide (23.1 % each).