Spacecraft Launch Research Articles

Simulator of pyroshock environment and effect rules of its adjustable parameters

During the launching of spacecraft, the on-board devices will undergo a series of pyroshock environments. In order to verify the reliability of these devices under these pyroshock environments, all of them are needed to take the shock test before launching. This paper has carried out an in-depth research on the simulation method of the pyroshock based on the true explosive excitation. In this study, a simulator containing multiple adjustment parameters is presented and the safety is considered by the design of the protective cover. And the working process of this setup is simulated with the explicit dynamic codes LS-DYNA. What’s more, the effects of the adjustment parameters on the three factors of shock Response Spectrum (SRS) of the resonant board are explored carefully. The rules achieved in this paper are verified by a typical example. The results indicate that the improved simulator can avoid the danger of explosive and make full use of the advantage of actual explosive excitation. And the test condition can be quickly realized at the simulator according to the effect rules of the three adjustable parameters.

Numerical study of ways to prevent side loads in an over–expanded rocket nozzles during the launch stage

The paper aims to study an issues of ensuring the safe launch of spacecraft. The phenomenon of a turbulent boundary layer separation from the rocket engine nozzle wall intensively studied in the overexpansion mode that is realized at the start–up stage at sea–level. The emergence of the so–called restricted flow separation leads to the side loads on the wall of the nozzle, the shaking the structure and the destruction of the engine. The conditions for the realization of restricted or free shock separation in the thrust–optimized contoured nozzles of rocket engines are studied numerically. A hysteresis was obtained when changing the type of separation at the stages of increasing and decreasing pressure at the nozzle inlet with the same ambient pressure. Two ways of controlling the transition from one type of separation to another one were studied numerically. The numerical simulation of separated flow in a model nozzle with a back cone shaped extension was carried out. The numerical results confirm that in the nozzle with the large back cone the restricted shock separation can arise and this result appears to be in agreement with experiment.

Self-Compensation for Disturbances in Differential Vibratory Gyroscope for Space Navigation

In order to meet in many aspects contradictory requirements to gyroscopes used in space applications, this paper proposes to use the multimode Coriolis vibratory gyroscope, which operates in a new differential mode in addition to well-known rate and rate-integrating modes. The differential mode of operation can provide high resistance to external disturbances acting on the gyroscope under operating in launch, docking, and landing regimes of spacecraft. The differential mode of operation differs from the other two known modes by the presence of two, X and Y, measurement channels. Each of the channel measures angle rate with opposite sign. This mode of operation can be implemented by the disposition of standing wave pattern between the drive and sense electrodes. The paper presents promising capabilities of the differential mode of operation to self-compensate for external disturbances. The condition for the standing wave angle disposition that provides maximum compensation for the disturbances is proposed. Numerical and graphical test results on the determination of shock suppression coefficient, vibration sensitivity reduction coefficient, and also reduction of sensitivity to constant and variable magnetic fields are presented.

GNSS potential to monitor unsuccessful spacecraft launches

Using GPS signals we studied TEC disturbances caused by unsuccessful spacecraft launches from Baikonur space center in 1999 and 2006---2016. We considered 15 unsuccessful launches. Six of them were launches into unplanned orbits. Nine launches failed, they ended with the destruction of the launch vehicle and spacecraft during ascent. It is shown that there is a considerable difference in the shapes of TEC disturbances in the ionosphere during launches into unplanned orbits and failed launches. Launches into unplanned orbits caused TEC disturbances in the form of single pulses similar to disturbances during successful launches. Failed launches generated TEC disturbances in the form of wave packets typical of explosions. The revealed effect provides the possibility to determine if the failed launch was followed by the catastrophic destruction of the launch vehicle and spacecraft using data of the ionosphere sounding with GNSS signals.

Project Lyra: Sending a spacecraft to 1I/’Oumuamua (former A/2017 U1), the interstellar asteroid

The first definitely interstellar object 1I/‘Oumuamua (previously A/2017 U1) observed in our solar system provides the opportunity to directly study material from an other star system. Can such objects be intercepted? The challenge of reaching the object within a reasonable timeframe is formidable due to its high heliocentric hyperbolic excess velocity of about 26 km/s; much faster than any vehicle yet launched. This paper presents a high-level analysis of potential near-term options for a mission to 1I/‘Oumuamua and potential similar objects. Reaching 1I/‘Oumuamua via a spacecraft launched in a reasonable timeframe of 5-10 years (launch in 2022-2027) requires an Earth departure hyperbolic excess velocity between 33 and 76 km/s for mission durations between 30 and 5 years, respectively. Different mission durations and their velocity requirements are explored with respect to the launch date, assuming direct impulsive transfer to the intercept trajectory. In addition, missions using a powered Jupiter gravity assist combined with a solar Oberth manoeuvre are explored, using solid rocket engines and Parker Solar Probe heat shield technology. For such a mission, a Falcon Heavy-class launcher would be able to launch a spacecraft of dozens of kilograms towards 1I/‘Oumuamua, if launched in 2021. An additional Saturn gravity assist would allow for the launch of a New Horizons-class spacecraft. Further technology options are outlined, ranging from electric propulsion, and more advanced options such as laser electric propulsion, solar and laser sails. To maximize science return, decelerating the spacecraft at ’Oumuamua is highly desirable, compared to the minimal science return from a flyby. Electric and magnetic sails could be used for this purpose. It is concluded that although reaching the object is challenging, there seem to be feasible options based on current and near-term technology.

HEP quantification strategy based on modified CREAM

The accurate quantification of human error probability (HEP) has long been a main pursuit for most human reliability analysis (HRA) methods. This paper proposes a strategy to analyze and predict human error on the basis of a further modification of the existing cognitive reliability and error analysis method (CREAM) in HRA. Through providing a broader definition for the key parameters used in the quantification method, this paper produces a relatively flexible strategy to determine the nominal HEP (HEP 0 ). Basing on this strategy, the actual quantification of HEP in CREAM is able to be applied to more verified conditions. To prove the validity of the method proposed, a case study of spacecraft launch is also introduced, in which the prediction derived from the method is proved consistent with real field data.

Evaluation of Kp and Dst Predictions Using ACE and DSCOVR Solar Wind Data

AbstractModels that predict the Kp and Dst indices are evaluated using solar wind data at L1. The models consist of ensembles of neural networks that have been developed using ACE Level 2 data from the period 1998–2015. The use of ensembles is motivated by the difficulty of generating functions that generalize well in regions of the input‐output space that are poorly sampled, which typically occurs during stronger events. Since the launch of the DSCOVR spacecraft, providing measurements from about August 2016, new and independent data have become available to test the models. ACE Level 2 data for almost the same period are also available, representing another independent set collected after 2015. We evaluate the models using plots and various statistical measures. We also study the performance of the predictions for lead times up to 3 hr. The results show that the models perform better when using ACE or cleaned DSCOVR data compared to the real‐time DSCOVR data and that lead times of L1‐Earth travel time plus a maximum of 1 hr are possible. As the models use only solar wind for their inputs, and the temporal dynamics of Kp and Dst are very different, we see significant differences in the error distributions that we believe are related to long‐term changes in Dst that are not captured by the Dst prediction model.

ЭКСПЕРИМЕНТАЛЬНОЕ ПОДТВЕРЖДЕНИЕ РАБОТОСПОСОБНОСТИ КАПИЛЛЯРНОГО ЗАБОРНОГО УСТРОЙСТВА (СЕТЧАТОГО РАЗДЕЛИТЕЛЯ) ПРИ ПРОГРАММНОМ РАЗВОРОТЕ

One of the problems of providing a group launch of spacecraft is to ensure the re-launch of the propulsion engine for the transition to another orbit, after the separation of the first vehicle. When the spacecraft is separated into a stage, an acceleration is applied directed toward the bottom of the tank and leading to an outflow of the fuel component from the mesh separator. This, in turn, leads to the denudation of the mesh separator and the penetration of gas under it, which can lead to a breakdown of the launch of the engine. In viewof the complexity of the hydrodynamic processes taking place in the tanks, experimental confirmation of the mesh separators chosen during the design is required.In this paper, we describe the methods for determining the model conditions (the choice of the scale of the experimental design, the type of the model fluid, the magnitude of the accelerations, etc.), a description of the stand designed specifically for the test data, and the results of the experimental confirmation of the capillary sampling device performance in the separation of the spacecraft. The results of the performed works confirm the correctness of the decisions made in the development of capillary sampling devices.

Defending the earth from long-period comets and sneaky asteroids: Short term threat response requires long term preparation

Comets, such as the recent Comet C/2013 A1 (Siding Spring) can impact terrestrial planets with very short warning times: for Siding Spring the time from discovery to closest approach to Mars (135,000 km) was less than 22 months. Short warning times are also possible for certain asteroids approaching on unfavorable orbits, e.g., from the direction of the Sun. The time required to design, build, test and launch a high-reliability Interceptor spacecraft is on the order of 5 years, a timescale incompatible with threat mitigation for such short warning times. A potential solution is to build the interceptor before the threat is detected and put it into storage until it is required. Launch of a stored spacecraft could be accomplished in much less than a year after authorization based on experience with the DSCOVR (Deep Space Climate Observatory) program managed by NASA's Goddard Space Flight Center.While asteroidal impacts are of order 100 times more likely than cometary impacts, comet impacts can carry more than 100 times the energy of a typical asteroid threat, making their destructive power nearly equal. Cometary threats have been largely ignored to date, a situation that needs to change as our Planetary Defense efforts become more mature. An idealized threat response system will be described.

Nonlinear Feed Forward Control of a Perturbed Satellite using Extended Least Squares Adaptation and a Luenberger Observer

Three adaptive approaches for a non-linear feed-forward controller are combined with two physics-inspired sinusoidal trajectory planners in a spacecraft attitude control model for large slew maneuvers. The basis of the model is a space based satellite sensor which has suffered an unwanted collision where the inertial matrix of the craft is no longer similar to the originally measured inertial matrix. This causes a large inherent error in the feedforward control needed for system maneuvers due to the mismatch of expected dynamics. Trajectory generation, feedforward control, feedback control, filters, observers, and system stability are discussed in relation to the non-linear dynamics under simulation. The adaptive feed-forward controllers discussed include a proportional-derivative (PD) adaptive controller, a Recursive Least Square (RLS) Method, and an Extended Least Squares (ELS) Method. Mean control effort stayed relatively constant between configurations. The controller configuration with ELS feedforward, PID feed-back, and an extended sinusoidal trajectory outperformed the baseline adaptive controller. Mean error was decreased by 23.4%, error standard deviation by 34.0%, and maximum error by 33.0% from a similar case using RLS adaptation. This improvement is entirely based on a need to correct for un-modeled or mis-modeled dynamics. This scenario occurs in actual operation during spacecraft launch, collisions with debris, or can be caused by fuel slosh or loose components.

Stellar kinematics in the IC 348 cluster

In 30 - 40s of the last century Viktor Ambartsumian has published several studies on the kinematics and dynamics of stellar systems. The scientific issues of these papers not only remain relevant up to nowadays, but also are of key importance for the construction of the generalized theory of star formation. The launch of the Gaia spacecraft on December 19, 2013, opens up new horizons in the study of young stellar systems origin and evolution. The accuracy of astrometric measurements (the typical uncertainty is about 0.04 - 0.7 mas for the positions and parallaxes, and about 0.05 - 1.2 mas yr-1 for the proper motions), undoubtedly, will allow lifting the possibilities of study the kinematics and dynamics of star-forming regions to a new level. We focus our present research on the young (∼ 2 - 3Myr) stellar cluster IC 348. It has a complex structure, which includes the older compact core and the younger subgroup in the vicinity of HH 211 outflow. Based on the Gaia DR2 data, we considered the kinematic properties of both the cluster as a whole and its subgroups. We found that special velocity dispersion significantly exceeds the virial velocity dispersion and, therefore, the IC 348 stellar cluster is a supervirial or gravitationally unbound system.

Deflection of fictitious asteroid 2017 PDC: Ion beam vs. kinetic impactor

Mission scenarios for the deflection of fictitious asteroid 2017 PDC are investigated. Two deflection options, kinetic impactor (KI) and ion beam shepherd (IBS), are studied and compared on the basis of deflection performance, safety, as well as mission schedule and political aspects. Firstly, we propose the launch of a medium-size rendevous spacecraft equipped with at least two ionic thrusters that can serve as propulsion means for the interplanetary trajectory up to rendezvous with the asteroid and as contactless actuators for a possible follow-up deflection mission. The asteroid, whose uncertainty ellipsoid is initially too large to establish whether (and how) it should be deflected, is reached by the rendezvous spacecraft after a low-thrust interplanetary trajectory of reasonable duration. Following rendezvous the spacecraft is placed in the vicinity of the asteroid to estimate its mass, study its structure and composition and, crucially, reduce its uncertainty ellipsoid by ground tracking to confirm or rule out an impact. Assuming that an impact is confirmed two main deflection scenarios are considered based on the actual asteroid size. Ion beam deflection is considered with the possibility of full deflection (the asteroid misses the Earth by a safe margin) or impact location adjustment (the impact footprint is diplaced to the nearest unpopulated region) depending on the asteroid size and the predicted impact location. The launch of a kinetic impactor mission is also considered with the employment of the rendezvous spacecraft to measure the deflection outcome and possibly to refine the deflection in case it is needed. The deflection performance of the two methods is compared.

Options and uncertainties in planetary defense: Mission planning and vehicle design for flexible response

This paper is part of an integrated study by NASA and the NNSA to quantitatively understand the response timeframe should a threatening Earth-impacting near-Earth object (NEO) be identified. The two realistic responses considered are the use of a spacecraft functioning as either a kinetic impactor or a nuclear explosive carrier to deflect the approaching NEO. The choice depends on the NEO size and mass, the available response time prior to Earth impact, and the various uncertainties. Whenever practical, the kinetic impactor is the preferred approach, but various factors, such as large uncertainties or short available response time, reduce the kinetic impactor's suitability and, ultimately, eliminate its sufficiency.Herein we examine response time and the activities that occur between the time when an NEO is recognized as being a sufficient threat to require a deflection and the time when the deflection impulse is applied to the NEO. To use a kinetic impactor for successful deflection of an NEO, it is essential to minimize the reaction time and maximize the time available for the impulse delivered to the NEO by the kinetic impactor to integrate forward in time to the eventual deflection of the NEO away from Earth impact.To shorten the response time, we develop tools to survey the profile of needed spacecraft launches and the possible mission payloads. We further present a vehicle design capable of either serving as a kinetic impactor, or, if the need arises, serving as a system to transport a nuclear explosive to the NEO. These results are generated by analyzing a specific case study in which the simulated Earth-impacting NEO is modeled very closely after the real NEO known as 101955 Bennu (1999 RQ36). Bennu was selected for our case study in part because it is the best-studied of the known NEOs. It is also the destination of NASA's OSIRIS-REx sample return mission, which is, at the time of this writing, enroute to Bennu following a September 2016 launch.

EXPERIMENTAL CONFIRMATION OF THE ACQUISITION PROPELLANT DEVICE (MESH SEPARATOR) OPERABILITY IN THE SPACECRAFT SEPARATION

One of the problems of providing a group launch of spacecraft is to ensure the re-launch of the propulsion engine for the transition to another orbit, after the separation of the first vehicle. When the spacecraft is separated, a stage is acted upon by the acceleration directed toward the bottom of the tank and leading to the outflow of the fuel component from the mesh separator. This leads to the mesh separator denudation and the gas penetration under it, which can lead to a breakdown of the engine launch. In view of the hydrodynamic processes complexity taking place in the tanks, experimental confirmation of the mesh separators chosen during the design is required.The methods for determining the model conditions (the choice of the scale of the experimental design, the type of the model fluid, the magnitude of the accelerations, etc.) are presented, a description of the stand designed specifically for the test data and the results of the experimental confirmation of the capillary sampling device performance in the separation of the spacecraft are given. The results of the performed work confirm the correctness of the decisions made in the development of capillary sampling devices.

A Viable Commercial Model Dealing with Space Debris

The problem of space debris has already been discussed for many years. How to deal with debris and create a viable commercial opportunity for removing debris has become a universal concern nowadays. First of all, our team builds a global fund model that offers a commercial opportunity and assists with orbital debris removal. The fund could be established basing on the simple economic approach to address the problem of space debris. Fund could be collected from any countries involved with the spacecraft launch, prior to all launches for 25 years. Five percentage of the total cost of various space-related missions, which include space activities both in the past and future, are given to the foundation. Approximately half of the payments into the fund would be distributed to those entities devoted in removing space debris, including private firms interested in entering the debris removal market. An attractive market would exist with the policy’s support, assuming that the orbital debris removal could be accomplished with low risk technically. Secondly, the situation about the main countries’ space-related missions is analyzed, and a total cost of each country spent on spaceflight in the past years is estimated. Utilizing the estimated data, the amount of credit that private firm gets from the fund could be derived. At the same time, we create a time-dependent function to predict the budgets for launching satellites in the coming years. Thirdly, the comparison of the different options of removing debris is done and our team chooses the proper methods for cleaning debris objects in varied size according to existed technologies. Right after that the cost taken to do the removal job is derived. Next, the risk is analyzed and the benefit could be predicted. Our model’s dependence on the policy or objective environment makes our result full of uncertainty. Finally, we come to a conclusion that by using our model there would be an economically attractive opportunity for removing space debris with the support of relative policy or international organizations.

Need for space governance for India and global space governance

Comets, such as the recent Comet C/2013 A1 (Siding Spring) can impact terrestrial planets with very short warning times: for Siding Spring the time from discovery to closest approach to Mars (135,000 km) was less than 22 months. Short warning times are also possible for certain asteroids approaching on unfavorable orbits, e.g., from the direction of the Sun. The time required to design, build, test and launch a high-reliability Interceptor spacecraft is on the order of 5 years, a timescale incompatible with threat mitigation for such short warning times. A potential solution is to build the interceptor before the threat is detected and put it into storage until it is required. Launch of a stored spacecraft could be accomplished in much less than a year after authorization based on experience with the DSCOVR (Deep Space Climate Observatory) program managed by NASA's Goddard Space Flight Center.While asteroidal impacts are of order 100 times more likely than cometary impacts, comet impacts can carry more than 100 times the energy of a typical asteroid threat, making their destructive power nearly equal. Cometary threats have been largely ignored to date, a situation that needs to change as our Planetary Defense efforts become more mature. An idealized threat response system will be described.

Loss Modelling and Experimental Verification of A 98.8% Efficiency Bidirectional Isolated DC-DC Converter

High efficiency in power converters means less power wasted, which implies reduction in heat sink requirement, thereby small in size and less in weight. This power loss reductions in space application leads to smaller solar panels and less fuel on-board, hence reduces the spacecraft size, weight and launch cost. One of the major challenges for power converters in space application is to boost their efficiency, to reduce thermal dissipation problems.In this paper, design and implementation of an ultra-high efficiency isolated bi-directional dc-dc converter utilizing GaN devices is presented. Loss modelling of the GaN converter is also included in this paper. The converter has achieved a maximum measured efficiency of 98.8% in both directions of power flow, using the same power components. Hardware prototype of the converter along with the measured efficiency curve is also presented in this paper.

Gravitational and Anti-Gravitational Applications

Because gravitation operates by a flow from the attracting object acting on the attracted object as shown in the prior papers Connecting Newton’s G With the Rest of Physics-Modern Newtonian Gravitation Resolving the Problem of “Big G’s” Value and The Mechanics of Gravitation-What It Is; How It Operates, it is now possible to partially deflect gravitation away from an object so that the gravitational attraction on the object is reduced. That effect makes it possible to extract energy from the gravitational field, which makes the generation of gravito-electric power technologically feasible. Such plants would be similar to hydro-electric plants and would have their advantages of not needing fuel and not polluting the environment. Physically, the action of deflecting away gravitational attraction, which of course is directed toward the gravitation source, produces an equal but opposite reaction on the mechanism that produces the deflection of the gravitational action [the deflector], a reaction directed away from the gravitation source. The result is the combination of reducing the gravitational attractive acceleration of the object toward the gravitation source plus the introducing of a reactive acceleration on the object in the direction away from the gravitation source. Such a deflector, engineered to enable controlled adjustment of the amount and the direction of its action, could provide for a spacecraft both launch levitation and deep space travel acceleration. It could provide both levitation and horizontal motion for a flying vehicle over a planet surface. This technology, which uses readily abundantly available materials and techniques, is ready now for research and engineering refinement.

The Capturing of Space Debris with a Spaceborne Multi-fingered Gripper

With the massive launching of spacecraft, more and more space debris are making the low Earth orbit (LEO) much more crowded which seriously affects the normal flight of other spacecrafts. Space debris removal has become a very urgent issue concerned by numerous countries. In this paper, using SwissCube as a target, the capturing of space debris with a spaceborne four-fingered gripper was studied in order to obtain the key factors that affect the capturing effect. The contact state between the gripper fingers and SwissCube was described using a defined contact matrix. The law of momentum conservation was used to model the motion variations of the gripper and SwissCube before and after the capturing process. A zero-gravity simulation environment was built using ADAMS software. Two typical kinds of capturing processes were simulated considering different stiffness of fingers and different friction conditions between fingers and SwissCube. Comparisons between results obtained with the law of momentum conservation and those from ADAMS simulation show that the theoretical calculations and simulation results are consistent. In addition, through analyzing the capturing process, a valuable finding was obtained that the contact friction and finger flexibility are two very important factors that affect the capturing result.

低成本运载火箭发展与分析

为适应未来军事和商业航天低成本发射需求，世界各航天机构纷纷提出各自的低成本运载火箭设计方案，并采取各种技术措施大幅降低发射成本。对各主要方案和技术进行对比，从美国某型运载火箭各组成部分与其质量和成本估计、猎鹰9R可重复使用运载火箭运载能力损失和总低成本估计两个方面分析了运载火箭成本与性能的关系，并为我国低成本运载火箭研制提出相应发展建议。 By the low cost requirement in military and commerce spacecraft launch tasks, many space institutions in the world have proposed multifarious design schemes of low cost launch vehicles, and all kinds of new technologies and measures have been adapted to reduce the total launch cost. The paper researched the major design schemes and key technologies comparatively. Then the relationship between of cost and performance of the launch vehicles are independently analyzed from mass percents and cost percents of the major components for a certain American launch vehicle and carrying capability loss and total low cost for a reusable launch vehicle. At last, several proposals were presented for the development of low cost launch vehicles in China.