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Copyright of Journal of National University of Defense Technology / Guofang Keji Daxue Xuebao is the property of NUDT Press and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

Based on the goals of "high reliability, high frequency, rapid launch, and low cost" for space launch sites, an integrated dual-sided deflector system for convective cooling and thermal protection is presented. The interaction process between the gas jet and liquid water jet and its effect on the flow field environment are thoroughly studied using numerical calculation methods. Furthermore, considering the phase-change heat transfer issue in a compressible gas–liquid two-phase flow, and the varying distribution of different bubble shapes and sizes at the gas–liquid interface, a modified Lee model is derived. The research results show that compared to the classical Lee model, the modified Lee model can achieve a higher numerical accuracy in predicting the heat and mass transfer processes in gas–liquid two-phase flows. Through comparative analysis with the traditional dual-sided deflector and the conventional cooling system, the integrated dual-sided deflector system exhibits significant performance advantages in gas flow regulation and flow field environment improvement at the near-ground region of the space launch site. It not only achieves effective flow deflection, but also mitigates the degree of erosion caused by the gas jet on the deflector. This conclusion can provide theoretical references for the thermal protection design of commercial launch vehicle systems at space launch sites. [ABSTRACT FROM AUTHOR]

This paper presents a real-time fault-tolerant guidance method, which solves the autonomous rescue problem in the event of thrust drop failure during the final stage flight of launch vehicles. The core of the method is the online and optimal reconstruction of both the degraded target orbit and the subsequent flight trajectory. Aiming at achieving satisfied onboard computing performance and making the best use of the launcher's residual energy, a unique relaxation-penalization model transformation method for the nonlinear orbit injection conditions is proposed, and a successive convexification algorithm is developed to solve the problem. Furthermore, both the two classic thrust drop failure situations, that is the mass flow rate drop and the specific impulse drop problems, can be resolved by the proposed algorithm. The average run-time of this orbit-trajectory joint optimization algorithm is below 150ms in an embedded ARM processor @1.5 GHz, which is adequate for onboard use. Using the above algorithm as a central infrastructure, a receding-horizon-strategy-based fault-tolerant guidance method is proposed. By performing the optimization repeatedly at a high frequency, the effects of parameter deviations and external disturbances can be absorbed for the most part, for example, the guidance algorithm is verified to be capable of tolerating at least ±5% thrust deviations. Comprehensive numerical and hardware-in-the-loop simulations are performed to demonstrate the computational performance, accuracy, and reliability of the proposed algorithms. [ABSTRACT FROM AUTHOR]

Copyright of Computer Measurement & Control is the property of Magazine Agency of Computer Measurement & Control and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

Copyright of Computer Measurement & Control is the property of Magazine Agency of Computer Measurement & Control and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

The object of this study was the motion of an ultralight class variable-length launch vehicle made of a polymer body along the active phase of the trajectory. The work considers the solution to the problem of designing low-cost means of delivery to orbit, namely to the assessment of the possibility of removing the payload by a carrier rocket with a polymer body of variable length beyond the dense atmosphere of the Earth. For this purpose, ballistic projection of the trajectory of the launch vehicle was carried out taking into account overloading; its aerodynamic characteristics and peculiarities of aerothermodynamic processes occurring during flight in the atmospheric phase of the trajectory were determined. The closeness (up to 10 %) of the obtained results with known experimental data is shown. The influence of the aerodynamic force on the parameters of the launch vehicle motion was studied. A flight simulation was conducted, the results of which showed the fundamental possibility of launching a CubeSat 24U class payload using a launch vehicle with a polymer body of variable length to a suborbital trajectory with an altitude of about 300 km. At the same time, the effective longitudinal overload on the body of the launch vehicle does not exceed 4 units, and the temperature on the surface of the body does not exceed 300 K. A feature of the research is the use of a multidisciplinary approach, which implies taking into account the interrelationship of aerodynamic, thermodynamic, and ballistic processes. The established motion parameters, aerodynamic characteristics, and the surface heating temperature of the launch vehicle body are key values for further research on the design and analysis of a launch vehicle with a polymer body of variable length. These data could be used to calculate the mechanical and thermal loads acting on the structure of the launch vehicle during flight. [ABSTRACT FROM AUTHOR]

When optimizing the staging of a launch vehicle, usually a reference mission with a fixed payload mass is used to determine the specific launch cost (or a surrogate thereof, e.g. total mass or dry mass). While this approach neglects suboptimal loading, it delivers sufficiently good results for the evaluation of an individual launcher if the design mission is well known. However, this approach does not capture the full range of payloads that a launch vehicle is expected to launch during its operational lifetime. In order to estimate the cost and subsequently optimize the staging of such launch vehicle families, a full launch scenario has to be considered. The cost of an individual launch vehicle loses importance and the deciding factor is the total cost associated with launching the specified scenario. However, estimates of the future market inevitably include large uncertainties. These uncertainties and their impact are ideally considered and quantified when assessing the specific recurring cost of a single launch vehicle. The consideration of a whole launch market scenario is especially relevant when evaluating a family of launch vehicles, an option currently being discussed for future European launchers. In this case, the distribution of the payloads onto the different members of the launch vehicle family has to be optimized in order to reach the lowest possible cost. Within this manuscript, a method that allows the determination of the recurring cost of such a launcher family based on a launch market scenario including uncertainties is proposed. Based on the launch numbers derived from the combinatorial optimization, parametric cost models for recurring and non-recurring cost are adapted and applied to three possible future launch vehicle families. • Novel launch vehicle family cost assessment method based on market scenarios. • Efficient payload assignment minimizes recurring costs. • Future market uncertainties considered using Monte Carlo method. • Methodology tested on three conceptual launch vehicle families. [ABSTRACT FROM AUTHOR]

Copyright of Journal of National University of Defense Technology / Guofang Keji Daxue Xuebao is the property of NUDT Press and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

Copyright of Journal of National University of Defense Technology / Guofang Keji Daxue Xuebao is the property of NUDT Press and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

The Pogo suppression method can be divided into active suppression and passive suppression. For practical engineering, passive suppression is generally achieved through accumulators, but this method sometimes cannot meet the design requirements of liquid rockets with uncertain model parameters and time-varying parameters. As for active suppression methods, state variables are mostly used as feedback thus leading to poor practical feasibility. In this study, an adaptive Pogo active suppression controller design method is proposed where measured acceleration is used as feedback. The eigenspace transformation theory is utilized to design two types of dimensional reduced models: an observation-oriented dimensional reduced model (O-ODRM) and a control-oriented dimensional reduced model(C-ODRM). The former works for observer design, and its state variables can be reconstructed by measuring only the acceleration and can be input to the adaptive controller after matrix operations. The latter is employed for adaptive controller design, and it can effectively process the issues of model parameter uncertainty and time-varying parameters. To demonstrate the effectiveness of the proposed method, a certain type of rocket was selected as the research object for simulation analysis. The simulation results show that the proposed method, which uses measured acceleration as feedback, can achieve Pogo active suppression for liquid rockets with uncertain and time-varying parameters. • An active Pogo suppression method based on acceleration feedback considering practical applications has been proposed. • The proposed method effectively solves the problem of difficult controller design for Pogo systems by measuring structural acceleration. • The proposed dimensionality reduction models for observation and control are the basis for controller design. • The proposed method has excellent adaptability to Pogo systems with parameter uncertainty and time-variation. [ABSTRACT FROM AUTHOR]

Copyright of Systems Engineering & Electronics is the property of Journal of Systems Engineering & Electronics Editorial Department and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

This paper presents a simplified method for launch vehicle coupled loads analysis (CLA) with spacecraft which is normally conducted in the process of the spacecraft development. The method allows spacecraft developers to calculate coupled vibration response of their own without performing transient analysis for the coupled whole system model. A new method, which allows us to obtain coupled vibration response by convolving non-loaded acceleration time history at the spacecraft boundary with corresponding impulse response of combined apparent mass matrices in frequency domain, has been proposed. The proposed method functions well in any CLA cases including lift-off forced events where loads accompany quasi-static component in nature. We propose a novel approach to escape the error caused in the process of convolution for the quasi-static event by making use of the method to convolve only dynamic acceleration response extracted from the original acceleration by inertia relief technique. Finally, we demonstrate the validity of the proposed method through the analysis using practical launch vehicle and spacecraft models.

With the development of new high-thrust rockets, the decreasing frequency of the bending modes of the rocket, as well as the factors such as flow separation and shock oscillations generated during transonic flight make it more prone to non-constant vibration. In this paper, a launch vehicle model with boost was taken as the research object, and the aerodynamic damping of the rocket during forced vibration was obtained through numerical calculation. The factors affecting the aerodynamic damping of the rocket were analyzed, including the position of structural nodes, the magnitude of vibration amplitude, pulsating pressure and so on. The study shows that the boost mainly plays the role of increasing aerodynamic damping and the front node mainly affects the aerodynamic damping of the contraction section. The vibration amplitude size and the pulsating pressure have a negligible effect on the aerodynamic damping.

A launch vehicle needs to adapt to a complex flight environment during flight, and traditional guidance and control algorithms can hardly deal with multi-factor uncertainties due to the high dependency on control models. To solve this problem, this paper designs a new intelligent flight control method for a rocket based on the deep reinforcement learning algorithm driven by knowledge and data. In this process, the Markov decision process of the rocket landing section is established by designing a reinforcement function with consideration of the combination effect on the return of the terminal constraint of the launch vehicle and the cumulative return of the flight process of the rocket. Meanwhile, to improve the training speed of the landing process of the launch vehicle and to enhance the generalization ability of the model, the strategic neural network model is obtained and trained via the form of a long short-term memory (LSTM) network combined with a full connection layer as a landing guidance strategy network. The proximal policy optimization (PPO) is the training algorithm of reinforcement learning network parameters combined with behavioral cloning (BC) as the reinforcement learning pre-training imitation learning algorithm. Notably, the rocket-borne environment is transplanted to the Nvidia Jetson TX2 embedded platform for the comparative testing and verification of this intelligent model, which is then used to generate real-time control commands for guiding the actual flying and landing process of the rocket. Further, comparisons of the results obtained from convex landing optimization and the proposed method in this work are performed to prove the effectiveness of this proposed method. The simulation results show that the intelligent control method in this work can meet the landing accuracy requirements of the launch vehicle with a fast convergence speed of 84 steps, and the decision time is only 2.5 ms. Additionally, it has the ability of online autonomous decision making as deployed on the embedded platform. [ABSTRACT FROM AUTHOR]

Based on the goals of “high reliability, high frequency, rapid launch, and low cost” for space launch sites, an integrated dual-sided deflector system for convective cooling and thermal protection is presented. The interaction process between the gas jet and liquid water jet and its effect on the flow field environment are thoroughly studied using numerical calculation methods. Furthermore, considering the phase-change heat transfer issue in a compressible gas–liquid two-phase flow, and the varying distribution of different bubble shapes and sizes at the gas–liquid interface, a modified Lee model is derived. The research results show that compared to the classical Lee model, the modified Lee model can achieve a higher numerical accuracy in predicting the heat and mass transfer processes in gas–liquid two-phase flows. Through comparative analysis with the traditional dual-sided deflector and the conventional cooling system, the integrated dual-sided deflector system exhibits significant performance advantages in gas flow regulation and flow field environment improvement at the near-ground region of the space launch site. It not only achieves effective flow deflection, but also mitigates the degree of erosion caused by the gas jet on the deflector. This conclusion can provide theoretical references for the thermal protection design of commercial launch vehicle systems at space launch sites.

In this article, guidance and control of the reusable launch vehicle at pre-landing phase have been studied. A plenty of vertical take-off and vertical landing (VTVL) reusable launch vehicle have been gradually developed to access the space allowing for reducing the flight operation costs and realizing reusability such as Falcon 9. The goal of this article is increasing the accuracy and safety of the guidance and control system of the launch vehicle in order to successfully land on the ground station. Integrated guidance and control, being a novel method in guidance and control, is used to guide the launch vehicle in the re-entry phase and landing phase on the ground. The returning object studied is a Falcon 9 rocket for which we find the 6 DOF model. This vehicle has a vertical landing phase and implementing an integrated guidance and control system is a novel approach in this area. The controller used in this article is based on sliding-mode. This controller has an acceptable performance against the uncertainties and indefinite values of the system parameters. GPOPS software is used to create the initial flight path of the launch vehicle. This software calculates the optimal landing path with respect to dynamic constraints, control constraints and a cost function. Genetic algorithm is used to find the constant parameters for guidance and control. In this work, a fast terminal sliding mode control is proposed for a VTVL reusable launch in the present of disturbance and uncertainties. Compared with the previous works, the proposed controller has satisfactory performances and achieves smaller steady-state error under the consideration of multiple simulation results. Results show that integrated guidance and control needs less bulk standard deviation than common guidance and control due to simultaneous solving of guidance and control equations. This causes the defined constraints not to surpass the acceptable values anywhere in the path.

The hard landing of the spent stage of promising launch vehicles (LV) in the designated areas of the fall leads to the collapse of the structure, the spillage of the remnants of the guaranteed fuel reserve, the outbreak of fires, and, as a result, require large expenditures for the complete elimination of man-made consequences. Residues of the guaranteed fuel are additional stocks of propellant components in the rocket blocks of the stage of LV, designed to compensate for the disturbing factors acting in flight on the LV. They are characterized by carcinogenic and mutagenic effects on the environment. The article considers the issue of extraction of unprocessed residues of guaranteed fuel. An experimental stand and a program for conducting experiments have been developed to study the extraction process. The purpose of the study is to conduct a comparative analysis of the results of mathematical modeling and physical modeling. Based on the results of the physical experiment, a mathematical model, and the results of a comparative study of two types of models are presented: The calculated heat carrier (HC) temperatures at the experimental model setup (EMS) inlet (according to the first law of thermodynamics and the Navier-Stokes equations) and the measured HC temperature at the EMS inlet during experiments were compared.

Radar Cross Section (RCS) plays a significant role in detecting and tracking the space-based objects such as launch vehicles, missiles, aircrafts etc. In space applications, Radar systems are used to track and provide real-time trajectory information of the satellite launch vehicles after the lift off from the launch pad for range safety purpose. RCS is a critical key parameter that determines tracking performance of the Radar and it is highly dependent on both Radar operating parameters and the target characteristics. For space-based applications, a good quantity of RCS is required for quick detection by the Radar for continuous tracking. In order to choose the best Radar tracking configuration for real time tracking of the launch vehicle, it is required to model and simulate the launch vehicle’s RCS fluctuations prior to launch in order to predict the real time Signal to Noise Ratio (SNR) for its complete dynamic trajectory. This modeling and simulation methodology will help to choose the optimum Radar configuration for obtaining a good quantity SNR in the real-time launch. This study also provides good guidance to Radar operators for the effective Radar operation during real time space object tracking. This paper demonstrates, the real-time RCS fluctuations of a typical ISRO launch vehicle through simulation for its dynamic trajectory using physical optics based EM software prior to launch. Furthermore, the simulation results are validated with real time monostatic Radar tracking data, which showed good agreement. [ABSTRACT FROM AUTHOR]

The hard landing of the spent stage of promising launch vehicles (LV) in the designated areas of the fall leads to the collapse of the structure, the spillage of the remnants of the guaranteed fuel reserve, the outbreak of fires, and, as a result, require large expenditures for the complete elimination of man-made consequences. Residues of the guaranteed fuel are additional stocks of propellant components in the rocket blocks of the stage of LV, designed to compensate for the disturbing factors acting in flight on the LV. They are characterized by carcinogenic and mutagenic effects on the environment. The article considers the issue of extraction of unprocessed residues of guaranteed fuel. An experimental stand and a program for conducting experiments have been developed to study the extraction process. The purpose of the study is to conduct a comparative analysis of the results of mathematical modeling and physical modeling. Based on the results of the physical experiment, a mathematical model, and the results of a comparative study of two types of models are presented: The calculated heat carrier (HC) temperatures at the experimental model setup (EMS) inlet (according to the first law of thermodynamics and the Navier-Stokes equations) and the measured HC temperature at the EMS inlet during experiments were compared. [ABSTRACT FROM AUTHOR]

Copyright of Computer Measurement & Control is the property of Magazine Agency of Computer Measurement & Control and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

Sixty percent of the failures of launch vehicles in the ascending phase occur in the propulsion system. Among them, the vibration generated by the engine is an important factor in the occurrence of failure. At present, health assessment methods in the aerospace field are mostly for specific equipment, and scholars mostly assess the real-time health status of launch vehicle engines which can only reflect the current health status of the launch vehicle. Existing methods cannot be applied to different equipment, and there is a lack of research on health assessments of fuzzy and complex mechanical systems. In this article, we propose a multi-layer and multi-factor predictive evaluation method for a fuzzy and complex system and conduct experiments on real vibration data of rockets. First, we divide the health assessment level according to the vibration data that affect the normal operation of the rocket. Secondly, we obtain the future trend of vibration signals based on five data prediction methods and calculate the health status interval of the rocket engine's working conditions based on the boxplot method. At the same time, we calculate the single health evaluation set of every vibration signal. We obtain the weights of each level and factor for the health value based on an analytic hierarchy process (AHP). The optimization of this step avoids an over-reliance on expert experience. Finally, we complete a fuzzy comprehensive evaluation of the engine system from the bottom up to obtain the final health value. The minimum evaluation error is 0.0193% on the test data of the Long March series launch vehicle engine, which shows that the proposed method can successfully predict and evaluate the launch vehicle engine. [ABSTRACT FROM AUTHOR]

In order to better solve the problem of Pogo stability considering parameter uncertainty and time-variation in the large launch vehicle, the model-reference adaptive control based on the dimensionality reduction model (MRAC-DRM) is proposed. Firstly, the state space Pogo model with an active controller is established for control system design by the improved Rubin method. Secondly, a greatly reduced dimensional model of the Pogo control system is derived from the eigenvalue transformation theory, which preserves the features of Pogo instability and is appropriate for control design. Then, the reference model and adaptive control law are designed by the model-reference adaptive control method based on the reduced dimensional model. Finally, two examples with different types of propulsion systems are used to demonstrate the effectiveness of the proposed MRAC-DRM. Simulation results demonstrate that the proposed method not only has excellent adaptability to the Pogo systems with parameter uncertainty and time-variation but also is insensitive to the initial value of feedback gain. • The MRAC-DRM method is proposed for Pogo systems with parameter uncertainty and time-variation. • The proposed method can effectively suppress Pogo vibration of different types of launch vehicles. • The proposed method is robust as it is insensitive to the initial value of feedback gain. • A reduced-dimension Pogo model is derived, which plays an important role in the controler design. [ABSTRACT FROM AUTHOR]

The article aims at reviewing the drop regions (DR) of the launch vehicles (LV) separating parts (SP) and methods of their determination. The DRs include sea and land areas; going beyond them is associated with a number of environmental, economic, and political factors. Their combination dictates the need to ensure the safety of the people, transport, infrastructure, and environment from the negative impact of LV SPs and fuel residues. The Monte Carlo method is mostly used to determine the impact areas. It enables an estimation of the probability of the SPs of LVs falling in certain areas, constituting the DRs. These points are varied according to a set of different initial parameters. The methods of controlling the impact areas are contingently divided into engineering (based on a change in the design appearance of the LV), mathematical (which includes the changes in or optimization of the LV's trajectory or its SP), and "soft landing" (implying the return of the LV's to the spaceport or to a certain prepared area). The present analysis can be used as a starting point when choosing a method for determining and controlling the projected LVs and the SPs' area of impact. [ABSTRACT FROM AUTHOR]

The paper describes how the attitude of a sounding rocket, launch vehicle or satellite with respect to the Earth can be estimated from camera images of the Earth horizon. Details about detecting the horizon in the camera image, fitting hyperbolae or ellipses to the detected horizon curve and deriving the Earth nadir vector and the corresponding error covariance from the fitted conic section are given. The presented method works at low heights, where the projected horizon mostly appears to be hyperbolic, as well as at large heights, where the projected horizon mostly appears to be elliptic and it is irrelevant if the Earth is fully or only partially in the field of view of the camera. The method can be universally used to estimate the direction vectors and attitude with respect to any spherical celestial body such as the Sun or Moon. Using the example of a sounding rocket mission with two cameras aboard, it is illustrated how the estimates of the Earth nadir and the Sun direction vectors are fused with the measurements of a strapdown inertial measurement unit and a GPS receiver to obtain an accurate and continuous estimate of the three-dimensional orientation of the sounding rocket with respect to the Earth. [ABSTRACT FROM AUTHOR]

Copyright of Computer Measurement & Control is the property of Magazine Agency of Computer Measurement & Control and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

Generally, the guidance and the control are designed as separate modules, which are suitable for many common and normal applications, but in order to achieve less error in deploying the satellite into the orbit increase the accuracy in performing the mission, we need to develop integrated Guidance and Control algorithms (IGC). By using such algorithms, it is possible to design the closed loop guidance for the last stage of the launch vehicle in vacuum condition. Examples of this are iterative guidance mode (IGM) and power explicit guidance (PEG). In this paper, the guidance of the final phase along with the guidance of the first and second stages of a typical solid rocket launch vehicle is modeled and simulated. The results of this article show that the IGM guidance can be well compatible with the Pitch Program guidance of the previous steps and successfully complete the deployment in the orbit. [ABSTRACT FROM AUTHOR]

Copyright of Systems Engineering & Electronics is the property of Journal of Systems Engineering & Electronics Editorial Department and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

The manufacture of aerospace structures is often accomplished by machining solid plates or forgings, sourced to fulfil the envelope of each component. As such, any external features such as reinforcement ribs, bosses or mounting brackets expand the envelope of the structure, significantly increasing the raw material costs. The present investigation assessed the feasibility of using friction welding to join such features on large spacecraft structures. Two Linear Friction Welding (LFW) case studies were considered: (i) Joining bosses for line attachments on cryogenic tanks in AA2195-T8 for heavy launch vehicles; and (ii) Adding connection tabs to Ti-6Al-4V fuel tanks for the Eurostar Neo satellite. A series of LFW trials was undertaken to demonstrate the weldability of the materials, initially using representative flat-plate coupons, which were characterised via metallography, hardness and tensile testing. The weld properties reported met the acceptance criteria defined by the end-users, Airbus UK and ArianeGroup. Representative technology demonstrators were designed and manufactured enabling to reach Technology Readiness Level (TRL) 3. A Life Cycle Analysis (LCA) indicated that the alternative LFW manufacturing routes have the potential for reducing the raw material usage and its waste, as well as the cost of manufacturing and the associated environmental impact. [ABSTRACT FROM AUTHOR]

A launch vehicle needs to adapt to a complex flight environment during flight, and traditional guidance and control algorithms can hardly deal with multi-factor uncertainties due to the high dependency on control models. To solve this problem, this paper designs a new intelligent flight control method for a rocket based on the deep reinforcement learning algorithm driven by knowledge and data. In this process, the Markov decision process of the rocket landing section is established by designing a reinforcement function with consideration of the combination effect on the return of the terminal constraint of the launch vehicle and the cumulative return of the flight process of the rocket. Meanwhile, to improve the training speed of the landing process of the launch vehicle and to enhance the generalization ability of the model, the strategic neural network model is obtained and trained via the form of a long short-term memory (LSTM) network combined with a full connection layer as a landing guidance strategy network. The proximal policy optimization (PPO) is the training algorithm of reinforcement learning network parameters combined with behavioral cloning (BC) as the reinforcement learning pre-training imitation learning algorithm. Notably, the rocket-borne environment is transplanted to the Nvidia Jetson TX2 embedded platform for the comparative testing and verification of this intelligent model, which is then used to generate real-time control commands for guiding the actual flying and landing process of the rocket. Further, comparisons of the results obtained from convex landing optimization and the proposed method in this work are performed to prove the effectiveness of this proposed method. The simulation results show that the intelligent control method in this work can meet the landing accuracy requirements of the launch vehicle with a fast convergence speed of 84 steps, and the decision time is only 2.5 ms. Additionally, it has the ability of online autonomous decision making as deployed on the embedded platform.

Application of composite lattice structures in aerospace application can bring about considerable weight savings, thus allowing for increased payload weight. This study is devoted to reliability analysis of composite anisogrid lattice interstage structure (CALIS) used in typical launch vehicle (LV). CALIS is usually formed in to thin-walled cylindrical or conical shell via helical and circumferential ribs that can act as a lightweight load bearing support structure interfacing any two stages of the LV. Thus, the current study is of practical importance as the interstage structure is acted upon by rather high axial and bending forces emanating from LV external aerodynamics, thrust and inertial forces. As there are always uncertainties associated with the material properties, structural geometry, loading conditions as well as the manufacturing processes, CALIS is faced with stochastic load and strength characteristics. In this sense, CALIS structural reliability is of importance and is thus investigated with respect to material strength, local and global buckling via different reliability techniques. The analytical reliability results are verified against a Monte Carlo simulation and analysis. In addition, a comprehensive sensitivity analysis is performed for some of key CALIS design variables that dominantly affect its structural reliability. [ABSTRACT FROM AUTHOR]

Analytical target cascading (ATC) is a method for coordinating hierarchical system design optimization with a decomposition-based framework. Since a launch vehicle (LV) is usually powered by two or more stages of rocket motors, the overall design of the LV clearly has a hierarchical structure, including system level (conducted by the general design department) and subsystem level (conducted by the motor stage design department). In particular, the subsystem level contains stage-divided elements rather than discipline-divided elements. Therefore, ATC is inherently suitable for the overall design of the LV. This paper presents an ATC decomposition framework for LV design according to practical engineering. The feasibility of the multi-island genetic algorithm (MIGA) used in the ATC decomposition is verified by a mathematical programming test, in which non-linear programming with the quadratic Lagrangian (NLPQL) algorithm is set as a comparison. The multi-disciplinary analysis modules of a hybrid rocket motor (HRM) propelled LV, including propulsion, structure, aerodynamics and trajectory, are established. A hierarchical decomposition is proposed for this multi-level design with a multi-disciplinary model. The application and optimization results verify the feasibility of the ATC decomposition framework with MIGA in the preliminary design of the LV and the final orbit accuracy is better than that of the MDF method. In addition, the final design schemes also prove that HRMs can be considered as a feasible choice of propulsion system for a small payload at low earth orbit. [ABSTRACT FROM AUTHOR]