Spaceship

28.3.1 Sources and Transport of Spacecraft Contamination

Spacecraft contamination can be defined as molecular or particulate matter on or near a spacecraft surface that is foreign to that surface. Sources of spacecraft contamination can include thruster propellants and burn residue, outgassing of spacecraft materials, vented gases from spacecraft systems, fluids released from the spacecraft by dumping or leakage, micrometeoroids and orbital debris, and particles generated or redistributed during spacecraft mechanical operations or astronaut extravehicular activity (EVA) operations (Chen, 2001). Comprehensive data on outgassing of spacecraft materials is found in Walter and Scialdone (1997). Space environment interactions with materials can also produce contaminants, such as volatile products of atomic oxygen reactions and UV-induced or radiation-induced chain-scission products in polymer materials, and residual nonoxidative films left free-standing due to atomic oxygen erosion of underlying material. Space environment effects, such as atomic oxygen, UV, and radiation interactions, can further modify contaminant species.

Spacecraft contaminants can either deposit onto spacecraft surfaces or remain in the vicinity of the spacecraft. Molecular contaminants can transport from surface-to-surface through various mechanisms, including line-of-sight transport, nonline-of-sight transport through reflection or scattering, and attraction of positively ionized contaminants to a negatively charged, sunlit spacecraft surface (Tribble, 1995). These transport mechanisms can put critical spacecraft surfaces at risk for contamination effects.

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Degradation of Spacecraft Materials

Joyce Dever, ... Sharon Miller, in Handbook of Environmental Degradation of Materials (Second Edition), 2012

24.3.1 Sources and Transport of Spacecraft Contamination

Spacecraft contamination can be defined as molecular or particulate matter on or near a spacecraft surface that is foreign to that surface. Sources of spacecraft contamination can include thruster propellants and burn residue, outgassing of spacecraft materials, vented gases from spacecraft systems, fluids released from the spacecraft by dumping or leakage, micrometeoroids and orbital debris, and particles generated or redistributed during spacecraft mechanical operations or astronaut extravehicular activity (EVA) operations.44 Comprehensive data on outgassing ofspacecraft materials is found in Ref. 45. Space environment interactions with materials can also produce contaminants, such as volatile products of atomic oxygen reactions and ultraviolet-induced or radiation-induced chain scission products in polymer materials and residual non-oxidative films left free-standing due to atomic oxygen erosion of underlying material. Space environment effects, such as atomic oxygen, ultraviolet, and radiation interactions, can further modify contaminant species.

Spacecraft contaminants can either deposit onto spacecraft surfaces or remain in the vicinity of the spacecraft. Molecular contaminants can transport from surface-to- surface through various mechanisms, including line-of-sight transport, non-line-of-sight transport through reflection or scattering, and attraction of positively ionized contaminants to a negatively charged, sunlit spacecraft surface.46 These transport mechanisms can put critical spacecraft surfaces at risk for contamination effects.

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Orbital Operations Safety

Tommaso Sgobba, ... Eugene Levin, in Safety Design for Space Operations, 2013

Current balance models

Spacecraft surface and structural charging interactions are modeled with charge balance or charge accumulation equations that describe the relationship between spacecraft voltage, Vs, relative to some reference voltage, and the magnitude of the positive (+I) and negative (–I) charge particle electrical currents impinging on the spacecraft. If both ions and electrons can contribute to spacecraft charging, the net charge accumulation on the spacecraft is zero at steady state. We therefore have:

(6)+I(Vs)+−I(Vs)=0

The steady state model embodied in eq. (6) is a useful approximation in many applications. However, in cases where the vehicle capacitance is large or during highly dynamic space flight environments, the net current cannot be neglected and electric charge of one sign or the other accumulates on the spacecraft at various rates while spacecraft voltage changes. If charged particle impingement and emission is the dominant spacecraft charging process, spacecraft voltage is a function of time and will depend on the net charging current, charged particle kinetic energy spectrum, and voltage of the target structure so that:

(7)Vs(t)=F(Vp,I−,I+)

Note that the electromotive force (measured in volts) generating the structure voltage may originate either external to (e.g., energetic charged particles) or internal to (electrical power supplies) the spacecraft. For a given net charging current, the capacitance of the spacecraft will determine the time dependence of the spacecraft voltage as shown in the section below.

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Spacecraft Configuration Design

Pasquale M. Sforza, in Manned Spacecraft Design Principles, 2016

Abstract

The spacecraft environment and its effect on design are discussed. Crew volume allowance in spacecraft is contrasted with passenger volume allowance on commercial, business, and military aircraft. Vehicle mass characteristics as evidenced by the successful manned spacecraft and the resultant ballistic coefficient are explored. Human factors in spacecraft design are assessed, and the major areas of thermal control and management of the habitable volume are described. Environmental control and life support systems including heating, ventilating and air conditioning, air and water purification, waste management, fire and emergency control systems, and communications are illustrated. Basic structural design issues, spacecraft propulsion units, and spacecraft power systems are discussed.

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Orbital Maneuvers

Howard D. Curtis, in Orbital Mechanics for Engineering Students (Third Edition), 2014

Section 6.8

6.33

Spacecraft A and B are in concentric, coplanar circular orbits 1 and 2, respectively. At the instant shown, spacecraft A executes an impulsive delta-v maneuver to embark on orbit 3 to intercept and rendezvous with spacecraft B in a time equal to the period of orbit 1. Calculate the total delta-v required.

{Ans.: 3.795 km/s}

6.34

Spacecraft A is in orbit 1, a 10,000 km radius equatorial earth orbit. Spacecraft B is in elliptical polar orbit 2, having eccentricity 0.5 and perigee radius 16,000 km. At the instant shown, both spacecraft are in the equatorial plane and B is at its perigee. At that instant, spacecraft A executes an impulsive delta-v maneuver to intercept spacecraft B 1 h later at point C. Calculate the delta-v required for A to switch to the intercept trajectory 3.

{Ans.: 8.117 km/s}

6.35

Spacecraft B and C are in the same elliptical orbit 1, characterized by a perigee radius of 7000 km and an apogee radius of 10,000 km. The spacecraft are in the positions shown when B executes an impulsive transfer to orbit 2 to catch and rendezvous with C when C arrives at apogee A. Find the total delta-v requirement.

{Ans.: 5.066 km/s}

6.36

At time t = 0, manned spacecraft a and unmanned spacecraft b are at the positions shown in circular earth orbits 1 and 2, respectively. For assigned values of θ0(a) and θ0(b), design a series of impulsive maneuvers by means of which spacecraft a transfers from orbit 1 to orbit 2 so as to rendezvous with spacecraft b (i.e., occupy the same position in space). The total time and total delta-v required for the transfer should be as small as possible. Consider earth's gravity only.

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System Assessment

Jiuping Xu, Lei Xu, in Integrated System Health Management, 2017

5.2.1 General background

MSA engineering refers to the electronic integration of every component on a spacecraft to allow for a smooth trouble free operating unit [5]. In a manned spacecraft, avionics refers to all the electronics-based instrumentation that enables essential capabilities such as communication, navigation, flight control (FC), data handling (DH), and vehicle management for the MSA [6]; therefore, as the avionics' health condition and effectiveness directly influence manned space flight safety and mission success [7], an accurate assessment of both the avionics' health condition and the subsystem's effectiveness is of great importance. To avoid and manage system failures, an ISHM-oriented hierarchical effectiveness MSA evaluation is required.

In recent years, there has been significant research focused on MSA health management. Wilkinson [8], in a discussion on prognostics for avionics remaining useful life, commented that there were two options to be considered when mechanizing the estimation of remaining life; an on-board function such as a CMC or the environmental history that could be downloaded to a ground-based maintenance facility for offline computation of remaining life. In the following year, Orsagh [9] studied PHM technologies for avionics power supplies and Banerjee [10] proposed a method based on discriminant analysis for the prognostics of an avionics system. However, little research has paid attention to evaluating effectiveness at the subsystem level from an ISHM perspective. Similarly, although some studies have made progress in evaluating the effectiveness of MSA subsystems, little importance has been attached to the overall health condition. In this study, as the test targets are selected from the MSA system-level health condition assessment [11], the MSA health condition is first assessed at the system level before an effectiveness evaluation is conducted at the subsystem level. To do this, a holistic approach is required for a system effectiveness and condition assessment that concurrently considers all multidimensional criteria and takes fuzziness and uncertainty into account.

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The Space Environment

Gerald Griffith, Tateo Goka Ph.D., in Safety Design for Space Systems, 2009

2.6.1 Introduction to the Thermal Environment

A spacecraft in proximity to a planet or a moon experiences natural environmental heating from three sources: the Sun; solar energy reflected from the planet, that is, albedo; and infrared energy emitted from the planet, that is, planetary infrared or outgoing long wave radiation. These environmental heating sources, in concert with orbital parameters, spacecraft attitude, and the design of the vehicle, determine the induced thermal environment and, hence, the thermal response of the spacecraft.

This section provides the reader with an introduction to the natural and induced thermal environments for orbiting spacecraft. Whereas the focus is primarily on Earth orbiting spacecraft, extensions to the theory provided in this section permit an understanding of the thermal environments experienced while in proximity to other bodies. Furthermore, planetary surface environments are discussed.

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Performance Evaluation

Jiuping Xu, Lei Xu, in Integrated System Health Management, 2017

4.1.1 General background

Spacecraft rely on a complex assemblage of components, all of which need to continue working to ensure operational continuance and longer mission durations. Because of the high costs and the intricacies of spacecraft structure, high LSRs are vital, especially for manned spacecraft [1,2]; therefore, spacecraft require highly reliable components and rigorous maintenance to provide the safety margins necessary to avoid mission failure. However, many factors can affect spacecraft safety and space mission success, such as the functionally graded material used in the building of the spacecraft [3–6]. Integrated system health management (ISHM), which consists of performance assessments, in situ monitoring, and fault diagnostics and prognostics, is conducted to deal with spacecraft safety requirements. To evaluate spacecraft system performance and monitor and manage system status, ISHM uses advanced algorithms and intelligent models to make full use of different data sources, with the most important being the LSR evaluation.

Four experiments are needed for the assessment of the spacecraft launch success rate; a Monte Carlo simulation based on mathematics, a hardware-in-the-loop simulation (HS), an external field test (EFT) simulation, and a real FT. Over the four assessments, as the experimental condition moves closer to reality, the test data must also be based on real situations; however, because of cost and other limitations, only a small number of experiments can be performed. Therefore, compared with general device performance assessments, assessing lunar spacecraft launch success rates has typical "high-performance, small samples, different overall" features. "High performance" refers to the need for evaluation accuracy and reliability. "Small samples" refers to the small amount or very small amount of information (such as real flight experiments) available for the experimental stages and especially for the stereotypes stage, which can reduce effectiveness when using traditional statistical theory based on frequency stability and large samples. In addition, because of the significant differences between experimental realization and cost, one small experimental data sample may differ by several orders of magnitude from another experimental data sample, with very large samples possibly dominating small samples, resulting in unreliable assessments. "Diverse populations" refers to when the experimental information obtained from the different stages and sources deviates, making the experimental data somewhat abnormal and undermining the premise of common assessment methods such as Bayes estimation theory.

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Space environment

José Meseguer, ... Angel Sanz-Andrés, in Spacecraft Thermal Control, 2012

2.1 Introduction

A spacecraft is designed to fulfil all the requirements of a given space mission. Because of that, since the very beginning of the design phase the integrity of the spacecraft must be guaranteed, as well as the correct functioning of the spacecraft itself, the different subsystems, and, obviously, the payloads.

Keeping the above in mind, it is clear that the environmental conditions where the spacecraft has to develop its mission are of paramount importance in the design process, from the first step of the design until the end of the spacecraft's operating life. These environmental conditions must include not only the extreme environment that the spacecraft will face in outer space, but also the potential sources of damage the vehicle can encounter on Earth, from the phase where the spacecraft is only a collection of materials and components to be integrated, until the launching and orbit insertion.

In general the effect of space medium can be grouped into five categories: (1) vacuum, (2) electrically neutral particles, (3) plasma, (4) radiation, and (5) micrometeoroids and orbital debris. The interactions associated with each one of these environmental conditions are summarized in Table 2.1.

Table 2.1. Space environment effect

EnvironmentEffectVacuum■

Pressure differentials

■

Sur face degradation due to solar ultraviolet radiation

■

Contamination

Electrically neutral particles■

Mechanical effects (aerodynamic drag, physical sputtering)

■

Chemical effects (atomic oxygen attack, spacecraft glow)

Plasma■

Spacecraft charging (shift in electrical potential)

■

Electrostatic discharge and dielectric breakdown

■

Enhanced sputtering

■

Re-attraction of contamination

Radiation■

Total dose effects (electronic degradation, crew safety hazards)

■

Single event effects (upsets, latch-up, burnout)

Micrometeoroids and orbital debris■ Surface damage due to hypervelocity impacts

Source: After Tribble (2003).

This book is mainly devoted to the thermal control subsystem, the task of which is to maintain the temperature of all spacecraft components, subsystems, engineering equipment, payloads and the total flight system, at safe operating and survival levels throughout the entire lifespan of the spacecraft for all flight modes. Like any other subsystem, the spacecraft thermal control subsystem is essential to ensure the reliable operation and long-term survival of any spacecraft.

Environmental effects affect the different subsystems of the spacecraft. In the case of the thermal control subsystem, the environmental thermal loads are one of the main factors that drive the design of the subsystem. Furthermore, the environmental effects may affect the performance of the subsystem through the degradation of the thermo-optical properties of the spacecraft's external surfaces once in orbit. Since the target of the thermal control subsystem is to keep the different parts of the spacecraft within their operating temperature ranges, it usually requires the evacuation of heat from dissipating parts which are prone to overheating, or to heat those parts which are too cold. The thermal control process generally requires the transfer of heat between different parts of the spacecraft, as well as between the spacecraft and outer space. Obviously, this last heat exchange becomes seriously affected when the surface's optical properties (solar absorptance, α, and infrared emissivity, ε) are modified by environmental conditions. In general, environmental effects related to conditions such as vacuum, electrically neutral particles, radiation, and micrometeoroids and orbital debris, modify the absorptance to emissivity ratio of the spacecraft's external surfaces, whereas those related to plasma affect the re-attraction of contamination.

Since the space environment can cause severe problems for space systems, great efforts have been devoted to characterize the space environment, in order to properly assess its potential effects on spacecraft. To this end, a considerable amount of technical publications on related topics dealing with space environment have been published. Furthermore, the different space agencies have developed space environment standards which aim to assist in the consistent application of space environment engineering to space products through the specification of required or recommended methods (ECSS-E-ST-10-04C, 2008).

In the next section, a brief review of the ground environment is presented. In Section 2.3 a summary of the launch environment is included. The following sections are devoted to the specific environmental aspects to be considered once the spacecraft is in orbit, including the determination of the external thermal loads on to a space vehicle.Skip to Main content

Space Shuttles

Related terms:

Energy Engineering

NASA

Space Stations

Spacecraft

Hydrogen

Astronauts

Payload

Propellant

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Overview

In Electro Hydraulic Control Theory and Its Applications Under Extreme Environment, 2019

VI Atlantis

The sixth space shuttle, Atlantis, weighed about 77 tonnes, started its first flight on 3rd October 1985 and flew 18 times. Atlantis was launched successfully at the Kennedy Space Centre on 8 July 2011 and executed a 12-day mission, which would be the last flight of the space shuttle. Atlantis landed safely at the Kennedy Space Center in Florida at 5:57 on 21 July 2011, finishing its 'curtain call tour'. Thus, the USA used the space shuttle for 30 years. Since 2011, a new generation of manned spacecraft, named Orion, designed to replace the space shuttle has been under development by NASA in conjunction with the ESA.

The design of the space shuttle is very complicated, including over 3000 important subsystems and more than 3 million components. A problem with one subsystem or a key component can cause a major accident. The crash of Columbia in 2003 was caused by a piece of cellular insulation falling from the external fuel tank and hitting the left wing of the space shuttle; this was enough to lead to tragedy. The space shuttle is much heavier than a conventional rocket, so accelerates more slowly during take-off, producing forces of about 3g, compared to 4–4.5g for rockets. The space shuttle combines the properties of rocket, satellite and plane. It can be vertically launched into space orbit like rocket, move in orbit just like a satellite, and make a glide landing in atmosphere just like a plane; it is a new type of multifunctional spacecraft. As the most complex man-made machine so far, the actual launching success rate reaches 98.5%, this is a target that most disposable rockets cannot achieve.

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Polymers for aerospace structures

In Introduction to Aerospace Materials, 2012

13.15 Case study: space shuttle Challenger accident

The space shuttle Challenger (STS-51) exploded just over one minute after take-off on 28 January 1986, killing seven astronauts. After an exhaustive investigation by NASA and other US agencies the cause of the accident was found. The space shuttle is fitted with two solid rocket boosters that generate an extraordinary amount of thrust during take-off that launches the main vehicle into space. Without the boosters the shuttle cannot generate enough thrust to overcome the gravitational pull of Earth. There is a booster rocket attached to each side of the external fuel tank, and each booster is 36 m long and 7.3 m in diameter (Fig. 13.18). The boosters are constructed from hollow metal cylinders, with the joint connecting the cylinders containing two O-rings made with an elastomer. The elastomer is needed to create a tight seal to prevent hot gases escaping from the rocket motor during take-off.

13.18. Rocket boosters on the space shuttle.

Photograph reproduced with permission from NASA.

The Challenger accident was caused by several factors, with a critical problem being that one of the elastomer O-rings in a booster rocket did not form a tight seal owing to cold weather during take-off. Elastomers shrink and lose elasticity at low temperature and, at take-off, the O-ring was unable to expand sufficiently to form a seal between two cylinders. This caused hot combustion gases (over 5000 °F) inside the rocket motor to rapidly degrade the elastomer O-ring, thus allowing hundreds of tons of propellant to escape and ignite, thereby causing the space shuttle to explode (Fig. 13.19).

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Fire Safety

Gary A. Ruff Ph.D., ... Paul T. Johnson, in Safety Design for Space Systems, 2009

Space Shuttle Detector

The Space Shuttle is designed for extended missions of typically 2-wk duration. It exhibits a large pressurized volume (74 m3), and operationally there are periods of time measured in hours when the entire crew can be asleep. In addition, during the launch sequence, many avionics systems cannot be accessed readily by the crew. Consequently, the vehicle is designed with a fire detection system and a centrally activated fire suppression system for the avionics systems. During the 1970s, when the Space Shuttle was being designed, ionization smoke detectors were becoming readily available for consumer home use, and there was extensive progress in the development of smoke detectors for terrestrial applications (Bukowski and Mulholland 1978). Ionization detectors, which used a radioactive source and whose stability was much greater than the expected lifetime of the electronics, readily were available, whereas photoelectric, that is, scattering or obscuration, detectors generally were unavailable due to the difficulty of producing stable light sources. No data were available regarding smoke particle size distributions in low gravity, and the database of normal gravity smoke characteristics was only a fraction of what is available today. Additionally, no data were available with regard to spacecraft dust particle size distributions, but the absence of gravitational settling suggested that the particles would be larger than those seen on Earth.

Given the state of knowledge at that time, it is quite reasonable that the Space Shuttle design (Figure 27.13) is an extension of the most common ground based approach, the use of ionization detectors. The Space Shuttle has nine particle ionization smoke detectors in the avionics cooling air return lines in the middeck and flight deck. Spacelab has six additional particle ionization smoke detectors in the avionics cooling air return lines (Martin and DaLee 1993). Although the design rationale for the Space Shuttle detector is not completely known, Celesco® (later Brunswick Defense®) based its design on data that suggested incipient fires could be discriminated by looking for particulate in the 0.4 to 0.7 μm range (Barr 1977). The device combined a dual chamber ionization detector with a small vane pump. The addition of a pump provided the opportunity to employ an inertial particle separator that rejected particulates larger than 1.0 μm because of their greater inertia. This focus on the smaller particulate was consistent with the understanding at the time that incipient smoke particles were smaller than 1.0 μm. Further, because the response from ionization chambers is affected by the ambient air velocity, the implementation of an ionization detector in a flow duct probably was facilitated by the use of an air pump to control the airflow through the smoke detector. This air pump increased the power requirements and reduced the operational life for the detectors.

Figure 27.13. Brunswick Defense smoke detector used in the NASA Space Shuttle fleet. The inlet is on the right, and the gas is expelled out the small plate on the top

(Courtesy of NASA).

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Space research

Marcello Lappa, in Fluids, Materials and Microgravity, 2004

1.2.5 The Space Shuttle

The Space Shuttle is a spacecraft (Fig. 1.6) that often carries special laboratories for conducting scientific research. The Shuttle can stay in orbit around the Earth for up to 17 days, which is ideal for experiments that need longer periods in microgravity to be successful.

Fig. 1.6. The Space Shuttle.

While a free-fall reduces the effects of gravity, being in an orbiting laboratory introduces, however, other accelerations that cause effects that are indistinguishable from those due to gravity. When a spacecraft is in orbit around the Earth, the orbit is actually defined by the path of the center of mass of the spacecraft around the center of the Earth. Any object in a location other than on the line traversed by the center of mass of the spacecraft is actually in a different orbit around Earth. Owing to this, all objects not attached to the spacecraft move relative to the orbiter center of mass. Other relative motions of unattached objects are related to aerodynamic drag on the vehicle and spacecraft rotations. A spacecraft in a low-Earth orbit experiences some amount of drag due to interactions with the atmosphere. An object within the vehicle, however, is protected from the atmosphere by the spacecraft itself and does not experience the same deceleration that the vehicle does. The floating object and the spacecraft therefore are moving relative to each other. Similarly, rotation of the spacecraft due to orbital motion causes a force to act on objects fixed to the vehicle but not on objects freely floating within it. On average for the Space Shuttles, the quasisteady accelerations resulting from the sources discussed above (position in the spacecraft, aerodynamic drag, and vehicle rotation) are on the order of 1 × 10−6 g, but vary with time due to variations in the atmospheric density around the Earth and due to changes in the Shuttle orientation. In addition to these quasisteady accelerations, many operations on spacecraft cause vibrations of the vehicle and the payloads (experiment apparatus). These vibrations are often referred to as "g-jitters," because their effects are similar to those that would be caused by a time-varying gravitational field. Typical sources for vibrations are experiment and spacecraft fans and pumps, motion of centrifuges, and thruster firings. With a crew onboard to conduct experiments, additional vibrations can result from crew activities. The combined acceleration levels that result from the quasisteady and vibratory contributions are generally referred to as the microgravity environment of the spacecraft. On the Space Shuttles, the types of vibration-causing operations discussed earlier tend to create a cumulative background microgravity environment of about 1 × 10−4 g, considering contributions for all frequencies below 250 Hz.

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Metal matrix, fibre–metal and ceramic matrix composites for aerospace applications

In Introduction to Aerospace Materials, 2012

16.7 Case study: ceramic matrix composites in the space shuttle orbiter

The space shuttle orbiter is the world's first, and so far only, reusable spacecraft. (The Russian built Buran was developed in the 1980s as a reuseable vehicle similar to the space shuttle which performed just one unmanned space mission). The construction of the orbiter is similar to a conventional airliner; the body is built mostly of high-strength aluminium alloys and the payload doors are fibre–polymer sandwich composite material. Although speciality materials are used in highly stressed components of the mainframe, such as titanium and metal matrix composite, most of the structure is built with the same aluminium alloys used in civil and military aircraft. Chapter 3 provides more information on the structural materials used in the orbiter.

The orbiter is covered with ceramic tiles that protect the vehicle, payload and crew from the extreme heat generated during re-entry into the Earth's atmosphere. Figure 16.9 shows the temperature profile over the orbiter surface during re-entry; the nose and leading edges are heated to 1000–1400 °C whereas the maximum temperatures of other sections are 200–1000 °C. The properties of aluminium and polymer matrix composite demand that the orbiter's structure is kept below 150–200 °C, and therefore the heat insulation tiles are essential.

16.9. Temperature surface profile of the space shuttle orbiter during re-entry.

The orbiter is covered with over 25 000 reusable ceramic matrix composite tiles (Fig. 16.10). All the tiles are brittle and can crack when stressed or impacted, as tragically proven when the Columbia broke up during re-entry on flight STS-107 (February 2003). Chapter 18 describes the impact fracture of the carbon–carbon tiles that caused the Columbia incident. Because the aluminium structure of the orbiter expands and contracts owing to temperature changes over the course of a flight mission, the tiles are not mounted directly onto the skin. Instead, compliant adhesive felt pads are used to bond the tiles to the aluminium.

16.10. Heat insulation tiles on the space shuttle orbiter where the black regions are carbon–carbon composite and white regions are mostly high-purity silica ceramic. RCC is reinforced carbon–carbon composite; HRSI is high-temperature reusable surface insulation; LRSI is low-temperature reusable surface insulation; FRSI is fibrous reusable surface insulation.

The forward nose cap and leading edges of the wings are covered with carbon–carbon composite tiles. These tiles are coated with black silicon carbide for oxidation resistance, with the black colour helping to radiate heat during re-entry. Carbon–carbon tiles are used in the hottest regions where the temperature exceeds about 800 °C. The cooler regions, where the temperature is 200–800 °C are covered mostly with white ceramic tiles that reflect solar radiation to keep the space shuttle cool. The tiles consist of porous high-purity silica ceramic covered with borosilicate glass. Two other types of silica-based tiles are used on the orbiter: fibrous refractory composite insulation (FRCI) or toughened unipiece fibrous insulation (TUFI). FRCI is used in a few selected regions whereas the TUFI is applied over the extreme back of the orbiter near the engines.

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Fleet impact resulting from a space shuttle Columbia main engine controller wire failure during Mission STS-93

Steven J. McDanels, in Handbook of Materials Failure Analysis with Case Studies from the Aerospace and Automotive Industries, 2016

1 Space Shuttle Columbia Wiring Hardware Overview

The space shuttles each had over 200 mile of wiring, weighing over two-and-a-half tons, in addition to associated cables, conduits, and trays as well as hardware to house, route, and contain all of the materials. Although the initial investigation focused on only several inches of the wiring, the ramifications extended to all 200-plus miles of wiring aboard the Columbia, and impacted the entire Shuttle fleet [2].

The subject wire provided power to the digital control units and was a 14 American wire gauge (AWG) polyimide-insulated twisted three-wire conductor, with Kapton® polyimide insulation surrounding the nickel-plated copper conductor. The insulation was topcoated with an aromatic polyimide resin color coded to indicate AWG size, with green corresponding to 14 AWG and red to 20 AWG. Although the proximate cause of the loss of signal redundancy was found in the form of a short circuit caused by the wire arcing to an adjacent screw, the root cause of the arcing event had to be ascertained.

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Thermal protection systems

José Meseguer, ... Angel Sanz-Andrés, in Spacecraft Thermal Control, 2012

17.3.1 Design example: Space Shuttle

The Space Shuttle thermal protection system was based on the use of surface materials with a high temperature capability, in combination with an underlying thermal insulation to reduce heat conduction to the interior of the vehicle. The heat from the aerodynamic heating was thereby radiated back into space. Furthermore, the aim of the Space Shuttle thermal protection system was to protect the vehicle not only from the hot re-entry, but also from the extremely cold conditions experienced during the night phase of each orbit. A design life of one hundred missions without major refurbishment was also expected.

Due to the difference of temperature on the surface of the vehicle, ranging from 600 to 1750 K, four different materials were selected (Dotts et al., 1983): reinforced carbon-carbon (RCC), high-temperature reusable surface insulation tiles (HRSI), low-temperature reusable surface insulation tiles (LRSI), and felt reusable surface insulation blankets (FRSI). A map of the distribution of the materials on the vehicle surface is shown in Figure 17.4.

Figure 17.4. Space Shuttle thermal protection system

Key: 1 – reinforced carbon-carbon; 2 – high-temperature reusable surface insulation; 3 – low-temperature reusable surface insulation; 4 – coated Nomex felt; 5 – metal or glass.

Source: After Dotts et al. (1983).

Reinforced carbon-carbon is a light grey all-carbon composite used where the temperature exceeds 1500 K (this is on the leading edges of wings and the nose cap). The reinforced carbon-carbon material is highly resistant to fatigue loads, has sufficient strength to withstand launch and re-entry aerodynamic loads, and a low coefficient of thermal expansion, which provides it with excellent resistance to thermal stresses and shock. Although strong, and capable of withstanding extremely high temperatures, reinforced carbon-carbon is also thermally conductive. This requires the use of insulation blankets and tiles behind the reinforced carbon-carbon panels to protect the vehicle structure and attach fittings from heat radiated from the rear side. Each wing leading edge had 22 reinforced carbon-carbon panels with Inconel foil-wrapped ceramic insulators protecting the metallic attach fittings. The nose cap was a reinforced carbon-carbon monoconic shell with a combination of Nextel/silica fibre blankets and internal tiles to protect the area behind. Reinforced carbon-carbon was also used in the arrowhead area at the forward external tank attach point, for shock protection during pyrotechnic separation.

The Space Shuttle was equipped with about 24 000 reusable surface insulation tiles. These tiles were used for areas with temperatures ranging from 900 to 1500 K, whereas low-temperature reusable surface insulation tiles were used for areas with temperatures ranging from 650 to 900 K. Both were made of the same base material but with different coatings. Most tiles were made of a material called LI-900, 99.9% pure silica glass fibres. The material was designed to minimize thermal conductivity and weight, which compromised its strength. Thus, for high stress areas, a higher strength version of the material, LI-2200, was used. In 1981, a new material, FRCI-12, with 22% of Nextel fibres, was introduced to replace LI-900 and LI-2200 tiles, and in 1996 NASA introduced a fourth material, the AETB-8 (alumina strength thermal barrier), which improved the properties of the existing materials by adding small quantities of alumina. Other upgrades were the replacement of most low-temperature reusable surface insulation tiles on the upper surface with fibrous insulation blankets (FIB), and the use of toughened unipiece fibrous insulation (TUFI) tiles.

The reaction-cured glass coatings were made from blended glass powders mixed with thickeners and pigments. The coatings were sprayed, dried, and then fired in a kiln at 1200 °C for 90 minutes. The coatings were between 0.25 and 2.5 mm thick. High-temperature reusable surface insulation tiles were coated with a black borosilicate glass coating that provided high emissivity. The low-temperature reusable surface insulation tiles had a white borosilicate glass coating, with a low solar absorptance for in-orbit thermal control.

Felt reusable surface insulation blankets were nylon felt blankets coated with a white silicone rubber. They were used to protect areas exposed at temperatures below 650 K.

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Atmospheric Entry Mechanics

Pasquale M. Sforza, in Manned Spacecraft Design Principles, 2016

6.6.4 Moderate L/D Entry: Space Shuttle Orbiter

The Space Shuttle Orbiter follows a trajectory that yields a smooth velocity variation with altitude with no skip like that followed by the Apollo capsule. A plot of the trajectory for a typical entry of a Space Shuttle Orbiter from LEO is shown in Figure 6.41. In this case, the nominal trajectory for Space Transportation System STS-5 is depicted. The entry velocity, flight path angle, and altitude are Ve=7.7 km/s, γe=−1.5°, and ze=120 km, respectively. The altitude drops rapidly at essentially constant velocity until the vehicle reaches an altitude of about 80 km (50 mi) after which time the drop in velocity and altitude is more gradual. During the descent, the Space Shuttle Orbiter is pitched through decreasing angles of attack α(t), while also rolling back and forth through a series of different bank angles ϕ(t) as shown in Figure 6.42. These pitching and rolling maneuvers are used to modulate the effective lift and drag so as to control the reentry behavior. Banking back and forth modulates the lift in the vertical plane, while keeping the average side force approximately zero so that the vehicle remains on a trajectory lying within a vertical plane.

Figure 6.41. Plot of a nominal Space Shuttle Orbiter entry trajectory (STS-5) showing the variation of speed, altitude, and angle of attack as a function of time.

Figure 6.42. Plot of a typical Space Shuttle Orbiter entry trajectory showing the nominal variation of Orbiter bank angle ϕ, angle of attack α, and altitude z as a function of time.

Along this path, the dynamic pressure, the deceleration, and the convective heat transfer, as calculated from the STS-5 nominal trajectory data according to the equations of the previous sections, all continually increase from essentially zero to a maximum and then decrease again, as shown in Figure 6.43. Note that the three variables reach their maxima at different times and therefore at different altitudes in the typical mission shown. That the nominal entry trajectory of STS-5 shown in Figure 6.41 is one which lies within the acceptable corridor for manned spaceflight is illustrated in Figure 6.44.

Figure 6.43. The nominal variation of stagnation point heat transfer rate, deceleration, and dynamic pressure is shown as a function of time for the typical Space Shuttle Orbiter entry trajectory of Figure 6.41.

Figure 6.44. The nominal trajectory of STS-5 is shown to lie within the dynamic pressure corridor considered acceptable for manned spaceflight.

Because the Space Shuttle Orbiter is actually a spaceplane, it has aerodynamic characteristics very much like an airplane. That is, the lift and drag are functions of angle of attack α, and, as mentioned in the previous section, the lift in the vertical plane is a function of the angle of bank ϕ as well. The lift and drag characteristics are described in some detail in Chapter 8 and are summarized for hypersonic flight (M>5) in Figure 6.45. Obviously there is a substantial variation in lift and drag coefficients over the range of angle of attack covered by STS-5. In addition, the history of the bank angle of STS-5 is not in hand and changes in lift may also occur in addition to those due to angle of attack.

Figure 6.45. Approximate lift and drag characteristics for the Space Shuttle Orbiter in hypersonic flight (M>5).

However, we may use the STS-5 trajectory results to assess the behavior of numerical solutions to the equations developed thus far. In Figure 6.46, the nominal velocity history for STS-5 shown in Figure 6.41 is compared to numerical solutions of Eqns (6.34)–(6.36) for the simple case in which constant lift and drag coefficients are assumed to exist throughout the flight. Using a nominal mass m=90,000 kg, a reference area S=250 m2, and a drag coefficient CD=0.86 yields a normalized B′=mgE/CDS=0.0406 atm which corresponds to flight at α=40°, which is the case for over half the flight time, as can be seen in Figure 6.42. Using this constant value of B′ along with an assumed constant value L/D in Eqns (6.34)–(6.36) leads to the results shown in Figures 6.46 and 6.47 where it is seen that the case of B′=0.0406 and L/D=0.9 provides a reasonably good fit to the nominal flight data for velocity and altitude as a function of time from entry.

Figure 6.46. Comparison of the velocity history of a typical Space Shuttle Orbiter (STS-5) entry trajectory with numerical solutions for constant values of L/D and B′=0.0406 atm.

Figure 6.47. Comparison of the altitude history of a typical Space Shuttle Orbiter (STS-5) entry trajectory with numerical solutions for constant values of L/D and B′=0.0406 atm.

Note that the computed velocity and altitude histories in Figures 6.46 and 6.47 exhibit a long-period phugoid-type motion at high altitudes. This is characteristic of computations for high L/D (>1) vehicles entering the atmosphere at small angles. Damping due to increasing density reduces these oscillations as lower altitudes are reached, as shown in Figure 6.46. Flight path angle measurements were not available for STS-5 but the general behavior is as illustrated in Figure 6.48.

Figure 6.48. Comparison of the computed flight path angle history for constant values of L/D and B′=0.0406 atm.

Although this simple approach is instructive, there is no method for choosing appropriate constant values for B′ and L/D. If we use the angle of attack history for STS-5 shown in Figure 6.41, along with the curve fits for the approximate lift and drag characteristics shown in Figure 6.45, in Eqns (6.34)–(6.36), we obtain the velocity and time histories shown in Figures 6.49 and 6.50, respectively. There are two cases considered, one with a zero bank angle and one with a variable bank angle. It is clear from Figures 6.49 and 6.50 that the case with a zero bank angle leads to considerably higher velocities and altitudes than those for STS-5. This result suggests that the lift produced is too high leading to excessive entry time and too great range. However, if the lift is reduced by reducing the angle of attack the drag will also be reduced and again the velocity will be increased, so some other means of reducing lift is necessary. Rolling the vehicle such that the lift in the vertical plane of the trajectory is reduced by the cosine of the bank angle provides such lift control. A typical such maneuver for the Space Shuttle Orbiter was shown in Figure 6.42.

Figure 6.49. Computed velocity history for the STS-5 vehicle assuming a variable bank angle and a bank angle of zero compared to the nominal STS-5 data.

Figure 6.50. Computed altitude history for the STS-5 vehicle assuming a variable bank angle and a bank angle of zero compared to the nominal STS-5 data.

For illustration purposes, the simpler maneuver shown in Figure 6.51 is suggested. The corresponding CL and CD characteristics are illustrated in Figure 6.52. The spacecraft response to angle of attack variation with and without roll modulation is shown in Figures 6.49 and 6.50. It is clear that the roll maneuver works to reduce the flight time even though the angle of attack variation is fixed. It should also be recalled that the STS-5 trajectory is being used for illustration because the angle of attack history is available. The actual roll maneuvers of STS-5 were not available and are not used here, being more complicated, like the typical variation shown in Figure 6.42.

Figure 6.51. Notional banking maneuver to reduce lift coefficient at fixed angle of attack.

Figure 6.52. Variation of CL and CD for the notional rolling maneuver shown in Figure 6.51.

The range covered with and without roll modulation of the lift is shown in Figure 6.53. It appears clear that the roll maneuver has essentially put a time shift into the trajectory without changing its character. This may also be seen in the history of the flight path angles shown in Figure 6.54 which display a behavior similar to that seen for the constant L/D trajectories in Figure 6.48.

Figure 6.53. Range as a function of spacecraft altitude with and without roll modulation of lift.

Figure 6.54. Comparison of the computed flight path angle history with and without roll modulation of lift.

Also worth noting again is that the computed trajectories in Figures 6.47, 6.50, and 6.53 display a skip-like behavior unlike the STS-5 flight data in Figure 6.50. As mentioned previously, this arises because Eqns (6.34)–(6.36) involve small differences of small numbers, particularly at high altitudes (the density varies by eight orders of magnitude over the flight path), which results in details of the flight characteristics being magnified. Thus, though the computed flight path angles remain small through most of the flight, Figure 6.54 shows that the sign of the angle fluctuates resulting in the phugoid motion observed in the calculated trajectories. This is best appreciated by first comparing the altitude variation of velocity with and without roll modulation of the lift to that for the STS-5 data shown in Figure 6.55. The computed velocity exhibits the waviness described previously, particularly above an altitude of 60 km. The roll maneuver takes place between 60 and 70 km altitude, and below 60 km the velocity–altitude behavior of the spacecraft with and without roll modulation of the lift is essentially the same. The computed results do not closely match the STS-5 data because, although its angle of attack history was used, its actual roll history is not known and the simple roll maneuver described previously was substituted for illustrative purposes.

Figure 6.55. Altitude variation of spacecraft velocity with and without roll modulation compared to that for the STS-5 data.

This variation of computed altitude coupled to the exponential nature of the atmosphere leads to corresponding oscillations in computed flight characteristics like stagnation point convective heat flux, dynamic pressure, and acceleration level, as illustrated in Figures 6.56–6.58, respectively. The convective heat flux depends upon ρ1/2V3, as described in Section 6.4, and therefore is sensitive to the waviness in the trajectory data. As can be seen in Figure 6.56, the variations in convective heat flux are most noticeable in the altitude range 60 km

Figure 6.56. Altitude variation of stagnation point convective heat flux with and without roll modulation compared to that computed from the STS-5 data.

Figure 6.57. Altitude variation of dynamic pressure with and without roll modulation compared to that computed from the STS-5 data.

Figure 6.58. Altitude variation of normalized acceleration with and without roll modulation compared to that computed from the STS-5 data.

The peak dynamic pressure for the STS-5 data occurs at an altitude well below that for which oscillations in the computed values of q appear, as seen in Figure 6.57. The computed values of q do not show a peak and their magnitude is about three times that of the STS-5 data because the computed velocity is not in very close agreement with the flight data as shown in Figure 6.55. The peak value of dynamic pressure in the STS-5 data occurs in the supersonic range, M≤3, where the lift and drag coefficients depart from the hypersonic values used in the computations. In Figure 6.58, the computed acceleration levels show the altitude variations discussed previously but are all within the range of the STS-5 data which has a maximum of a/gE~−1. The Space Shuttle Orbiter was designed to carry scientific specialists who were not necessarily pilots trained for and capable of tolerating the relatively high deceleration levels common to previous flights with space capsules. The lifting capability of a spaceplane was crucial to providing such a benign environment for the crew, although very small entry angles were still necessary. As a result, the calculation of the entry trajectory of a spaceplane is dependent upon accurate lift and drag information for use throughout the flight.

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Materials and material requirements for aerospace structures and engines

In Introduction to Aerospace Materials, 2012

3.4 Space shuttle structures

The space shuttle is a complex system consisting of an external fuel tank, two solid rocket boosters and the Space Transportation System (STS) orbiter vehicle. In this section, we only examine the structure and materials of the orbiter. The orbiter resembles a conventional aircraft with double-delta wings, and uses many of the same materials. The orbiter is divided into nine major structural sections (Fig. 3.12). Most of the sections are constructed like a passenger airliner using aircraft-grade aluminium alloys. The major structural assemblies are connected and held together by rivets, bolts and other fasteners, again much like an airliner. However, some materials used in the space shuttle are unique, and are not found in fixed- or rotary-wing aircraft. One distinguishing feature of the orbiter is the reusable thermal insulation system. Over 25 000 ceramic and carbon–carbon composite tiles, that can withstand temperatures of about 1200 °C and above 2000 °C, respectively, are used to insulate the underlying structure during re-entry.

3.12. Main sections in the space shuttle orbiter.

The forward fuselage section is robustly designed to carry the high body bending loads and nose gear landing loads. The body skin panels, stringers, frames and bulkheads in the forward section are made with the same aluminium alloy (2024 Al) found in conventional aircraft structures. The windows are made using the thickest ever pieces of optical quality glass. Each window consists of three individual panes: the innermost pane is 15.9 mm (0.625 in) thick tempered aluminosilicate glass whereas the centre and outer panes are 33 mm (1.3 in) and 15.9 mm fused silica glass. This design can withstand the extreme heat and thermal shock during re-entry when temperatures reach 600–700 °C.

The mid-fuselage section is the 18.3 m (60 ft) long structure that interfaces with the forward and aft fuselage sections and the wings. The mid-fuselage includes the wing carry-through structure, which is heavily loaded during reentry, and the payload bay (including its doors). The fuselage is constructed with monolithic and honeycomb sandwich panels of aluminium, which are stiffened with load-bearing vertical and horizontal frames. The frame is constructed using 300 struts of metal matrix composite (boron fibre/aluminium tubes), which has exceptionally high stiffness and provides a weight saving of 45% compared with a conventional aluminium construction. The payload bay doors are a sandwich composite construction (carbon fibre–epoxy skins and Nomex core) with carbon-fibre composite stiffeners. This construction reduces the weight by over 400 kg (900 lb) or 23% compared with an aluminium honeycomb material.

The aft fuselage consists of an outer shell, thrust section and internal secondary structure, and it supports the manoeuvring/reaction control systems pods, main engines and vertical tail. The aft fuselage skins are made of aluminium alloy reinforced with boron fibre-epoxy composite struts. These struts transfer the main engine thrust loads to the mid-fuselage and external tanks during take-off. At take-off the two solid rocket boosters generate a combined thrust of 25 MN (5.6 million lb), which is over 200 times the twin engine thrust of a Boeing 737. Owing to the extreme thrust, titanium alloy strengthened with boron–epoxy struts is used near the engines.

The wing and vertical tail is constructed mostly with aircraft-grade aluminium alloy. The outboard wing section is made with high temperature nickel honeycomb sandwich composite and the inboard wing section of titanium honeycomb. The elevons, used for vehicle control during atmospheric flight, are constructed of aluminium honeycomb.

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Corrosion control in space launch vehicles

L.M. Calle, in Corrosion Control in the Aerospace Industry, 2009

9.2.3 Orbiter flight and ground environment

The Space Shuttle experiences corrosion environments from benign to severe as it progresses from flight to flight. The relative severity of each environment is illustrated in Fig. 9.2. At KSC, the orbiters are stored in the Orbiter Processing Facility (OPF) under temperature and humidity control. OPF temperature is maintained at 21.1 ± 2.8 °C. Relative humidity is 50% maximum. The OPF is the prime facility for orbiter processing: preflight and post landing. Access is provided to all external and limited internal surfaces of the orbiter to perform the following operations: draining and purging all fuel systems, ordinance removal, repair and replacement of damaged components, inspection and refurbishment of the Thermal Protection System (TPS), inspection and testing of orbiter systems (landing gear, main and auxiliary propulsion, power units, flight instrumentation, communications and orbiter hydraulics), payload bay (configuration and testing), and payload installation, connection, and removal. The orbiter is towed a short distance from the OPF to the adjacent Vehicle Assembly Building (VAB) where it is mated to the external tank (ET), which is already mated with the SRBs on the mobile launcher platform (MLP). Unlike the OPF, the VAB is not under temperature and humidity control but it offers protection from wind, rain, salt spray, and sunlight. When moved to the launch pad, the orbiter is subjected to an almost constant salt spray from the nearby Atlantic Ocean. The high humidity allows the formation of condensation on all surfaces open to the atmosphere. Once the orbiter reaches low earth orbit, any water that may have collected during earlier exposure evaporates in the vacuum of space. Corrosion is not a concern in space. However, the environment of space can have detrimental effects on materials due to the presence of high energy protons and electrons, ultraviolet radiation, atomic oxygen, high and low temperature extremes, high vacuum, galactic cosmic radiation, micro meteors, and man-made debris. These effects must be considered in selecting materials and corrosion protection finishes. As part of NASA's Materials International Space Station Experiment (MISSE), hundreds of samples representative of a variety of materials, including coatings, have been attached to the exterior of the International Space Station for long-term periods of exposure and brought back to earth for evaluation.5 On landing, the orbiter is then exposed to the arbitrary environment of the landing site. If the orbiter lands in California, it must be ferried across the country atop the Shuttle Carrier Aircraft (SCA) to KSC. Until the orbiter returns to the OPF, this is additional exposure to an uncontrolled environment. All these factors were taken into account when the overall corrosion protection scheme for the orbiter was developed.

9.2. Variation of orbiter corrosion environment during typical flow.

The Space Transportation System (STS) consists of the orbiter spacecraft, two SRBs for launch, and an ET for the liquid hydrogen and oxygen fuels which feed the three main engines housed in the orbiter aft fuselage. STS processing includes a scheduled period at the pad for payload installation, final servicing, and checkout before launch. The average pad stay is approximately 31 days. On occasion, problems with the flight hardware have extended the time normally scheduled at the pad. The longest time an STS has spent at the pad was 166 days before Space Shuttle Columbia's tenth flight. Extended stays are of concern because the pad is located a few hundred meters from the ocean. The coastal exposure is very severe due the heat, high humidity, salt air, and the daily condensation of dew deposited onto the orbiter structure

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