Integrated Design for Space Transportation System

The satellite systems provide a number of services to mankind in the areas of societal applications. To derive maximum benefits from the satellites, they have to be positioned into the right orbital slot in space. Space transportation systems (STS) play a key role in transporting the satellites from Earth and placing in the specified orbits depending on the application requirements. With the limitations of the technologies available as on date, STS has to be configured with multi-stages and the vehicle subsystems have to be totally autonomous and automatic. To reduce the cost of launching satellites by a specified STS, the STS must deliver maximum payload mass into the specified orbit with very high accuracy. The vehicle must be highly robust and this poses conflicting requirements on the design of each of the subsystems. The robustness demand of STS increases many fold when the humans are onboard, and in such cases, the transportation systems have to meet the stringent requirements of human rating. Therefore, various aspects of the optimum and robust design of such STS satisfying all specified conditions are discussed in this chapter. The functional requirements of STS to transport these satellites to space or to bring back to the Earth, typical subsystems and operating environments are also included. Development cycles and cost aspects of STS are explained. The design challenges for the complex subsystems and the need for the integrated design using systems approach to arrive at an optimum, cost effective and robust design are also included. The development of advanced technologies involving design, characterization, testing, and qualification has a major implication on the cost and schedule for realization of the vehicle and their salient aspects are discussed.

The design of a launch vehicle demands close involvement of many disciplines like propulsion, vehicle structures, aerodynamics, flight mechanics, navigation, guidance and control (NGC) systems, vehicle avionics, stage auxiliary systems, thermal systems, etc. It is essential to integrate the requirements of these systems to obtain the best design for the specified mission, and the design iterations are inevitable to arrive at a good design. The interfaces among disciplines are not always well defined. Design of a launch vehicle is essentially making suitable compromises to achieve a good balance between many conflicting requirements and interactions. The design process also has to depend a great deal on analyses, experimentation and simulations. To arrive at an optimum launch vehicle design, it is essential to apply systems approach involving various steps like: (a) mission definition, (b) requirements analyses, (c) subsystems definition, (d) concept analyses and definition and (e) design. Design iterations are necessary to understand the interaction among subsystems, interfaces between various disciplines and relation between various design functions to meet the specified requirements. There are severe dynamic coupling between various systems which further aggravates the complexity of the design process. The clear understanding of the interfaces between various disciplines and their functional interdependencies with respect to various operating conditions is a vital factor. Therefore, it is essential to integrate various subsystems through a structured multidisciplinary process. The complex design problem of a launch vehicle can be handled only by understanding the roles of each of the vehicle subsystems and by utilizing the systems engineering approach right from the concept stage. As space missions are expensive, the quality, reliability and cost should get maximum attention all through the design and development phases. Various aspects of integrated design of STS using systems approach, which are essential to arrive at an optimum, robust and cost effective design of STS are explained in this chapter. An integrated approach for STS design evaluation using suitable vehicle system models to arrive at robust and optimum design is also discussed in brief.

To specify the functional requirements of STS, it is essential to understand the orbital motion of the injected satellites under the influence of a central gravitational force and other disturbing forces. This chapter deals with astrodynamics which explains the motion of celestial bodies as well as human-made satellites under the influence of gravitational force field of celestial bodies and other external forces. Orbital motions of Low Earth Orbit (LEO) satellites are the solutions of two-body problems i.e. the Earth and the satellite, in the specified reference frame, considering the Earth’s gravitational force as the primary central body force field. The deviation of the gravity force away from the central force field and other disturbance forces affect the orbital motion of the satellites. In addition to the central force field, gravitational forces of other planets and Moon also influence the higher altitude orbital motions. Solutions for such motions are achieved by solving restricted three body problem, considering the Earth’s gravity force as the central gravity field whereas the perturbing gravitational force is from the third body such as Moon. Depending on the type of trajectory, different reference frames are used and theses aspects are explained first. Then, this chapter discusses the orbital mechanics of satellites and various aspects of orbital motions of two-body problems. The restricted three body problem and the resulting orbital motion are also briefly explained. Even though the launch vehicles are capable of injecting the satellites in the near Earth orbits, for certain scientific applications, these satellites have to reach and orbit around Moon or distant planets. The interplanetary trajectories of such satellites from the Earth bound orbits to the target planets are also included. In an integrated mission management, optimum strategies like transferring the satellite at suitable time from the initial orbit to the required one in an optimum fashion are required. Various optimum orbital transfer maneuver strategies are explained in this chapter.

The function of space transportation system is not only to lift the specified satellite from Earth surface, travel through space and to inject it precisely into the defined orbit but also to achieve all specified orbital specifications simultaneously. To achieve these objectives, the STS has to provide the required mechanical energy to the satellite, defined by the orbit size and satellite mass. It is also necessary to have an optimum location of launch site on the surface of the Earth and launch direction (launch azimuth). Therefore, each orbital mission demands specified location of launch site and launch azimuth. But, due to geographical constraints, it is not possible to have an optimum launch site, and the allowable launch azimuth directions are also limited due to range safety related issues. In such cases, launch vehicle has to provide extra energy to reach the defined target. The influence of Earth’s gravity field and aerodynamic drag during atmospheric flight phase causes the energy loss and to compensate these losses extra energy has to be provided. All these effects need to be considered for configuring STS to provide the necessary energy to the satellite. The satellites requirements are widely varying, ranging from low Earth orbits to high altitude orbits, depending on their applications. Therefore with the effective utilization of available energy in the STS, the maximum performance has to be achieved by adopting suitable strategies and the details of all these methods are explained in this chapter. The functional requirements for STS design have to start with the orbital mission requirements and the satellite mass to be placed in the specified orbit. The step by step procedure starting from mission design to STS design process, which satisfies the overall functional requirements, is also described. Various errors occurring during the operation of several subsystems of STS, their effects on overall mission and how to alleviate them are also explained here.

In all space missions, the main objective during the design phase is to define a suitable STS configuration which has the needed capability to inject the defined spacecraft into the specified orbit. To achieve these objectives, STS is configured with suitable propulsion systems to impart the required equivalent velocity to the satellite, after accounting for the various velocity losses the vehicle encounters during its mission. This chapter discusses the STS configuration, selection criteria to provide the required energy to inject the satellite into the orbit, while meeting the various other subsystems requirements. The optimum STS configuration has to provide the maximum performance with respect to the vehicle size, maximum reliability, reduced unit cost for the vehicle, minimum cost for design and development and a reasonable schedule for the realization of the vehicle. To arrive at an optimum vehicle sizing and configuration, exhaustive trade-off studies are required. The studies have to address existing technologies and advanced new technologies with respect to performance improvements, performance requirements, demonstrations, process technologies, advanced technology developments, overall vehicle reliability, unit cost of the vehicle, technology risks, increased development cost and schedule, limitations on the realization of the systems, etc. The selection methodology for STS configuration, design guidelines and the various processes involved in the design of an optimum configuration are discussed in this chapter. Design strategies for various vehicle subsystems which influence the vehicle performance and the vehicle configuration are highlighted. The constraints imposed by the existing technology limitations and the possible future technology improvements are explained. The requirements of multi-staging and their relative merits and demerits are highlighted. The optimum staging aspects, the design sensitivities and various factors which influence the vehicle performance are elaborated, Attention to ground and launch operations while finalising the configuration of the vehicle is of utmost importance because it has large influence on the technology options for the vehicle, schedule and cost. The infrastructure needed for ground operations like launch complex, launch tower, propellant servicing facilities, etc. has to be closely linked to the vehicle configuration. Therefore these aspects are addressed in brief at the end of the chapter.

The space transportation systems and their subsystems have to perform normally under all kinds of adverse environments during their operation in flight. The vehicle experiences very severe external environments during atmospheric regime of flight and subsequent to that faces steady and dynamic loads during the remaining phases of flight till injection of satellite. The vehicle structure has to withstand all these extreme flight environments to achieve the mission successfully. To counteract the disturbance and to achieve the intended function, the various subsystems have to stretch their functional limits. The additional response caused by such efforts influences the performance of other related subsystems. Similarly, performance dispersions of a particular system have functional impact on other subsystems. Therefore, there is a strong coupling between environment and performance of subsystems in the vehicle. The disturbances are originated either from the external source to the vehicle or from a specific system within the vehicle which acts on the vehicle. The vehicle and subsystems have to be designed to operate against the expected environment disturbances. Typical external operating environments are gravity, atmosphere and aerothermodynamics. Atmospheric wind has specific characteristics and has major influence on the performance vehicle systems. The full understanding of the thermal environment which can cause severe effect on the performance of the subsystems is essential. Dynamic environments experienced by the vehicle systems, by both external and vehicle internal sources have the potential to induce the coupling between subsystems which can cause severe degradation in performance and at times failure too. This chapter describes in detail all such external, internal and dynamic operating environments experienced by the vehicle subsystems. How to deal with all such hostile environments during the design phase and how to enhance the robustness of the systems to work in such environments are discussed. Methodologies for understanding the vehicle operating environments thoroughly, predicting them accurately and utilizing those values in the design and qualification process are explained. The various parameter dispersions to be considered for the vehicle design and overall mission are also discussed.

During the entire phase of space transportation system mission, from lift-off till satellite injection, various constraints and requirements applicable not only to the mission but also to vehicle systems, ground systems, range safety and tracking systems are to be satisfied. Considering these constraints and requirements, an optimum feasible trajectory has to be designed to meet the mission requirements. The mission design process involves the utilization of the available energy for realizing the defined orbital mission by devising suitable strategies of directing the energy along the suitable path and sequencing the energy addition process. Optimum mission design strategies have to be arrived at to achieve the maximum performance, ensuring the defined mission under nominal and off-nominal flight environments and system parameter dispersions. The trajectory shaping satisfying vehicle loads and radio visibility for continuous tracking coverage during ascent phase are other essential parts of mission design. There are number of constraints like thermal loads on the spacecraft, the vehicle subsystems during ascent phase and jet plumes of reaction control system thrusters interacting with the spacecraft. The passivation requirements of the final stage after spacecraft separation are to be carefully worked out. Another important aspect of mission design is to finalize the flight events/sequences which generate various commands to separate the stages and to initiate the subsequent flight events. In such complex systems close interactions among various disciplines exist, and the mission design requires several iterations. In this chapter all these aspects of mission design are discussed in detail and various activities involved in mission design process explained. The mission design strategies and importance of the same for the vehicle design process are included. Mission requirements, constraints, design and analysis aspects and trajectory design constraints during various phases of trajectory are presented. Mission sequence design considerations and all other mission-related studies like satellite orientation requirements for multiple satellite launch and passivation requirements to ensure the safety of the spent stage are also highlighted.

The flight mechanics is an important ingredient in vehicle and subsystem design, its performance evaluation and its validation. It deals with forces and moments acting on bodies and the response of the bodies to the applied forces and moments. Typical applications are trajectory design, optimum payload estimation, navigation, guidance and control (NGC) system algorithm design, estimation of loads on vehicle and various subsystems, mission design, vehicle sequencing, performance evaluations of subsystems, validation of NGC systems and evaluation of mission performance. Flight mechanics consists of two processes, namely, modelling and solution where the modelling represents subsystems, vehicle, forces and moments acting on it, their operating environment and the dynamics of vehicle and subsystem. The solution involves obtaining the solutions to the mathematical models, which truly represent the response of the vehicle and subsystems. To achieve the error-free design, the various models used for the vehicle, subsystems, environment, forces and moments generated by the respective systems as well as the dynamics of the systems under the influence of the forces and moments have to represent very close to the physical system and process. The systems models vary from simple three-dimensional model for the translational motion of the vehicle centre of gravity to the detailed six-degrees-of-freedom model along with the flexible vehicle structural dynamics. In addition, to evaluate the integrated system performance, the modelling of the vehicle onboard system elements such as sensors, navigation, guidance and control systems, the corresponding algorithms and signal flows simulating the delays in data transmission among various systems are required. In this chapter, the role of flight mechanics in the STS design process and the need for the integrated design approach are explained. The different coordinate systems used to represent the mathematical models, vehicle attitude sign conventions and the coordinate transformation to transfer the data between the reference frames are described. Subsequently, various models necessary for representing vehicle, subsystem and environment and the methodology for evaluating the system response are included. The usage of the flight mechanics models in the design process is also highlighted.

Propulsion systems in space transportation systems have to impart the necessary energy to the vehicle to achieve the desired orbital conditions for the specified satellite. Different categories of propulsion systems such as chemical propulsion, electric propulsion, nuclear propulsion, solar sail, etc. are being used in various applications of space missions, depending on the requirements. For the boost phase with the present day technologies, chemical propulsion systems are generally being used and the energy source is the chemical reaction. There are two types of chemical propulsion systems like air breathing propulsion and non-air breathing propulsion. In air breathing propulsion, the oxygen available in the atmosphere and the fuel stored vehicle onboard are used for the combustion process, whereas in the case of non-air breathing propulsion system, both oxidizer and fuel stored vehicle onboard are used for generating the necessary thrust. Depending on the type of propellants used, the chemical rocket propulsion systems are classified into solid motors rockets, liquid engines and hybrid propellant rockets. In solid propellant rocket motor, the propellant is stored in combustion chamber and propellant burns from the surface. The liquid propellants are categorized into bipropellant and monopropellant. Depending on the propellants used, the bipropellant systems are further classified into Earth storable systems, cryogenic systems and semi-cryogenic systems. In the hybrid rocket propulsion system, generally the fuel is solid and the oxidizer is liquid. The payload capability of a multi-stage vehicle has to be maximized and in achieving this objective the propulsion module for each stage has to be carefully selected. The selection of the number of stages and type of propulsion system depend on the mission objectives, state-of-the-art technologies in propulsion, technology base available, development lead time, and cost and reliability requirements. The selection and design of a suitable propulsion system has to be carried out by considering various factors and also its close interactions with several other major subsystems of the vehicle. This chapter addresses important system engineering aspects of propulsion systems like the selection and design of propulsion modules for a STS, the basics of rocket propulsion, several propulsion options and their relative merits and demerits. The staging aspects and criteria for selection of suitable propulsive modules are highlighted. The qualification process for the propulsive stages is also included. A brief discussion on air breathing propulsion system also outlined.

Aerodynamics plays a vital role in STS design process. The steady and unsteady aerodynamic loads generated due to the relative motion between vehicle body and surrounding air during the atmospheric flight phase of the vehicle, its aerodynamic characteristics, external environment and flow characteristics are the major inputs for the vehicle systems design. The aerodynamic axial force is used for the mission design and vehicle performance evaluation. The steady aerodynamic disturbance moments and loads acting on the vehicle are required for vehicle control and structural systems design. Further the aerodynamics-related issues become complex since the maximum dynamic pressure, maximum aerodynamic force and moment coefficients, unsteady loads, maximum wind velocity, large wind shear and measured uncertainties of wind occur almost simultaneously. The vehicle sizing with clustered propulsion systems complicates the aerodynamic flows with unfavourable additional aerodynamic loads. The unsteady pressure fluctuations are due to shock oscillations during the transonic regime and local aeroacoustic loads due to flow separation because of protrusions form a major input for the qualification of structural and other sensitive systems. This chapter presents the aerodynamic characteristics of a launch vehicle during various phases of its atmospheric flight, different types of aerodynamic loads acting on the overall vehicle/its subsystems and their importance for the launch vehicle systems design. Methods of estimating these loads under different conditions are also outlined. Subsequently, aerodynamic force and moment on overall vehicle, its impact on the vehicle systems design and the methods of aerodynamic characterization using the computational fluid dynamics and wind tunnel studies are briefly covered. The salient features of aerodynamic configuration design of a launch vehicle are also included.

The structure of the space transportation system is the physical body of the vehicle which houses all major or minor subsystems. It consists of propellant tanks, motor cases, payload fairing and interstage structures. The entire structure of the vehicle has to support all associated secondary elements like mechanisms, pyros, avionics, actuators, etc. Structures are classified as primary and secondary structures. Load-bearing structures like the solid stage motors, liquid stage tanks, interstage structures, payload fairing and interface joints are all primary structures. Gas bottles, fuel tanks for control systems, brackets, avionics packages, etc. which do not experience direct loads are known as secondary structures. The design of structural elements has to ensure the structural integrity of the vehicle during its various phases of flight starting from lift-off till satellite injection. Loads, materials, their characteristics, structural construction, manufacturing processes, structural dynamic response, stability characteristics, development schedule, cost, etc. strongly influence the structural designs. During the structural design process the mass of the structure has to be kept minimum to maximize the vehicle performance while ensuring adequate design margin. This makes the launch vehicles very flexible and hence the detailed structural dynamic studies are essential. Therefore, selection of suitable materials for structures and their shape and construction methods are important. To ensure that the designed product meets the specified requirements, the compliance has to be checked through detailed analysis and testing. Structures designed and analyzed for various load cases are to be qualified through a series of structural static and dynamic tests. This chapter discusses in detail the STS structural design requirements, load analysis and different configurations to meet different requirements. The materials’ selection for structures, their shape and construction methods are included. Static and dynamic analyses are described and some insight on design tools is presented. The various static, environmental and dynamic tests needed for the qualification of the structures are highlighted.

The STS during its entire flight regime is exposed to different thermal environments, causing severe thermal loads to the vehicle, structural elements and several other sensitive elements. All thermal effects such as local heating rate and total heat load are to be analyzed in detail to understand the thermal loads during flight. Thermal environment of the vehicle and subsystems caused by aerodynamic heating depends mainly on their external configuration, vehicle surface material characteristics, flow field characteristics and vehicle trajectory. The thermal load caused by propulsion system depends on the type of propulsion system, vehicle operating altitude, nozzle expansion ratio and the vicinity of the subsystem elements with respect to the propulsion elements. Thermal protection materials and thickness in turn decide the thermal protection system mass depending on thermal environment, type of thermal load, materials used in the system and temperature constraints specified for the various subsystems. Thermal protection systems (TPS) are passive, semi-passive and active depending on the application. While appropriate TPS is used to ensure the normal function of the subsystem to meet the specified functions, the mass of the integrated vehicle has to be minimized. During the initial development phase, an integrated system design approach is required to arrive at optimum structural and thermal designs for the vehicle subsystems. Once the suitable thermal protection materials are chosen based on the detailed analyses, it is essential to carry out thermophysical and mechanical property tests for these materials within the temperature range they are expected to experience in flight. This chapter presents the thermal design aspects of vehicle and subsystems for a launch vehicle. The impact of the thermal environments, on vehicle and subsystems, the need for the integrated design strategy, the requirements of various subsystems which need thermal protection, design constraints and approach for optimum thermal design for each of the subsystems are highlighted. The various aspects of the heating environment due to jet exhaust are described. Thermal response analysis and the methodology for the analysis are covered. Tests for thermal protection systems and their qualification methods are also included.

The launch vehicles are generally configured with multiple stages to achieve the required orbital velocity. To achieve the required orbital conditions, the burnt-out stages are to be separated from the vehicle immediately after meeting their intended functional requirements. These separation processes are mission-critical functions as the inadvertent collision during separation can lead to vehicle failure. All these systems are generally termed as stage auxiliary systems (SAS) and carry out the mission critical separation processes in the vehicle as per the specified requirements, thus ensuring the vehicle safety and successful mission. The stage auxiliary systems are configured with high-energy systems and they are built with pyro-elements, which, once assembled in the flight systems, are not amenable for ground tests. Thus, the stage auxiliary systems have only one chance to operate, that too directly in the flight after its manufacturing and these systems have to operate successfully in the first attempt itself. Therefore, these systems have to be highly reliable. They have to be necessarily robust with suitable built-in redundancy to achieve the safer and successful mission. The design of SAS is closely linked with all other subsystems of the vehicle. The high-energy pyro-systems during their operation induce severe environment to the vehicle systems. Design of the SAS is based on the vehicle systems inertial properties, vehicle subsystems, performance parameters and the vehicle operating environment. Therefore, integrated design approach is essential to achieve the robust and highly reliable designs for the complex, high-energy stage auxiliary systems. During the vehicle flight from lift-off till satellite injection, the vehicle can deviate from its nominal path due to abnormal behaviour or due to failure of any of the subsystems onboard. In such cases where the deviation is beyond the safe allowable corridor, the flight has to be terminated using vehicle destruct system. Additional care should be taken to avoid the inadvertent activation of the vehicle destruct system during normal flight. The stage auxiliary systems’ requirements of a launch vehicle, their functional aspects, integrated design aspects of SAS and the various elements involved in the SAS design process are explained in this chapter. The details of the actuators used in these systems based on pyro-mechanical devices, the jettisoning and destruct systems, their performance in a typical vehicle and the validation strategy of these important subsystems are also presented. Detailed analysis of the separated body dynamics with respect to the ongoing vehicle is very vital in the separation system design. The various aspects of system analysis are included.

The navigation, guidance and control (NGC) system, the brain of the vehicle, is responsible for directing the propulsive forces and stabilizing the vehicle along the desired path to achieve the orbit with the specified accuracy. The NGC system has to define the optimum trajectory in real time to reach the specified target and steer the vehicle along the desired path and inject the spacecraft into the mission targeted orbit within the specified dispersions. The navigation system measures the instantaneous state of the vehicle, and using this information, the guidance system generates the optimum trajectory to achieve the target and desired vehicle steering command to realize the optimum trajectory in real time. The vehicle control system, comprising of autopilot and control power plants, receives the steering commands from the guidance system and steers the vehicle to follow the desired attitude in the presence of all disturbances. The guidance system charts out the remaining path continuously, recalculating the desired attitude of the vehicle to achieve the mission target. Final mission objectives and allowable orbital dispersions dictate the choice of a suitable guidance system. The vehicle autopilot has three major functions namely, to ensure that the vehicle loads are always well within the specified limits, to stabilize the vehicle all through the flight and to steer the vehicle to follow the desired attitude as decided by the guidance system. The autopilot generates the control commands as per the defined control law which is used to drive the actuation systems to generate the necessary control forces. Proper choice of inertial sensors, inertial systems, guidance schemes, control actuation systems and control laws depend on the mission objectives and requirements. This chapter starts with the functional requirements and systems requirements of NGC system of a launch vehicle. The need for the integrated design of NGC system using systems engineering approach is explained next. Various elements involved in NGC system like navigation, guidance and control, the selection of a suitable scheme and their design aspects are also explained. The performance evaluation of the integrated design of NGC system carried out using different test beds is described.

After orbiting the spacecraft, in many cases the materials are to be brought back from space to Earth. It is also essential to bring back the humans safely from space after completion of the space exploration experiments. During the return of the vehicle from space, it re-enters the Earth’s atmosphere with orbital speed and remains in the atmosphere for the entire duration of the mission till the vehicle is brought to rest at the specified location. During this regime, the vehicle travels with very high hypersonic speeds at relatively lower altitudes and this causes harsh environments to the vehicle along with different complex flight regimes. Design of such a vehicle which has to fly safely in the severe operating environment with highly varying flight operating regimes along with large dispersions in both flight and vehicle parameters and ensuring the safe landing at the specified location is quite complex. All the reentry space transportation systems are having a wide range of flight regimes with large dispersions in flight parameters, severe flight environments and different functional requirements. For the case of reentry missions, the aerothermal operating environment is entirely different for different missions and strongly depends on the trajectory of a particular reentry mission which makes the vehicle design as unique for the specified reentry mission. The reentry systems are truly interdisciplinary with strong coupling between various vehicle systems viz., aerodynamic configuration, thermal protection systems, structure and vehicle trajectory. Therefore, integrated systems approach is essential for the optimum reentry systems design. This chapter addresses the design guidelines, complexity of the flight environment and the design for reentry systems. The reentry dynamics and reentry vehicle configuration aspects are discussed. The aerothermodynamics aspects which involve aerodynamic design, structural design, thermal environment and thermal protection system design of reentry vehicles are included. The details of the thermal protection systems are presented. The salient aspects of the trajectory design, reentry guidance schemes and mission management are highlighted.