Heavy Launch Vehicle Research Articles

Finite-time state-constrained adaptive fault-tolerant control for heavy launch vehicles with input quantization

This paper proposes the finite-time state-constrained adaptive fault-tolerant control method for a class of heavy launch vehicles subject to quantized input signals and actuator faults. Firstly, the attitude dynamic model of the heavy launch vehicles suffering from the actuator faults and disturbances is formulated. Secondly, to deal with the completely unknown nonlinear functions, radial basis function neural networks (RBFNNs) are introduced for approximation and compensation. Meanwhile, the barrier Lyapunov function (BLF) is introduced to ensure that all the states in the closed-loop system are bounded and the constraints of the system states are satisfied. As a result, the finite-time state-constrained adaptive fault-tolerant control structure is constructed for the heavy launch vehicles and the semi-globally practical finite-time stability (SGPFS) of the closed-loop control system is proved. Finally, numerical simulation results show the effectiveness and the satisfactory performance of the proposed control algorithm.

PREREQUISITES FOR CREATING ULTRA-LIGHT LAUNCH VEHICLES WITH POLYMER BODIES

The relevance of the problem of creating new light and ultra-light launch vehicles from the point of view of the development of the modern space industry is shown. First of all, this problem is related to the expansion of the market and the scope of use of small satellites, most of which are now launched into orbit as passing load by medium and heavy launch vehicles, which is not at all prompt and too expensive. The estimated forecast of the small satellite launch market, which is presented herein, indicates a growing demand for their launch services. A low-cost launch vehicle for the launch of small satellites would greatly expand their use, development, and production. This indicates the need for creation of special-purpose cheap transport vehicles for the operative decision of tasks of launching such satellites into orbit. It is shown that the main directions for increasing the efficiency and reducing the cost of launches are reducing the cost of making a transport space system, increasing of mass the payload, increasing the multiplicity of reusable uses, and reducing one-time costs per start. World trends in solving this problem are connected with the solution problem of creating ultra-light launch vehicles. The historical review of studies on prototype rockets that use the principle of reducing the mass of a launch vehicle during flight by burning spent structural elements indicates the possibility of creating launch vehicles on this principle by using complex hydrocarbon polymers as structural materials. It is shown that polymer materials, in particular polyethylene, can become the main material for the manufacture of fuel tanks for solid rockets, which use the new principle of burning the design during flight. The use of fuel tanks as fuel will give a possibility to reduce the cost of the propulsion system, which is one of the most expensive parts of the launch vehicle. The development of such launch vehicles and the commercialization of scientific research will allow domestic aerospace companies to occupy a niche in the market of small satellite launches.

Vector acoustic intensity analysis of noise from the Delta IV Heavy launch vehicle during liftoff

The noise sources within a turbulent rocket plume are not well understood, let alone the radiation from multiple rocket nozzles. Even less is known about the noise sources during launch vehicle liftoff. This paper seeks to simultaneously address these noise source analysis challenges using data collected during the NROL-82 Delta IV Heavy (DIVH) launch from Vandenberg Space Force Base. The three-core DIVH’s liftoff noise was measured by a ground-based four-microphone array at a distance of 330 m. Hart et al. [Proc. Mtgs. Acoust. 45, 040003 (2022)] previously used this array with a cross-correlation technique to identify the overall noise source radiation locus as it switched from the flame trench exit to about 55 m downstream of the nozzle exit plane. Here, vector intensity is used to localize the frequency-dependent noise source axial distribution using a variant of the phase and amplitude gradient estimator method for acoustic intensity [Thomas et al., J. Acoust. Soc. Am. 137, 3366–3376 (2015)]. The results help clarify a complex noise source generation process that is both time and frequency-dependent and involves both free and impinging jet noise phenomena.

Assessing azimuthal asymmetry in the noise radiation from a three-core launch vehicle during liftoff

Creating accurate rocket noise models is important for assessing impacts on humans, the environment, and payloads. The United Launch Alliance Delta IV Heavy launch vehicle is unique because of the separation of the three cores and their associated RS-68A nozzles. This makes it a good candidate for determining how the asymmetry of nozzle configuration affects noise radiation, which can affect noise models. The NROL-82 and NROL-91 missions both launched from Vandenberg Spaceforce Base using Delta IV Heavy Vehicles. For both of these launches, acoustic data were recorded between ∼0.9 and ∼5.2 km from the vehicle at different azimuths to determine the extent of azimuthal asymmetry in noise radiation. Maximum overall sound pressure level, spectra, and overall power level were determined for each launch. Methods for comparing the datasets and results will be discussed.

Optical effects of running spacecraft engines in the lower thermosphere

This paper provides a brief overview on optical effects during operation of spacecraft (SC) onboard engines in the lower thermosphere according to observational data from the ISTP SB RAS Geophyisical Observatory. We present the results of detected disturbances in the night airglow during operation of SC vernier engines in the F2-region of the ionosphere in the “Radar–Progress” space experiment. With weights of combustion products of ≤10 kg injected by SC vernier engines, the atmospheric emission of [OI] 630.0 nm atomic oxygen is enhanced. We also show optical effects from the launches and passages of heavy launch vehicles “Energiya” from the Skif-DM spacecraft on May 15, 1987 and “Proton-M” from the Yamal-601 spacecraft on May 30, 2019 from Baikonur in the zone far from the launch site. We explore the possibility of enhanced [OI] 557.7 nm atmospheric emission due to chemical modification of the ionosphere in the E-region during the flight of the Energiya space system.

Оптические эффекты работы двигателей космических аппаратов на высотах нижней термосферы

This paper provides a brief overview on optical effects during operation of spacecraft (SC) onboard engines in the lower thermosphere according to observational data from the ISTP SB RAS Geophyisical Observatory. We present the results of detected disturbances in the night airglow during operation of SC vernier engines in the F2-region of the ionosphere in the “Radar–Progress” space experiment. With weights of combustion products of ≤10 kg injected by SC vernier engines, the atmospheric emission of [OI] 630.0 nm atomic oxygen is enhanced. We also show optical effects from the launches and passages of heavy launch vehicles “Energiya” from the Skif-DM spacecraft on May 15, 1987 and “Proton-M” from the Yamal-601 spacecraft on May 30, 2019 from Baikonur in the zone far from the launch site. We explore the possibility of enhanced [OI] 557.7 nm atmospheric emission due to chemical modification of the ionosphere in the E-region during the flight of the Energiya space system.

Elimination of root kissing bond in friction stir welding of thick plates

• The root-enhanced friction stir welding was proposed to eliminate root kissing bond. • The root tolerance had been expanded from 0.1 mm to 0.6 mm. • The enlarged-tip pin increased heat input and material deformation at the bottom. • The tensile properties were increased to 345 MPa. Root kissing bond in thick-plate friction stir welded joints limits the application of plates in fuel tanks of heavy launch vehicles. Here a root-enhanced friction stir welding from aspects of the on-demand distribution of heat input and plastic deformation through an enlarged-tip pin was proposed. The pin was designed with an enlarged tip for enhanced plasticizing and forging on root materials. Root tolerance, defined by the distance between the pin tip and the lower surface of the plate without a root kissing bond, was proposed to evaluate the feasibility of the designed pin. This value increased from 0.1 mm to 0.6 mm compared to the conventional conical pin. The ultimate tensile strength and elongation reached 345±4 MPa and 11.8±0.7%, which were 11.3% and 87.3% higher than the joint by the conical pin.

Europe’s Space Exploration and Korea’s Space Exploration Strategy from the Perspective of Science and Technology Diplomacy

Space exploration is an area where international cooperation takes place more actively than any other space activities such as Earth observation, communication and navigation. This is because a country cannot afford a huge budget to have full infrastructure for deep space exploration, such as a heavy launch vehicle, communication and energy infrastructure, and human habitats, and has learned that it is not sustainable. Korea expressed its willingness to join humanity’s epic exploration journey by signing the Artemis Accords in 2021 and launching Danuri lunar orbiter in 2022. The beginning of space exploration means that Korea’s space activities have expanded beyond the stage of focusing only on technology development to set norms necessary to accompany other countries and cooperate diplomatically to solve exposed problems. This paper analyzed European space policy and space exploration, which are most actively participating in the Artemis Program and exerting diplomatic power in the space field, from the perspective of science and technology diplomacy. The suggestions for Korea’s space exploration strategy from the perspective of science and technology diplomacy were drawn by examining the international cooperation strategies in Europe’s space activities ranging from space policy, space strategy, and space exploration program to project units.

Climate action and growing electricity demand: Meeting both challenges in the 21st century with space-based solar power delivered by space elevator

Global climate action is the grand challenge of the 21st century. Large reductions in greenhouse gas emissions are needed, with net-zero CO2 emissions by mid-century a requirement for many scenarios that restrain global warming below 1.5 °C by 2100. This will require massive deployment of renewables such as ground-based solar and wind, increased long-distance grid transmission capacity, and electrification of much of the global economy. Increased electrification, coupled with rising standards of living in developing countries, will drive ever-larger electricity demand to 2050 and beyond. And the intermittency, land-use, and transmission requirements of ground-based solar and wind present challenges to large-scale deployment to meet Net Zero targets. Space-based solar power (SSP) presents an enticing alternative, and with full reusability of upcoming heavy launch vehicles and modular SSP architectures, the economics of deployment are expected to become favorable in the coming decades. Efforts exist today to deploy demonstrators this decade, with initial production systems by 2050. However, deploying SSP at the scale required to address rising, global electricity demand will itself require orders-of-magnitude greater launch frequencies. If heavy launch frequency does not increase quickly enough, due to environmental concerns, launch site competition, or other bottlenecks, the dropping cost of launch will not be sufficient to enable SSP for global impact this century. A space elevator (SE) development program like that proposed by the International Space Elevator Consortium (ISEC) offers an alternative delivery mechanism for scaling SSP beyond 2050, hedging the risk of reliance on a single delivery system while reducing the time required for large-scale SSP deployment. To de-risk possible futures in which SSP is a major part of the global energy mix this century, a dual space access architecture is proposed, leveraging the strengths of heavy launch and SEs together. Such an architecture would secure the future for green energy on Earth, in 2050 and beyond.

Afterburning Effects on Thermal Environment of Liquid Oxygen/Kerosene Rocket Under Low Altitudes

Heavy launch vehicles use liquid oxygen (LOX) and kerosene propellants for their high specific impulse and controlling precision. However, incomplete burnt exhaust gas reacts easily with air, which will influence the thermal environment. To investigate the afterburning effect on the plume flowfield at low altitudes, we established a supersonic exhaust plume model by using the three-dimensional Reynolds-averaged Navier–Stokes (RANS) methods and realizable turbulence model. Also, the nine-species and fourth-step chemical reactions are adopted to simulate the afterburning process. The validity of our numerical model is confirmed by the good agreement between calculation results and experiment data. On this basis, the thermal environment of the LOX/kerosene rocket between frozen and reaction flow at five typical altitudes are calculated and compared. The numerical results show that the secondary combustion reactions mainly occur in the mixing layer, with about an 11.9–18.6% increase in the peak temperature after considering the afterburning effects. The increase in range of the maximum temperature of the engine exhaust plume gradually reduces as the flight height increases. At the same altitude, the afterburning has more significant effect on flowfield for a larger distance from the nozzle exit. The results provide a vital foundation of analysis and calculation for the design of the LOX/kerosene rocket thermal protection.

Chemical Rocket Motors for Space Launch Applications: Parametric Approach

Hybridrocket motors are widely used in heavy launch vehicles with maximum payload capacity. The main reason for the study in this field is that the hybrid rocket motor is very easy to handle and has a simple design. In this paper, we have studied the parameters involved in designing and optimizing the hybrid rockets which provides in achieving maximum performance value. The several parameters considered for parameters involve are- burn rate, specific fuel consumption, specific impulse, efficiency. The design constraints will be responsible in achieving a high thrust level through the nozzle. The study also highlights the propellants being used in hybrid rocket motor. The collaborative research will aid in briefing the design and the working of the hybrid motors that is frequently being utilized in the industry.

POSSIBLE ENGINEERING DESIGN SOLUTIONS FOR DEVELOPENT OF HEAVY AND SUPER-HEAVY OXYGEN-KEROSENE SPACE ROCKET

The paper presents the results of an independent design study into the feasibility of developing for the Super-Heavy Space Rocket (SHSR) with 140–170 ton payload capacity a modular Launch Vehicle (LV) in the form of a tri-pack cluster of three two-stage medium LVs Soyuz-5 and three side-mounted bi-pack clusters based on coupling together the first stages of LV Soyuz-5 based on NPO Energomash oxygen-kerosene main engines RD171MV, RD180, RD191M and a four-chamber version of the RD120 engine. In addition to this, the paper discusses the feasibility of using the tri-pack cluster from the SHSR as a standalone Heavy LV with an increased payload capacity (50–60tons) to assure its more frequent use, and thus upkeep the operational reliability of the more powerful SHSR. The paper demonstrates the feasibility of designing the second-phase reusable versions of the first-stage clusters for SHSR and Heavy LV, as well as the feasibility of manned lunar missions using not two, but only one SHSR. Key words: integrated launch vehicle, launch vehicle, main engine, tri-pack cluster, bi-pack cluster, crew transportation spacecraft, lunar take-off and landing vehicle.

Past experience and future prospects for launching our country's small spacecraft from Plesetsk launch site

The paper provides an integrated analysis of the problem of the fast growing field of development and operation of Small Spacecraft (SS) constellations and participation of domestic launch site Plesetsk in solving this problem. It reviews the history of launches from Plesetsk, and describes the current status of work to build up and replenish SS-based orbital constellations. It describes major methods of SS launches. When describing multiple launches of SS on heavy launch vehicles it cites foreign practices of SpaceX Company, as well as lays down the prospects for development of our country’s heavy launch vehicles. As for the small-lift launch vehicles, it cites the experience of launching from Plesetsk of launch-vehicles Rokot and Soyuz-2 Phase 1v. For the super-small-lift launch vehicles, it cites the experience of USA and PRC, as well as our country’s development project of the launch vehicle Irkut, which is planned to be launched from Plesetsk, among other places. As for changes in the configuration of ground space infrastructure, the paper points out a trend towards processing SS at a single common processing facility (there is no need to construct processing facilities for each specific small spacecraft). The paper presents general trends in the system for operation of space facilities. The trends towards standardizing mechanical handling equipment in particular, and processing facilities in general, make it possible to lower development costs and speed up development schedules for advanced small spacecraft. Reducing small spacecraft processing time has a beneficial effect on the operational efficiency of the launch campaign, while the use of mobile checkout equipment makes it possible to quickly put together mobile and flexible processing facilities, the use of which will undoubtedly contribute to shaping the future of our country’s launch system. An important point made in the paper is the conclusion that in spite of our country’s lagging behind in super-small-lift launch vehicles, it has a significant scientific and engineering potential for developing this type of launch vehicles. Regardless of the value of other conclusions of the paper, the key point is that the current state of the ground space infrastructure facilities and the accumulated experience of specialists at Plesetsk make it possible to address the full spectrum of SS launch operations. Key words: small spacecraft, launch vehicles, ground space infrastructure, system for operation of space facilities.

Effects of segmented thermal-vibration stress relief process on residual stresses, mechanical properties and microstructures of large 2219 Al alloy rings

Large 2219 Al-Cu alloy transition rings are extensively employed in propellant tanks of heavy launch vehicles. These rings have a diameter exceeding 5 m or even reaching 10 m, with a less than 2% thickness-to-radius ratio, and low stiffness which can cause machining deformation due to the residual stress. Thus, residual stress relief of these rings is necessary in their extreme manufacture. A novel effective method is thermal-vibration stress relief (TVSR), which integrates the conventional thermal stress relief (TSR) and vibratory stress relief (VSR); however, the existing TVSR equipment cannot meet the requirements of large rings, and the underlying mechanisms of TVSR remain unclear and a quantitative interpretation is still lacking. Therefore, a segmented TVSR (STVSR) process suitable for large rings was proposed and the corresponding experiment was carried out with a self-made STVSR experiment platform. Then this study investigated the evolution and distribution laws of the residual stresses, tensile properties, Vickers hardness, dislocations, precipitated phases and metallography during STVSR. Based on the experimental results, multi-scale mechanics theory and strengthening mechanisms were applied to quantitatively reveal the underlying mechanisms of residual stress relief by STVSR. The results showed that the circumferential and axial residual stress relief rates can reach 44.43% and 45.14% after STVSR, respectively. The residual stress relief after STVSR is attributed to the dynamic evolution of dislocations and precipitated phases in the material. The improvement of mechanical properties mainly depends on the precipitated phases and is also affected by the residual stress. The findings confirm the significant effects of STVSR on metal plasticity and provide valuable insight into the underlying mechanisms of TVSR.

Planning and Design Method for Heavy Launch Vehicle Launch site for Manned Mars Exploration Mission

The manned Mars exploration mission is an important application for the future use of heavy launch vehicle. With largger scale and more system, its launch site planning design and construction are difficult. Launch vehicle overall design, launch site planning design and test launch mode are considered as a whole of feedback iteration. For the huge system engineering of manned Mars exploration, the design of the launch site is mainly affected by the mission scale, launch vehicle overall design and test launch mode. Analysis of factors such as mission scale and window cycle, more than two heavy launch vehicle launch pads are needed for the manned Mars exploration mission. Therefore, it is necessary to consider the multi-pad construction. According to the preliminary genera parameters of heavy launch vehicle and the climate characteristics of Wenchang launch site, analysis the test launch mode that suitable for China’s heavy launch vehicle. And then, the horizontal planning and the vertical plannig method are suggestted for heavy launch vehicle launch site. Thereby achieving technically feasible, safe reliable and cost controllable purposes. Provid ideas and methods for China’s future large-scale space mission launch site planning.

Приближенная методика проектировочного баллистического расчета первых ступеней ракет-носителей

The initial stages of development of new types of launch vehicles involve performing the so-called design ballistic calculations. Currently, the relevance of design ballistics tasks is noticeably increasing in connection with the problems of deep space exploration and the creation of a new generation of space rocket systems, requiring significant financial investment. When developing design methods, motion models should not be inferior in accuracy to the uncertainties of the source data, and the algorithm of the technique should reflect the relationship of design parameters with the flight characteristics of the rocket. These principles formed the basis of the proposed methods for ballistic calculations of the active section of the trajectory of the first stages of heavy and super heavy launch vehicles. The technique is based on the results of the analysis of numerous trajectories corresponding to variations of various studied factors. The trajectory calculations were carried out using the method of numerical integration of differential equations of motion in a wide range of variation of design ballistic parameters and boundary conditions. The analysis made it possible to identify some general patterns of the controlled motion and to obtain the structure of the calculated dependencies in an analytical form.

Numerical Studies on Flowfield Characteristics of Four-Engine Rockets Impinging Jet with Different Flame Deflectors

[Objective] To explore the further deep space, many countries are developing the new generation launch vehicle which has higher effective loads [1, 2]. To upgrade capacity, a scheme with multiple engines working simultaneously has often been adopted in the propulsion system of heavy launch vehicles [3, 4]. Therefore, the rocket jet impact problem has been the vital factor which seriously affected the launch safety. Due to the rocket engines are close to the launch platform during launching phase, the exhaust jets of rocket engine not only disturb the attitude of rocket in the take-off, but also have a thermal impact on the rocket base and ground control system [5]. Potential dangers would threaten secure and stable of launch missions if these problems could not be solved properly. Therefore, carrying out research on the flow field characteristics of multi-engine rockets exhaust plume has a great significance for the risk aversion and improvement of achievement ratio in launching [6-9]. This paper researches the four-engine rockets flow field during launching phase, compares and analyzes the plume impact characteristics with two shapes of flame deflector by numerical simulation methods. [Methods] The study object of this paper is a three-stage liquid rocket which utilizes kerosene/liquid oxygen (LOX) propellant as propulsion system for the first stage. Structured meshes with tensor product structure have better numerical ability than unstructured grids that using a multi-block, structured mesh to improve the fidelity of the flow solutions and assure a good resolution of shock waves in the flow field. The liquid rocket exhaust is considered as an ideal gas with the continuum assumption. The numerical solution of the compressible, multi-species, Reynolds-Averaged Navier-Stokes equations for multi-component flow has been obtained with finite volume spatial discretization on a structured three-dimensional grid implemented. The supersonic jet impingement on deflector was computed the realizable k-ε turbulence model. The model has shown superior improvements over the standard-model where the flow features include strong streamline curvature and vortices. The accuracy and effectiveness of the impact model was verified by comparison between the calculated results and the measured date. [Results] The temperature, pressure, and contours on the symmetry plane with wedge-shaped and cone-shaped deflectors are obtained, three shock cells are established before the plume impacts the deflector surface. A high temperature region occurs immediately through excessive high temperature gas accumulation on the deflector surface. The core and developed region and boundary of plume are clearly visible downstream of flow field. The supersonic exhaust plume was diverted into outlet with two directions and circumferential direction in the wedge-shaped and cone-shaped flame deflector, respectively. It can be shown from the temperature and pressure contours in the deflector surface that the region of peak temperature and pressure occur directly under the rocket nozzle, and the temperature drops gradually from the center to the outlet of the flame deflector. The simulation results show the planar static temperature contours for the rocket motor gas impingement on wedge-shaped and cone-shaped deflectors at Z/D = 3, 4.5, and 6 in axial plume direction, where Z/D = 0 corresponds to the nozzle exit. The exhaust gas inlet of the wedge-shaped deflector has the same temperature as atmosphere except for jet flow region, and not affected by the reverse flow. The temperature edge region of planar in the cone-shaped deflector, as well as the central region, is much higher than ambient temperature. The temperature in the edge at Z/D =3 where is closest to the exhaust gas inlet may reach up to 800 K. [Conclusions] By comparative analysis between the four-engine rockets impinging jet on different deflector, the main conclusions are as follow. The region of peak temperature and pressure occurs directly under the rocket nozzle. The wedge-shaped and cone-shaped deflectors have different diversion directions. The maximum pressure and temperature in the wedge-shaped deflector are, respectively, 62% and 18% higher than those in the cone-shaped flame deflector. Compared with the wedge-shaped deflector, the cone-shaped deflector has better performance for deflecting, and increased the distance and optimized the structure of the sidewalls can further improve the exhaust gas flow conductivity. The results provide engineering guidance and theoretical significance for design of the flame deflector of the launch platform.The full version of this extended abstract will appear in Defence Technology in 2020.

A recovery system for the key components of the first stage of a heavy launch vehicle

In recent years, the space market has been pushing towards decreasing costs of launching spacecraft by reusing parts of the launchers. The purpose of this article is to present a feasibility study of a recovery system for the engine and engine frame of an existing, expendable heavy launch vehicle and present recommendations for further research.The concept developed is verified based on the Ariane 6. The recovery of the Vulcain Aft Bay (VuAB) is initialised by separation from the first stage at 157.7 km altitude while travelling at 6930 m/s. The study investigates an inflatable aeroshell for protection and deceleration during re-entry, after which it is proposed to further decelerate the VuAB using drogue parachutes. The final part of the concept entails retrieval of the VuAB by a helicopter in mid-air. To enable a controlled gliding flight during retrieval a parafoil is proposed. At launch, the recovery system will weigh 2789 kg with a payload penalty of 720 kg. The system can be integrated into the existing design of the launcher and will not interfere with nominal operations of the launcher. Implementing the recovery system can reduce the cost per launch of an Ariane 6 by 15%

CALCULATION OF SUBOPTIMAL HIGH-ELLIPTICAL ORBIT TO GEOSTATIONARY ORBIT TRAMSFERS FOR SPACECRAFT WITH LOW THRUSTERS

At present the geostationary orbit is where communication satellites are preferably placed. Conventional orbital insertion profiles using chemical propulsion are insufficiently effective and require the use of heavy launch vehicles. Combining electric thrusters with chemical propulsion increases the mass of payload. On the other hand, orbital injection of a spacecraft using electrical propulsion brings up the problem of looking for an optimal control law. The paper discusses the transfer of a spacecraft with low-thrust electrical thruster from a high elliptical orbit to geostationary orbit. It proposes a suboptimal control law for the thrust vector. It provides examples of simulations of the transfer using the control law for various initial conditions. Parameters of intermediate high-elliptical orbits were selected to minimize the time of transfer to the final orbit, and estimates were made of the effects of the residual atmospheric drag during flight in the vicinity of the perigee of the orbit. Considering the low level of error, simplicity and high computational speed the proposed method can be used for trajectory design calculations. Key words: Suboptimal control law, local optimization theory, electric thruster, high-elliptical orbit, geostationary orbit, math model of controlled motion, Pontryagin's maximum principle.

国外重型运载火箭的舱段结构概述

国外重型运载火箭的舱段结构概述