Crew Module Research Articles

Study on the safe landing of a crew module considering horizontal velocity using a DEM-FEM coupling algorithm

Due to the influence of shallow wind and reverse thrust on the ground, the horizontal velocity of the crew module poses a threat to its safety during landing. To address this issue, the discrete element and finite element coupling algorithm is used to investigate the landing process of the crew module with horizontal velocity in this paper. Firstly, the discrete element model of the soil at the designated landing site and the finite element model of the crew module structure are established. Subsequently, the discrete element-finite element coupling algorithm of the interaction between the soil and the crew module is applied. The above model is used to analyze the influence of horizontal velocity on the crew module landing process in vertical and inclined landing. The results indicate that the horizontal velocity significantly affects the landing process of the crew module during vertical landing. Additionally, during inclined landing, the dynamic characteristics of the crew module are compared under three working conditions: the horizontal velocity direction is the same as the oblique direction of the crew module, the reverse direction and the vertical direction. When the horizontal velocity direction aligns with the oblique direction of the crew module, the influence of horizontal velocity on the landing process is relatively minor. However, when the horizontal velocity opposes the oblique direction of the crew module, it has a more substantial influence on the impact force and the displacement of the crew module. The research results provide valuable theoretical insights to ensure the safe landing of the crew module.

A novel concept on Mid Air Recovery of crew module

Crew recovery is one of the most critical phase of Human space flight missions. The current recovery strategies include both water (sea) and land-based recovery operations. The land-based recovery operations mandate additional systems to reduce impact loads whereas water-based operations involve mobilization of many recovery assets. Considering the increased frequency of human space flight missions, reuse of crew capsules, increased crew comfort, minimization of assets and reduced cost, it is necessary to explore alternate options for the recovery of the Crew Module (CM).A novel concept on Mid-Air-Recovery (MAR) of CM is explored to achieve the above objectives. MAR of CM will significantly reduce the assets required in comparison to land-based and sea-based recovery operations. With MAR operations, the CM can be carried directly to the recovery base in minimum time. The loads experienced by the Crew are also expected to be lesser in case of MAR operations. Since the CM is not subjected to any impact loads and not exposed to seawater, the reusability of modules is enhanced. Another advantage of MAR operation is the simplification of the overall recovery procedures (SOPs) and reduction in trained manpower requirements. A helicopter is chosen for study, since it offers great manoeuvrability and remarkable cargo capabilities. A Recovery chute is added to the CM and an engagement hook is designed to capture this chute during MAR.The overall configuration for MAR and subsystems that need to be developed have been discussed as a part of this study. The components that are being added in CM are planned to be a standalone system so that the performance of other systems is not affected in case of failure of MAR operations. The impact of MAR operations on the Crew, in terms of physiological loads is also addressed.

Nonlinear, Rate-Dependent Evaluation of Flight Damping Derivatives of the Orion Crew Module

In the continued development of the Orion crew module, NASA is challenged with characterizing and predicting static and dynamic stability during the atmospheric reentry of the capsule. A thorough understanding of the aerodynamic behavior of the crew module in the subsonic regime is essential for safe return, as this flight regime presents the largest dynamic instability recently demonstrated by the Ascent Abort-2 (AA2) flight test. Forced oscillation wind-tunnel tests were conducted with varying frequency, amplitude, and offset angles to determine the rate-dependent static and dynamic derivatives. A novel data reduction technique was employed to obtain a nonlinear representation of the damping derivatives of the capsule as a function of both its angular position and rate from the aerodynamic force/moment measurements. Finally, a nonlinear, second-order model, two-degree-of-freedom ordinary differential equation was formulated to predict the response of the crew module based on initial angular positions, and , and angular rates, and . The accuracy of the model was assessed through direct comparisons with the NASA Ascent Abort-2 flight test. Through benchmark comparisons against the AA2 flight-test data and rate-invariant models, the utility of the second-order model was proven, establishing a low-cost, computationally efficient manner of predicting crew module motion and stability that could be leveraged in flight to improve guidance and navigation control algorithms.

Concept for 2033 Crewed Mars Orbital Mission with Venus Flyby

The 2033 Mars launch period provides a unique opportunity for a round-trip mission with a total flight time of only 1.6 years. A concept is presented to perform a crewed Mars orbital mission in 2033 that would minimize development and mission risk by using conventional hypergolic in-space propulsion stages with a common design. The Mars mission vehicle would include a Mars transit habitat, an Orion spacecraft, and chemical propulsion stages. It would be launched by a combination of Space Launch System and commercial rockets to be aggregated in high Earth orbit or at the Lunar Gateway. The crew would launch to the mission vehicle in Orion as the final element in the assembly. After transit to Mars, the crew would spend about 30 days in a high Mars orbit and then return to Earth via a Venus flyby and gravity assist. A sunshade would be deployed for thermal control of the mission vehicle inside 1 astronomical unit. The crew would return directly to Earth in the Orion crew module, with the transit habitat being expended. The 2033 mission would not be a “one-off” but could be a pathfinder for crew transport for landing missions to follow, perhaps starting as early as 2037.

Development of acoustic mufflers for cabin noise reduction in Orion spacecraft

Controlling cabin acoustic noise levels in the Crew Module (CM) of the Orion spacecraft is critical for adequate speech intelligibility, avoid fatigue, and prevent any possibility of temporary and permanent hearing loss to the crew. The primary source of cabin noise for the on-orbit phase of the mission is from the Environmental Control and Life Support System (ECLSS) which recycles and conditions breathing air and maintains cabin pressurization through its ducting network and components. Unfortunately, as a side effect, noise from the ECLSS fans propagates through theses ducts and emanate into the cabin habitable volume via the ECLSS inlet and outlets. To mitigate excessive duct-borne noise, two ECLSS mufflers have been designed to provide significant acoustic transmission loss (TL) so that the cabin noise requirements can be met. Each muffler is meant to be installed in the ducting of the ECLSS air inlet and outlet sides, respectively. Packaging constraints and tight volume requirements necessitated the mufflers to be of complex geometry and compatible with the bends of the ECLSS duct layout. To design and characterize the acoustic performance of the inlet and outlet mufflers, computational acoustic models were developed using the Finite Element Method (FEM) with software. Characterization of the acoustic material and perforations in the mufflers were addressed with poro-elastic theory. Once the mufflers were designed on paper and its TL predicted, prototypes of these mufflers were created using additive manufacturing. The muffler prototypes were subsequently tested for acoustic TL in the laboratory with various configurations of acoustic materials. Comparing the analytical predictions to the test performance yielded excellent correlation for acoustic TL and demonstrated significant broadband noise attenuation. The ECLSS mufflers are currently scheduled to be installed on the Artemis II CM of the Orion spacecraft and will provide significant cabin comfort to crew during the mission.

Onset Separation Aerodynamics of Crew Module from Crew Escape System

Aerodynamic characteristics of Crew Module (CM) during its separation from the Crew Escape System (CES) are essential to carry out separation dynamics analysis for confirming collision-free separation. As CM is at the base of CES, it has to clear the complex wake flow that prevails in the aft region. A 1 : 10-scale wind tunnel model is tested at 45 m/s freestream velocity in Indian Institute of Science, Bengaluru lowspeed tunnel. The interaction of CM with the base flows is studied for different attitude conditions in terms angle of attack, longitudinal and lateral movements. Results indicate that CM is recovered from forebody influence beyond X/D = 2, and it experiences reverse flow till it moves to X/D = 1.4D. Stability characteristics of CM are also analysed.

Flight Performance of Crew Escape System during Pad Abort Condition

As a prologue to the Gaganyaan project, ISRO successfully test fired one of the critical and essential technologies for human space flight, i.e. Crew Escape System (CES) Pad Abort Test on 5 July 2018. This is a strategically important flight as a part of various qualification tests that ensures emergency escape measure to quickly pull the Crew Module (CM) along with the astronauts to a safe distance from the launch vehicle in case of any eventuality. CES along with CM weighing of 12.6 tonnes lifted-off at 07.00 h from the Sounding Rocket Complex Launch Pad, Sriharikota. The vehicle was taken to an altitude of 3 km using a specially developed solid motor with multiple reverse flow and scarfed nozzles with 10 g acceleration. Around 330 sensors on-board were used to measure its performance. The data obtained ensured that the adopted design philosophy and approaches such as CM reorientation, aerodynamic data with jet-on environment, mission sequences, performance of special solid motors, deceleration system, etc. were in order. This provides more confidence to progress to the next level of demonstration as a part of Gaganyaan.

Mission Design, Preflight and Flight Performance and Observations for Pad Abort Test

Abort system is initiated to take the Crew Module (CM) away from the launch vehicle in case of an emergency at lift-off or at any point of time after launch for a mission with crew onboard. Crew Escape System (CES)-based abort is carried out from launch pad and during the atmospheric phase of ascent flight. The design and operation of CES play a crucial role in providing abort capability for escape from launch vehicle and return of the crew back to Earth during critical phase of ascent flight. CES motors are used to pull CM away from the launch vehicle during this mode of abort. Mission simulation and analysis is necessary for the design of CES-based abort mission and for its configuration. This article discusses the mission design, challenges faced during the design and strategies formulated towards the successful execution of Pad Abort Test.

Structural Design, Development and Qualification Tests of Grid Fin

Crew Escape System (CES) is one of the most critical subsystems in a human-rated launch vehicle. It is an emergency escape module to rescue the Crew Module (CM) with the crew to a safe distance from the launch vehicle in an exigency at the launch pad or during atmospheric phase. Pad Abort Test (PAT) is a technology demonstration on the capability of CES to evacuate the crew in case of any exigency at the launch pad, the first in a series of tests to qualify a CES, which is a crucial technology relevant for human spaceflight. There is a wide range of altitude–Mach number regime over which CES is expected to operate and deliver the required performance. Most important among them is the requirement of aerodynamic stability for the entire Mach number regime. The configuration of CES called for an additional lift producing surfaces at its bottom to get the required stability conditions. Accordingly, four grid fins are located at the aft end of CES to move the centre of pressure down for better stability. Grid fin is a lattice structure of several aerodynamic surfaces supported by rectangular outer frames. It is designed for a variable pressure load acting on the fin elements and outer frames. Grid fins are designed and developed with composite material because of its high specific stiffness and specific strength. Composite grid fins are designed, developed, qualified and successfully flown in PAT flight. This article presents the design methodology, development, validation through structural acceptance and deployment tests of composite grid fins meant for the PAT of CES.

Structural Design, Qualification and Post-Flight Assessment of Crew Module Fairing

Structural Design, Qualification and Post-Flight Assessment of Crew Module Fairing

Design, Development and Flight Performances of Deceleration System

Human Spaceflight Programme (HSP) of Indian Space Research Organisation is proposed with the objective of carrying two crew members to low Earth orbit and bring them back safely to a predetermined location on Earth. The deceleration system for the programme has been designed for a 4-tonne class payload and shall cater to the requirements of nominal as well as abort missions. In order to finalize the parachute configurations and deployment sequence, detailed studies and development tests, starting from wind tunnel tests to full-scale air-drop tests were carried out. After successful structural and functional qualification of the parachutes and the various subsystems, the system was used to safely recover the module in the Crew Module Atmospheric Re-entry Experiment, the first unmanned spaceflight of HSP. This article provides details on the system configurations, deployment sequence and numerous tests that have been carried out till now in order to make the system worthy of manned flights in future.

Wireless Instrumentation System Experiment

This article describes the design details of a Wireless Instrumentation System that was experimented in ISRO’s recent Human in Space Project (HSP) Pad Abort Test (PAT) mission in piggy-back mode. The system consists of a few Wireless Sensor Nodes (WSNs) that acquire parameter data and a Wireless Base Station that collects these over an IEEE 802.15.4 compatible single-hop RF link and forward it to the telemetry subsystem. The circuit configuration, communication link and protocol as well as the measurement plan as adopted for HSP-PAT flight test are described to bring out the scalability of the architecture. The performance of the system in HSP-PAT mission is discussed in detail by way of PFA analysis results. All the parameters monitored through the system, including inertial ones such as acceleration and rotation were compared with reference data obtained from functional Telemetry Telecommand and Power and Navigation Guidance and Control chains and showed normal signatures. Maintaining an uninterrupted wireless communication channel in a hostile and crowded chamber like the Crew Module for all the nodes has called for a robust Medium Access Control (MAC) layer based on the industry-popular IEEE802.15.4 RF PHY. The requirement of such a robust MAC layer is established when the radio frequency links outage encountered in one of the WSN links during the interval of high vehicle dynamics is seen to be taken care of by the diverse link of the same node. A future roadmap towards self-powered wireless sensors is also outlined.

Evaluation of Candidate Thermal Protection System Materials for Indian Manned Mission Gaganyaan in Plasma Wind Tunnel Facility

During atmospheric reentry of a manned crew module, the vehicle experiences severe thermal environments due to aerodynamic heating. Protection of crew module necessitates thermal protection systems (TPS) which shield the vehicle structure from the extreme thermal environments. Ground based testing in Plasma Wind Tunnels play a vital role in selection and qualification of TPS for re-entry missions. Three candidate TPS materials identified for the Gaganyaan mission of ISRO are evaluated in Plasma Wind Tunnel facility under simulated reentry environments. Tests are conducted at expected peak heat flux and total heat load conditions and performance of materials are compared based on thermal response and material surface recession characteristics.

Особенности компоновки пассажирского и служебного блоков на речных круизных пассажирских судах с учетом модернизаций и конверсий

Особенности компоновки пассажирского и служебного блоков на речных круизных пассажирских судах с учетом модернизаций и конверсий

Design and Selection Criteria of Main Parachute for Re entry Space Payload

Parachutes are used as a decelerator in the re-entry, descent, and landing of space recovery payloads, providingstability and desired descent rate for a safe landing. The selection of the main parachute is the most critical andimportant part of the space module recovery system. Parachute size is restricted by the required landing speed,materials, and weight of the payload. Parachute materials are selected based on the various forces experienced bythe parachute. An investigation has been carried out to design a parachute system which gives less impact velocity, less angle of oscillation and less impact load for the landing of a crew module. Therefore, in this paper, selection criteria for the main parachute have been discussed considering recovery of re-entry space payload of 500 kg (unmanned) and 3500 kg (manned) class. Based on analysis carried out on the parachute size, canopy filling time, velocity reduction, peak deceleration, and opening shock, authors have proposed a unique type of solid canopy with slots (slots of the minimum area equivalent to geometry porosity) for the main parachute rather than a complex ringsail or disk-band type canopy. With this new concept, the parachute has been designed, developed and qualified through testing, trials and maiden flight of space capsule in LEO and is propose to use in the next manned space mission program.

Rapid computation of rotary derivatives for subsonic and low transonic flows

Purpose The purpose of this paper is to develop a new approach for rapid computation of subsonic and low-transonic rotary derivatives with the available steady solutions obtained by Euler computational fluid dynamics (CFD) codes. Design/methodology/approach The approach is achieved by the perturbation on the steady-state pressure of Euler CFD codes. The resulting perturbation relation is established at a reference Mach number between rotary derivatives and normal velocity on surface due to angular velocity. The solution of the reference Mach number is generated technically by Prandtl–Glauert compressibility correction based on any Mach number of interest under the assumption of simple strip theory. Rotary derivatives of any Mach number of interest are then inversely predicted by the Prandtl–Glauert rule based on the reference Mach number aforementioned. Findings The resulting method has been verified for three typical different cases of the Basic Finner Reference Projectile, the Standard Dynamics Model Aircraft and the Orion Crew Module. In comparison with the original perturbation method, the performance at subsonic and low-transonic Mach numbers has significantly improved with satisfactory accuracy for most design efforts. Originality/value The approach presented is verified to be an efficient way for computation of subsonic and low-transonic rotary derivatives, which are performed almost at the same time as an accounting solution of steady Euler equations.

Prediction Method of Blast Wave Impact on Crew Module for Liquid Rocket Explosion on Launch Pad

The role of manned space flight in the field of space exploration and utilization is growing. However, the security system of the manned spaceflight is still imperfect. In the case that the rocket explodes, crew modules maybe damaged by the blast wave, which will threaten the safety of the crews. This research aims to obtain the necessary data and information to enable the designers of the launch vehicles and crew modules to develop safer launch systems. To this end, this paper proposes a numerical method using LS-DYNA to study the propagation law of blast waves caused by rocket explosion on the launch pad and to quantify the impact of the blast wave on crew module. The numerical results indicate that the final blast waveform of the model with rocket is conical in the upper and lower parts, and spherical in the middle. At the same time, the third-stage explosion is the most harmful to the crew module, while the first-stage explosion is the least. Furthermore, the model with rocket has a marked effect on explosion strength: the pressure enhancement factor is about 4–17 times. Most importantly, overpressure prediction formula acting on the crew modulesof explosion on the launch pad is established for quick peak overpressure predicting and damage evaluating.

Thermo-structural Analysis of Solid Rocket Scarfed Nozzle with Composite Ablative Liners for Crew Escape Solid Motor

Crew escape system (CES) is one of the most critical sub-systems in a human-rated launch vehicle. CES has four different types of solid motors. One is the high-altitude escape motor (HEM). This motor has the scarfed nozzle region at the divergent aft-end side which is different from conventional nozzle to accommodate the nozzle inside the envelope of the crew module shroud. The composite ablative phenolic liners are bonded to the nozzle metallic hardware by means of adhesive. The structural integrity of this scarfed nozzle region due to temperature and pressure plays an important role for the success of the motor. The nozzle hardware of this motor comprises of three subassemblies connected together by flanged joint. Composite ablative liners are primarily designed to satisfy the thermal and internal ballistics constraint requirements. Though metallic nozzle hardware is designed to bear the complete internal pressure loads, liners share a substantial part during operation due to its stiffness which is comparable with that of the metallic backup. In addition, the high thermal gradient also results in stresses near the inner surface of the liner. These stresses cannot be estimated by closed-form solutions considering the complexity in geometry, direction-dependent material property, and arbitrary temperature distribution arising during operation. The temperature of the liner at its inner surface is the highest due to its direct contact with hot gases. The temperature within the liner decreases across thickness. It is required to be ensured that the temperature at the liner–hardware interface does not exceed the safe-operating temperature. Thermo-structural analysis of composite ablative liners is essential to estimate the complete stress state in liners and to arrive at the minimum available structural margins. The temperature and scarfed geometry make the analysis all the more complicated. Challenges in modelling liners with the 3D contact surface elements are highlighted. Varying pressure load along the inner surface of the liners is simulated in an exclusive load step, and in addition, temperature data estimated from transient thermal analysis are applied in another load step. Temperature-dependent thermo-physical and structural properties are used for the analysis. The temperature distribution across the liner thickness and stresses in the liner and metallic nozzle are reported at different cross sections and their criticality is analysed. This paper highlights the integrated 3D finite-element modelling and analysis of composite liners of solid rocket scarfed nozzle and also covers the comparison of the mechanical strain and thermal parameters with post-static test values.

Attitude Control Schemes for Crew Module Atmospheric Re-entry Experiment Mission

The Crew Module Atmospheric Re-entry Experiment mission is a suborbital mission of flying Crew Module (CM) as the payload in a parent launch vehicle. After separation from the launcher, Crew Module (CM) performs suborbital re-entry flight from an altitude of 126 km. The module is actively controlled from the instant of its separation from the launcher up to the point of re-entry into the atmosphere at 80 km after which it follows a ballistic flight through the atmosphere. Three axis control with Reaction Control System (RCS) is envisaged during exo atmospheric descent to damp out the rates and ensure near zero degree angle of attack at re-entry. This paper gives different attitude control schemes proposed for the mission using RCS thrusters. The control strategy is based on proportional derivative controller with a modulator. The design techniques for different modulation schemes are discussed. Simulations results of crew module attitude control are also presented to demonstrate the obtainable performances from different modulation schemes.

Microstructural characterization of Ti-6Al-4V X-links from the Space Shuttle Columbia

Recovered artifacts from the Space Shuttle Columbia provide a unique opportunity to examine a broad spectrum of materials exposed to the extreme environment of earth re-entry. The degradation states of these components offer a unique perspective into material behavior in this highly reactive environment. The crew module X-links, which are composed of Ti-6Al-4V and provided structural support to the crew module within the fuselage, exhibit evidence of both oxidation and combustion phenomena. The conditions supporting ignition and combustion of bulk metals are not fully understood, due to the many compounding factors affecting variable oxidation behavior of metals. Materials characterization of the Columbia X-links will yield further insight into the discrimination between the oxidation, ignition, and combustion response of titanium alloys. Widespread use of titanium alloys in space vehicles warrants a thorough characterization of their potential failure modes to help ensure the safe and reliable operation of future spaceflight vehicles. Initial microstructural state of the X-links is assessed, and affected material exhibits microstructural gradients dependent on vehicle orientation. Three distinct microstructural states were observed across the length of the X-links, according to proximity of aggravated flow.