Browse Topic: Launch vehicles
Dynamic responses at critical locations of a spacecraft due to excitations expected during the ascent phase of a launch vehicle mission are usually estimated through a Coupled Loads Analysis (CLA) using the structural dynamic finite element model of the launch vehicle coupled with that of the spacecraft. Generally, the full physical structural dynamic model of a spacecraft has lakhs of degrees-of-freedom (DOFs). Coupling such a model with a similar model for the launch vehicle results in exorbitantly high computational costs for CLA. Hence, dynamic analysis of such large and complex structural assemblies usually employ sub-structure coupling or Component Mode Synthesis (CMS) methods. The most widely used CMS method for dynamic analyses is the Craig-Bampton (CB) method. Conventionally, a full launch vehicle CLA involves one level of CB-reduction wherein a reduced-order dynamic model of the spacecraft is first generated using the fixed-interface CB-method. This reduced-order model is coupled with the launch vehicle model through the interface DOFs and CLA is performed using this coupled dynamic model. For test missions of ISRO’s manned space program, a simulated Crew Module (CM) is interfaced to the launch vehicle in place of the spacecraft. For CLA of this launch vehicle, a CB-reduced model of the CM is required. The CM comprises several sub-systems, including the Crew Seat Assembly (CSA), and dynamic responses at critical locations of these sub-systems need to be obtained from CLA. Structural dynamic model of the CSA is complex and involves lakhs of DOFs. Interfacing this detailed finite element model of CSA directly with the full model of CM for CB-reduction was found infeasible considering the prohibitively high computational time and resources required. An alternate approach is utilized to overcome the problem wherein the reduced-order dynamic model of CM is generated with two levels of nested sub-structuring. The CSA is represented as a CB-reduced model within the full physical structural dynamic model of CM, which is further CB-reduced and coupled with the launch-vehicle model for CLA. This paper presents the approach adopted to extract internal point dynamic responses on CSA from a dynamic analysis using the reduced-order dynamic model of CM with two-level CB-reduction. To validate the proposed approach, structural dynamic model of a skeletal structure of CM is generated encompassing the full model of CSA. A reduced-order dynamic model of this skeletal structure is also generated with two-level CB-reduction. A typical transient force excitation is considered and dynamic responses at internal points of CSA are estimated from the reduced-order dynamic model using the proposed approach. These responses are compared with corresponding responses obtained from a similar analysis with the full physical model of the skeletal structure and validity of the approach is established. The presented approach is generic and can be conveniently extended to extract responses from any dynamic analysis with models having even more than two levels of nested sub-structuring. Using this approach, the computational time and resource requirements for dynamic analysis studies are minimized. Dynamic analysis capabilities of MSC Nastran software are used for this study.
Gaganyaan is an ambitious and recover safety mission for the Indian space program to launch humans into space. The success of the mission depends on the development of required technology and systems. A test vehicle is developed for the technological demonstration for all envisioned abort flight scenarios of Gaganyaan mission. A new configuration of launch vehicle with single liquid stage is planned for multiple flights. Coupled Loads analysis of launch vehicle system is a standard practice to estimate response and loads for the design of structures and generating sine vibration test levels. Usually a vehicle rests on the launch pad through base shroud with horizontal support and no vertical restraint. Upon ignition of the engine, thrust builds up and upon overcoming gravity the vehicle takes off. In the current analysis the launch vehicle is held in position using a holding / retracting mechanism and at a predefined time the vehicle is released. The boundary condition required a novel method to perform response analysis. The responses estimated from pre-flight analysis is correlated with flight response and a good correlation is observed. Additionally response analysis is performed at engine shut-off. Being the first flight, no flight data was available of engine thrust. In the current work, using an analytical thrust, responses are estimated. The analysis also helped in generating levels for vibration test of various sub-assemblies and payload.
The payload fairing of a launch vehicle is subjected to extremely high acoustic loads, with peak levels occurring during lift-off and transonic aerodynamic regimes. The external acoustic field penetrates the fairing, producing intense internal sound pressure levels that can challenge the integrity of spacecraft components. Accurate characterization of the vibroacoustic behavior of the payload fairing and its enclosed cavity is therefore essential to ensure spacecraft survivability. The internal acoustic field is governed by the coupled dynamics of the fairing structure and the spacecraft configuration, making it critical to quantify the acoustic environment for different payload arrangements. This study presents a detailed vibroacoustic analysis of a payload fairing with multiple spacecraft configurations to evaluate the resulting internal sound pressure distribution. Vibroacoustic finite element analysis is employed in the low frequency range, while statistical energy analysis is utilized for mid and high frequency ranges. Representative models are developed, and the predicted structural and acoustic responses are validated against experimental acoustic test measurements. The validated models are subsequently extended to other spacecraft configurations to perform sensitivity studies. The influence of various parameters on the internal sound pressure levels is assessed, and the resulting perturbations across frequency bands are quantified. The outcome of this study provides a comprehensive understanding of the internal acoustic environment within payload fairing, aiding in the specification of qualification acoustic test level for spacecraft.
Sealing systems in space applications must perform reliably under demanding conditions in engineering: cryogenic temperatures, vibration, leakage control, ultra-high vacuum, ionizing radiation, abrasive particulates, and repeated thermal cycling. Each factor strains conventional sealing technologies. In combination, they can rapidly cause failure in systems where margins are unforgiving and maintenance is impossible. As spacecraft architectures evolve toward longer operational lifetimes and broader mission profiles, sealing requirements continue to tighten. Launch vehicles, satellites, and exploration platforms now operate across wider temperature ranges and in contact with more aggressive propellants and media. As a result, both metal seals and engineered polymer alternatives are evaluated-and selected-against increasingly specific, measurable performance criteria.
The advent of EVs, ride sharing, global events such as the pandemic, chip shortage, and increasing dependency on suppliers are just some factors reshaping the automotive business. Consumer sentiment moving from product to experience resulted in more variants being launched at a record pace. Consequently, product development processes need to be more agile and yet more rigorous while bringing about cohesion and alignment across cross-functional teams to launch vehicles on time, on quality, and in budget. Automotive companies have been using Product Lifecycle Management (PLM) solutions for years to manage CAD, change, and BOMs. With changing business scenarios and increasing complexity of products, the sphere of influence of PLM solutions has expanded significantly over the last decade to manage all aspects of product development. Traditionally PLM software focused on integrating with different authoring tools and managing data in a central repository. The PLM solution had multiple such repositories to manage different types of data—CAD, manufacturing, simulation, requirements, engineering changes, and the like. This resulted in additional overhead of synchronizing and replicating this information across these repositories. This approach is not scalable to meet the dynamic needs of product development today. With a significant increase in the scope of PLM, a platform-centric approach that aims to eliminate rather than integrate silos is essential to the successful adoption of PLM software. The answer is next-generation PLM software such as DASSAULT SYSTÈMES ENOVIA PLM on the 3DEXPERIENCE platform that takes a platform-centric approach to managing data, people, and processes. ENOVIA PLM differentiates by being data-driven, eliminates silos, and models business processes that connect the dots throughout the product development process. Dongfeng Automobile Corporation realized 30% efficiency in the design process and 70% decrease in design problems when designing a van with design software and ENOVIA PLM on the 3DEXPERIENCE platform [1].
Physicists at the Naval Research Laboratory are collaborating with several universities throughout the U.S. to develop a small satellite that will detect the emission of short gamma-ray bursts. U.S. Naval Research Laboratory, Washington D.C. The U.S. Naval Research Laboratory (NRL), in partnership with NASA's Marshall Space Flight Center (MSFC), has developed StarBurst, a small satellite (SmallSat) instrument for NASA's StarBurst Multimessenger Pioneer mission, which will detect the emission of short gamma-ray bursts (GRBs), a key electromagnetic (EM) signature that will contribute to the understanding of neutron star (NS) mergers. NRL transferred the instrument to NASA on March 4 for the next phase, environmental testing. From there, the instrument will be integrated onto the spacecraft bus, followed by launch into Low Earth Orbit in 2027. StarBurst will be installed as a secondary payload via the Evolved Expendable Launch Vehicle Secondary Payload Adapter Grande interface with a mission duration of one year, with the option of extension.
A Coventry University design and materials engineer is leading an international team of researchers in the creation of a new material for liquid hydrogen storage tanks that are used to propel rockets into space. Coventry University, Coventry, UK The future of space travel is seemingly changing by the day and a Coventry University academic is doing his bit to stay at the front of the space race. Dr. Ashwath Pazhani along with an international team of researchers have created a new material for storing the liquid hydrogen used to propel rockets into space by the likes of NASA.
Researchers at the Max Planck Institute for Extraterrestrial Physics have developed a new way to produce and shape large, high-quality mirrors that are much thinner than conventional space-telescope mirrors. The final product is even flexible enough to be rolled up and stored compactly inside a launch vehicle.
Unsteady pressure fluctuations in launch vehicles can induce aerodynamic instabilities, potentially resulting in vibration, structural fatigue, and even catastrophic failure. These risks undermine structural integrity and jeopardize payload delivery, threatening mission success and crew safety. Therefore, precise measurements of unsteady pressure are vital for understanding dynamic pressure distribution and flow behaviour caused by phenomena like shock waves, vortices, boundary layer interactions, and flow separation. While ground-based wind tunnel tests have conventionally provided these insights, this paper presents an on-board system designed for real-time unsteady pressure data acquisition. The system addresses the challenge of accurately resolving high-frequency pressure variations over very high base pressure values. It can be integrated into re-entry vehicles and stage recovery experiments, providing confidence in acquiring data for complex geometrical shapes. Moreover, the capability to store and process data as per requirement during flight enhances its utility in practical scenarios. The system incorporates an 8-channel Sigma-Delta Analog to Digital Converter (Σ-Δ ADC) with necessary signal conditioning. A reprogrammable Field Programmable Gate Array (FPGA) handles data acquisition, ADC configuration and required processing for telemetry-based transfer and storage. Post-processing involves a Cascaded Integrator-Comb (CIC) compensation filter to enhance the overall frequency response within the intended bandwidth. This approach promises invaluable insights into launch vehicle dynamics and unsteady flow phenomena, with the system offering an accuracy of about 0.1% of full scale input range and supporting bandwidths up to 8 kHz.
In any human space flight program, safety of the crew is of utmost priority. In case of exigency in atmospheric flight, the crew is safely and quickly rescued from the launch vehicle using Crew Escape System (CES). CES is a critical part of the Human Space Flight which carries the crew module away from the ascending launch vehicle by firing its rocket motors (Pitch Motor (PM), Low altitude Escape Motor (LEM) and High altitude Escape Motor (HEM)). The structural loads experienced by the CES during the mission abort are severe as the propulsive, aerodynamic and inertial forces on the vehicle are significantly high. Since the mission abort can occur at anytime during the ascent phase of the launch vehicle, trajectory profiles are generated for abort at every one second interval of ascent flight period considering several combinations of dispersions on various propulsive parameters of abort motors and aero parameters. Depending on the time of abort, the ignition delay of PM, LEM and HEM are adjusted in order to minimize the lateral acceleration on the vehicle, at the same time meeting the horizontal range requirement. Aerodynamic load distributions on the vehicle and aero forces and moments on the Grid fin are generated for various Mach No., Angle of Attack (AoA) combinations for jet ON and jet OFF conditions of PM, LEM and HEM. In order to estimate the structural loads for CES during abort, inertia relief analyses are carried out for static load on a free-free finite element model at all the time instances in the abort trajectory as a time sweep simulation considering the appropriate aerodynamic forces, propulsive parameters (mass consumption and thrust) and trajectory parameters (Mach No., Dynamic Pressure and AoA) at each time instant in the trajectory. During abort at lower altitude along with LEM, Pitch Motor (PM) is fired perpendicular to the axis of the vehicle to turn the vehicle towards the sea thereby increasing the AoA and bending moment on the vehicle. The Pitch Motor thrust acting perpendicular to the vehicle excites the first bending mode and augments the Bending Moment. LEM thrust and HEM thrust acting along the vehicle axis also can excite the axial modes of the vehicle. Similarly the lateral aero dynamic forces on the vehicle excite the lateral dynamics of the vehicle whenever there is a sudden change in the AoA due to wind gust. To account for the flexible body dynamic forces, a load augmentation factor called Flexibility Factor is multiplied on the static loads to arrive at the limiting loads on the vehicle. This paper briefly explains the structural load estimation methodology for CES abort, various inputs required and different steps involved in it.
With regards to any aerospace mission, it is very useful to have awareness about the state of vehicle, i.e., the information about its position, velocity, attitude, rotational rates and other concerned data such as control surface deflections, landing gear touchdown, working of mechanisms and so on. The sensor data from the vehicle that is communicated to the ground can be difficult to perceive and analyze. A frame work for real-time motion simulation of an aerospace vehicle from onboard telemetry data is henceforth developed in order to improve the understanding about the current state of the mission and aid in real-time decision making if required. The telemetry data, that is transmitted through User Datagram Protocol (UDP), is received and decoded to usable format. The visualization software accepts the data in a fixed time interval and applies the required transformations in order to ensure one-to-one correspondence between actual vehicle and simulation. The transformations required for missions with various scope are formulated. Data handling scheme for low frequency data and data loss are also discussed. The importance of environment design emphasized, the 3D terrain is created from satellite imagery and heightmap or digital elevation model. The shader for volumetric atmosphere useful for surface to space missions, is created using an algorithm that simulated scattering of light in the atmosphere. For larger environments required for launch vehicle missions, a scaling scheme is worked out and is implemented in combination with the floating-origin algorithm to deal with the floating-point limitations posed by current generation graphic engines. Techniques such as quad-tree structures for level of detail (LOD) rendering are used for optimization of performance. Animation of flight events are rigged and get triggered based on flags in telemetry data. The developed framework was tested in critical demonstration missions for ISRO – The Reusable Launch Vehicle Landing Experiment (RLV-LEX) and Test Vehicle Demonstration Mission -1 (TV-D1). The framework can be used for a variety of missions such as launch vehicles, spacecrafts, UAVs and even landers & rovers.
Launch vehicle structures in course of its flight will be subjected to dynamic forces over a range of frequencies up to 2000 Hz. These loads can be steady, transient or random in nature. The dynamic excitations like aerodynamic gust, motor oscillations and transients, sudden application of control force are capable of exciting the low frequency structural modes and cause significant responses at the interface of launch vehicle and satellite. The satellite interface responses to these low frequency excitations are estimated through Coupled Load Analysis (CLA). This analysis plays a crucial role in mission as the satellite design loads and Sine vibration test levels are defined based on this. The perquisite of CLA is to predict the responses with considerable accuracy so that the design loads are not exceeded in the flight. CLA validation is possible by simulating the flight experienced responses through the analysis. In the present study, the satellite interface responses are validated for a launch vehicle with solid motors. The source of dynamic force in solid motor is mainly the motor ignition transient. Transient response analysis is carried out using the Finite Element models of launch vehicle coupled with the satellite to simulate the responses during solid motor ignition. The required excitation is generated from the motor ignition transients measured in the flight. The criticality in the simulation is to define and model the forcing function appropriately. The responses estimated from the analysis are compared with the responses measured in the flight and observed to be in good match in both temporal and frequency domain. The study confirmed that the forcing functions developed from the flight measured data are adequate. The post flight simulation studies helps to improve the prediction methodology for future missions.
In recent years, industry adoption of thermoplastic composites (TPCs) in lieu of thermosets and metallic structures has increased for the fabrication of air and launch vehicle components. Manufacturing of TPCs, performed via automated tape laying (ATL) and automated fiber placement (AFP), uses machines that place prepreg tow or tapes on molds in a unidirectional manner, which then undergo cure cycles, autoclaving, and other steps that require special tooling. The process is time, material, and energy intensive, requires large facilities to house equipment, and limits the size, mechanical properties and shapes of the parts manufactured. To address these limitations, NASA’s Langley Research Center has developed a simplified, tool-less automated tow/tape placement (ATP) system.
This SAE Aerospace Information Report (AIR) includes all missile and launch vehicle actuation systems, including electrohydraulic, electropneumatic, and electromechanical types. The data for many systems are not complete. As more information becomes available, periodic updates will be issued to complete existing data sheets and to add new ones. An index by type of vehicle and by type of actuation system is included. The actual data sheets in the body of the report are organized in alphabetical order.
Cranes for lifting and lowering heavy objects are an important and sometimes essential tool in modern industries such as construction, transportation, and manufacturing. NASA uses overhead and mobile cranes for assembly of load lines employed in full-scale testing of its Space Launch System (SLS), a super-heavy-lift launch vehicle for deep space human space exploration. Structural testing of the SLS requires precision placement of heavy objects with soft contact during mating connections, which proved to be problematic with the relatively coarse control available with motor-driven overhead cranes and the existing rigging devices.
Louisiana State University Baton Rouge, LA
Cranes for lifting and lowering heavy objects are an important and sometimes essential tool in modern industries such as construction, transportation, and manufacturing. NASA uses overhead and mobile cranes for assembly of load lines employed in full-scale testing of its Space Launch System (SLS), a super-heavy-lift launch vehicle for deep space human space exploration. Structural testing of the SLS requires precision placement of heavy objects with soft contact during mating connections, which proved to be problematic with the relatively coarse control available with motor-driven overhead cranes and the existing rigging devices.
Using rockets to launch satellites and people into orbit currently requires a lot of high-energy fuel, which is 95% of total rocket mass. Launching a pound of payload can cost $10,000 or more, so minimizing the total cost of launching rockets would maximize the scientific payloads and increase the feasibility of space exploration.
To comply with the stringent fuel consumption requirements, many automobile manufacturers have launched vehicle electrification programs which are representing a paradigm shift in vehicle design. Looking specifically at powertrain calibration, optimization approaches were developed to help the decision-making process in the powertrain control. Due to computational power limitations the most common approach is still the use of powertrain calibration tables in a rule-based controller. This is true despite the fact that the most common manual tuning can be quite long and exhausting, and with the optimal consumption behavior rarely being achieved. The present work proposes a simulation tool that has the objective to automate the process of tuning a calibration table in a powertrain model. To achieve that, it is first necessary to define the optimal reference performance. The calibration table then has its values optimized by the Genetic Algorithm to a single value that better matches the reference performance. A novel Iterative Histogram procedure is then used to identify which cells from the new table have the greatest contribution to the performance mismatch between the model and the reference. These values are optimized and the histogram is reassessed. This process is repeated until the mismatch target is achieved or the model results show saturation in its performance. The iterative nature of this process results in a powerful tool that gives its users the ability to easily conduct a simulation while simultaneously monitoring the results of each iteration until the target is met.
On September 1, 1961, NASA requested appropriations for initial land purchases on Merritt Island on Florida’s east coast to support the Apollo Lunar Landing Program. Designers quickly began developing plans for Launch Complex 39 facilities, which include the Launch Control Center, Pads A & B, and the huge hangar now known as the Vehicle Assembly Building (VAB).
In recent years there has been a trend towards the wider use of COTS (Commercial Off The Shelf) equipment in space missions. This trend has been mainly driven by the restrictions in R&D budgets and a growing demand for shorter design cycles. Funding Agencies are encouraging designers of spacecraft systems to identify and overcome the obstacles that previously prevented the use of COTS products for space missions. When it comes to space vehicle engineering, the tolerance of onboard electronics to radiation effects can be one of the most challenging aspects of the system design. The risk of failure for avionics equipment on-board spacecraft due to radiation exposure is determined by the vehicle's orbit trajectory and flight duration, during which the vehicle is exposed to trapped radiation as well as solar and cosmic radiation sources.
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