Materials used in Space Re-entry Vehicles
Seminar Report submitted by
Name:
Reg. Number:
Course code:
Semester/Batch:
Mentor:
Likith S
18ETAS012017
19ASP401A
7/2018
Dr Shashank Vadlamani
B. Tech in Aerospace Engineering
Department of Aerospace Engineering
Ramaiah University of Applied Sciences
University House, Gnanagangothri Campus, New BEL Road,
M S R Nagar, Bangalore, Karnataka, INDIA - 560 054
Declaration Sheet
Student Name
LIKITH S
Reg. No
18ETAS012017
Programme
B. Tech. ASE
Course Code
19ASP401A
Course Title
Seminar
Course Date
20th August 2021
Batch 2018
to
1st December 2021
Declaration
The seminar report submitted herewith is a result of my own investigations and that I
have conformed to the guidelines against plagiarism as laid out in the Student
Handbook. All sections of the text and results, which have been obtained from other
sources, are fully referenced. I understand that cheating and plagiarism constitute a
breach of University regulations and will be dealt with accordingly.
Signature of the
Date
student
Name
10/12/2021
Signature
Date
First Examiner
Second Examiner
Mentor
19ASP401A-Seminar
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Contents
Declaration Sheet............................................................................................................ i
LIST OF ABBREVIATIONS AND NOMENCLATURE .......................................................... iii
LIST OF FIGURES ............................................................................................................ iv
ABSTRACT ....................................................................................................................... 5
1. Introduction and scope of work................................................................................. 6
1.1 Introduction: .............................................................................................................. 6
2. Literature Review ....................................................................................................... 7
2.1 Literature Survey........................................................................................................ 7
3. Details of the Topic .................................................................................................... 9
3.1.1 Re-entry Vehicles .................................................................................................... 9
3.1.2 Thermal Protection System .............................................................................. 11
3.1.3 Requirement of Thermal Protection System .................................................... 11
3.1.4 Required Material Properties for TPS ............................................................... 12
3.2 Thermal Protection System Materials and their properties .................................... 13
3.2.1. Reinforced Carbon Carbon: ............................................................................. 13
3.2.2. High temperature reusable surface insulation HRSI tile ................................. 14
3.2.3. Low temperature reusable surface insulation LRSI tile ................................... 14
3.2.4. High-Efficiency Tantalum-based ceramics (HETC) composite ......................... 15
3.2.5. Phenolic Impregnated Carbon Ablator, PICA................................................... 16
3.2.6. Tantalum Carbide and Hafnium Carbide: ........................................................ 17
4. Challenges and Opportunities ................................................................................. 19
4.1 Challenges during Re-entry .................................................................................. 19
4.1.1 Challenges in Selection of Materials ................................................................. 20
6. Conclusion and Suggestions for Future Work.......................................................... 21
6.1 Conclusions .............................................................................................................. 21
6.2 Suggestions for Future Works .............................................................................. 21
REFERENCES ................................................................................................................. 22
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LIST OF ABBREVIATIONS AND NOMENCLATURE
TPS
Thermal Protection Systems
RCC
Reinforced Carbon Carbon
PICA
Phenolic Impregnated Carbon Ablator
HRSI
High Temperature Reusable Surface Insulation Tiles
LRSI
Low Temperature Reusable Surface Insulation Tiles
HETC
High Efficiency Tantalum Based Ceramics Composites
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LIST OF FIGURES
Figure 1 Atmospheric Re-entry Demonstrator Credits: CNES artist's concept ................... 6
Figure 2 Soyuz Descent Module reaches Entry Interface, where friction from Earth's
thickening atmosphere heats its outer surfaces. Credits: NASA ......................................... 9
Figure 3 China’s Shenzhou manned spacecraft and its reentry Credits: Enabling
Technologies for Chinese Mars Lander Guidance System, Xiuqiang Jiang ........................ 10
Figure 4 RCC used in leading edge of space shuttle orbiter .............................................. 13
Figure 5 HRSI Tile used in Landing Gear doors and some parts of orbiter under surface
Credits: Space Shuttle Tiles NASA ...................................................................................... 14
Figure 6 LRSI Tile used in some upper surface of space shuttle Credits: Orbiter TPS, NASA
............................................................................................................................................ 15
Figure 7 PICA formation process Credits: [3] ..................................................................... 16
Figure 8 Mars Science Laboratory (MSL) 4.5m PICA Heat Shield ...................................... 17
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ABSTRACT
Re-entry vehicle is a portion of the spacecraft which returns back to earths or another
planet atmosphere. A Re-entry vehicle could be a rocket, satellite, or a manned capsule.
When returning to earth or when landing on another planet, a safe reentry through the
atmosphere is needed. Safe re-entry can be difficult, because the very high speed of the
spacecraft creates very high temperatures, when entering through the atmosphere.
Due to very high temperature the material used in re-entry vehicle must be able to
withstand high temperature around 1260 ℃ and materials must have excellent heat
resistance in non-oxidizing atmosphere.
Thermal protection systems (TPS) are designed to protect re-entry space vehicles from
the severe heating encountered during hypersonic flight through a planets or the earth’s
atmosphere. A secondary goal is to protect from the heat and cold of space while on orbit.
The thermal protection system consists of various materials applied to the outer surface
of the orbiter to protect the orbiter/re-entry vehicle at extreme temperatures, primarily
during the re-entry into the atmosphere. These materials are the last defense before the
aluminum and graphite epoxy shell.
During re-entry, the TPS materials perform in temperature ranges from -250 ℉ in the cold
soak of space to entry temperatures that reach nearly 3,000 ℉. Because the thermal
protection system is installed on the outside of the orbiter skin, it establishes the
aerodynamics over the vehicle in addition to acting as the heat shield. The TPS is a passive
system consisting of materials selected for stability at high temperatures and weight
efficiency.
In this report firstly materials used in re-entry vehicles, Thermal protection systems for reentry vehicles. Also, in this report the challenges in selection of materials and difficulties
during re-entry is discussed. Secondly the requirement of thermal protection system and
advantages of materials used in re-entry vehicle is discussed. Finally, the most suitable
material for re-entry vehicle is suggested.
Keywords: Re-entry, Thermal Protection System, Aerodynamic Heating, Materials,
Atmospheric Drag
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1. Introduction and scope of work
1.1 Introduction:
Atmospheric re-entry is the entry of an object from the outer space to the earth
atmosphere or atmosphere of outer planet. The portion of a spacecraft which returns
back to earth is the re-entry vehicle. During atmospheric re-entry the major problem faced
is aerodynamic heating and atmospheric drag. Space craft, missiles, satellite and space
probe travel at high velocities several times than speed of sound (Mach No greater than
5). Increase in velocity decrease the thickness of the velocity boundary layer. This results
in the velocity gradient near the wall and hence the viscous stresses decrease. The kinetic
energy of the fluid stream results in an increase in internal energy of the fluid near the
wall. This results in increase in the temperature of fluid near the wall and increase heat
transfer to the wall. This is called aerodynamic heating.
Figure 1 Atmospheric Re-entry Demonstrator Credits: CNES artist's concept
Due to skin friction, on the surface of the body, the kinetic energy is converted into heat.
The gravity of earth is an important factor which causes major physical changes and
physical phenomenon to the vehicle. The natural thermal resistance depends upon the
shape of the re-entry vehicle. While comparing streamlined and blunt body concepts it is
clear that blunt body offers natural thermal protection to a large extent. Since re-entry is
at hypersonic speeds the vehicle should not be streamlined. During re-entry high heat
energy is produced by the vehicle because of high speed and the lethal to the vehicle. The
shape of the re-entry vehicle should not be streamlined because the ballistic coefficient
of streamlined body is very high. So, the shape should be blunt, because ballistic
coefficient is low for blunt body.
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2. Literature Review
2.1 Literature Survey
Ahana Arshad S, Dinu Vijay, Sajith R, Jincy J C (2019) [1] “Analysis of materials used for reentry vehicle”. Ahana Arshad reviewed about the different types of materials used in reentry vehicle and their temperature withstand capacity. Also, Arshad discussed about
various types of materials that can be used in re-entry vehicle and their advantages. Also,
Ahana Arshad reviewed about thermal protection system and high temperature coatings
for re-entry vehicle. Finally, Ahana Arshad concluded that superalloys and nanomaterials
are best for re-entry vehicle.
Sylvia M. Johnson (2015) [2] “Thermal Protection Materials and Systems: Past and Future”
40th International Conference and Exposition on Advanced Ceramics and Composites
Sylvia M. Johnson discussed about the thermal protection system, reusable and ablative
materials. Johnson also reviewed about the new trends in thermal protection system.
Ryan Oakes “Space Shuttle Ceramic Tiles” [3]. Oakes discussed about the various surface
insulation tile used in space shuttle orbiter and also reviewed about RCC and surface
insulation blankets.
J. Thornton, W. Fan “PICA Variants with Improved Mechanical Properties” [4]. J. Thornton
and W. Fan reviewed about phenolic impregnated carbon ablator and how to improve the
mechanical properties of the existing PICA. J. Thornton and W. Fan focused on improving
existing low-density PICA-like ablators by reducing brittleness and increasing strength
without increasing density. Understand key parameters that control the thermal and
mechanical properties of low density, porous ablators using PICA as a model system.
Discover and develop new advanced ablators
G. Pulci, J. Tirillo “Carbon–phenolic ablative materials for re-entry space vehicles:
Manufacturing and properties” [5]. G. Pulci and J. Tirillo developed a carbon–phenolic
ablative TPS and tested with the aim of fulfilling the thermal and mechanical requirements
corresponding to the actual loads experienced by a vehicle during a moon earth re-entry.
From G. Pulci experiment there is a clear indication that a carbon/phenolic composite
material made of a rigid non-woven graphitic felt impregnated with a resole phenolic resin
can act as a very effective insulator in a very wide range of temperatures (at least from
room temperature to about 2000 oC).
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Alvaro Rodriguez, Cooper Snapp “Orbiter Thermal Protection System”
[6].
Alvaro and
Cooper discussed about the space shuttle orbiter thermal protection system and materials
used in TPS. Alvaro also discussed about the environment during re-entry of the orbiter
and the different types of testing carried for materials used in re-entry vehicle. Alvaro &
Cooper concluded that Advanced materials and coatings were key in enabling the success
of the shuttle in high-temperature environments.
Sylvia M. Johnson, (2001) "Thermal Protection Materials for Reentry and Planetary
Applications"
[7].
Thermal protection materials and coatings used in reusable launch
vehicles will be reviewed, including the needs and directions for new materials to enable
new missions that require faster turnaround and much greater reusability. The role of
ablative materials for use in high heat flux environments, especially for non-reusable
applications and upcoming planetary missions, will be discussed.
Ali Asgar, Sarath Raj N.S, (2019) “Ablative Heating Technology in Hypersonic Re-entry
Vehicles and Cruise Aircrafts” [8]. Ali Asgar analyzed some re-usable insulation tile thermal
protection system types such as high temperature reusable surface insulation tiles
(H.R.S.I), fibrous refectory composite insulation tiles (F.R.C.I), low temperature reusable
surface insulation tiles (L.R.S.I) and gradually move on to ablative thermal protection
systems with the advent of reinforced carbon-carbon’s application in astronautics and
aeronautics respectively.
H K M Al-Jothery, M A Abdullah (2019) “A review of ultra-high temperature materials for
thermal protection system” [9]. H K M Al- Jothery and M A Abdullah review covers briefly
all main types of Thermal Protection Systems (TPSs) and all the materials are used to
fabricate them with the maximum operational temperatures. Also, it covers the
promised UHTMs (SiC, ZrB2, HfB2, SiB6 and B4C) which are currently using for several
aerospace applications, especially for TPS. Besides, it discusses the oxidation of SiC,
B4C, SiB6, ZrB2 and HfB2.
Osama Gaballa, (2012) “Processing development of 4TaC-HfC and related carbides and
borides for extreme environments” [10]. Osama Gaballa discussed about different types
of Carbides, nitrides, and borides ceramics and also wear-resistant coatings. Osama
Gaballa mainly focused on Tantalum and Hafnium carbides.
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3. Details of the Topic
3.1.1 Re-entry Vehicles
A re-entry vehicle is the part of a spacecraft that is designed to return through Earth's
atmosphere. It is built to survive intense heating during high-velocity flight through the
atmosphere and to protect the crew and/or instruments until it brings them safely to
Earth. Although the technology has changed over time, re-entry vehicles since the early
Mercury program have used the same basic design concept: a blunt shape protected by a
heat shield. During re-entry, the shuttle is going so fast, it compresses the air ahead of it.
The compression of the air layers near the leading edges of the shuttle is quick, causing
the temperature of the air to rise to as high as 3000 degrees Fahrenheit! Being in contact
with the shuttle, it heats the shuttle’s surface. Normally, this high temperature will melt
almost any material- from the rock of a meteor to the metal skin of a space shuttle.
Figure 2 Soyuz Descent Module reaches Entry Interface, where friction from
Earth's thickening atmosphere heats its outer surfaces. Credits: NASA
The shuttle enters or "attacks" the atmosphere at such an angle that its nose and
underside contact and compresses the air and absorbs most of the heat generated. On
the underside are incredibly heat resistant, insulating silica tiles. They conduct heat very
poorly and thus keep heat from penetrating to the metal skin of the shuttle. Without the
tiles, the 3000 degree air touching the shuttle’s metal skin would melt through it. The
shuttle can lose an occasional tile and not incur much heat damage. However, if many
tiles are missing the heat can do severe damage. Friction also creates heat as the air
molecules rub across the shuttle’s surface. This further contributes to the heat the shuttle
must endure.
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Re-entering Earth is all about attitude control. Rather, it refers to the angle at which the
spacecraft flies. Here's an overview of a shuttle descent:
•
Leaving orbit: To slow the ship down from its extreme orbit speed, the ship flipped
around and actually flew backward for a period. The orbital maneuvering engines
(OMS) then thrust the ship out of orbit and toward Earth.
•
Descent through atmosphere: After it was safely out of orbit, the shuttle turned
nose-first again and entered the atmosphere belly-down (like a belly-flop) to take
advantage of drag with its blunt bottom. Computers pulled the nose up to an angle
of attack (angle of descent) of about 40 degrees.
Figure 3 China’s Shenzhou manned spacecraft and its reentry Credits: Enabling
Technologies for Chinese Mars Lander Guidance System, Xiuqiang Jiang
•
Landing: If you've seen the movie "Apollo 13," you might remember that the
astronauts return to Earth in their command module and land in the ocean where
rescue workers pick them up. Space shuttles looked and landed much more like
airplanes. Once the ship got low enough, the commander took over the computers
and guided the shuttle to a landing strip. As it rolled along the strip, it deployed a
parachute to slow it down.
The trip back to Earth is a hot one. Instead of the ablative materials found on the Apollo
spacecraft, space shuttles had special heat-resistant materials and insulating tiles that
could sustain re-entry heat.
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3.1.2 Thermal Protection System
The term thermal protection system (TPS) refers to various materials applied externally
to the outer structural skin on an orbiter to maintain acceptable temperatures, especially
for the reentry phase of a mission. Materials used for a TPS are selected for their hightemperature stability and weight efficiency.
Space vehicles that enter the earth’s atmosphere require thermal protection systems to
protect them from aerodynamic heating. The TPS system used by space vehicles inhibits
the conduction of heat on the interior of the vehicle by combining an underlying layer of
the Friction with the atmosphere during re-entry produces extreme temperatures that
require specialized shielding systems to protect space vehicles.
In addition to heat, space vehicle thermal protection systems also shield systems and the
airframe from the extremely cold conditions that occur during parts of orbit thermal
insulation with high-temperature resistant surface materials.
Thermal protection system (TPS) acts as a protective covering for the re-entry vehicle with
a temperature of 1650 ℃ (3000 ℉) heat at the time of atmospheric re-entry.
The following are the TPS materials used in re-entry vehicles
•
Reinforced Carbon Carbon
•
High Temperature Reusable Surface Insulation Tile
•
Low Temperature Reusable Surface Insulation Tile
•
Advanced Flexible Reusable Surface Insulation Blanket
•
Flexible Reusable Surface Insulation Blanket
•
Phenolic Impregnated Carbon Ablator
3.1.3 Requirement of Thermal Protection System
The amount of pressure and aerodynamic heating that occurs during launch and re-entry
varies according to vehicle type, shape, and trajectory.
•
To provide adequate protection, the following requirements must be met for all
thermal protection systems
– Heat load: Regulating the flow of heat into and out of the vehicle is the
main role of thermal protection systems. The TPS system must be able to
withstand high temperatures without excessive degradation of material
properties.
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– Mechanic loads: The TPS must be able to withstand extreme aerodynamic
pressure, as well as in-plane inertial, dynamic, and acoustic loads without
failure.
– Chemical deterioration: High surface temperatures during re-entry make
the TPS susceptible to oxidation.
– Deflection limits: The TPS shapes the vehicle’s aerodynamic profile. Surface
deflections of the TPS need to be below a certain limit to maintain this
aerodynamic profile and prevent local overheating and system failure.
– Impact loads: The TPS can be subjected to many types of impact during
installation, launch, flight, and landing. Having adequate impact resistance
is an important requirement of a TPS.
– Lightweight: Due to the large amount of space that a TPS occupies, it makes
up a majority of the launch weight. To prevent the need for increased fuel
requirements, a TPS must be as lightweight as possible.
3.1.4 Required Material Properties for TPS
– It must be Light weight, Reusable and cost effective.
– The material used in thermal protection system should be environmentally
friendly.
– Low density
– The material must have high temperature handling capacity or the capability of
minimal thermal conductivity
– High Damage Tolerance
– High Melting point
– High Strength and corrosion resistance
– Fatigue and Fracture resistance
– High strength
– Most durable material in its design with a low vulnerability to orbital debris.
– High abrasion resistance
– Higher Melting point
– High creep and rupture strength
– Oxidation resistance
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3.2 Thermal Protection System Materials and their properties
3.2.1. Reinforced Carbon Carbon:
RCC is a composite material consisting of carbon fiber reinforcement in a matrix
of graphite. Reinforced Carbon-Carbon composite is an amorphous carbon matrix. Both
reinforcing fibers and matrix are pure carbon. It was developed for the re-entry vehicles
of intercontinental ballistic missiles, and is most widely known as the material for the nose
cone and wing leading edges of the Space Shuttle orbiter.
Figure 4 RCC used in leading edge of space shuttle orbiter
RCC can withstand temperature up to 1260℃ (2300℉) during atmospheric re-entry. So
mostly used in leading edge of the wing, nose cap etc.
Properties of Reinforced Carbon Carbon (RCC)
– Excellent thermal shock resistance
– Low density (1830 kg/m3)
– High modulus of elasticity (200GPa)
– Low Coefficient of Thermal Expansion
– High fatigue resistance
– High strength
– Low coefficient of friction (in the fiber direction)
– Excellent heat resistance in non-oxidizing atmosphere.
– High abrasion resistance
– High electrical conductivity
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3.2.2. High temperature reusable surface insulation HRSI tile
The HRSI tiles are made of low-density, high-purity silica; it is a 99.8-percent
amorphous fiber. Because 90 percent of the tile is air and the remaining 10 percent is
material, the tile weighs approximately 9 pounds per cubic foot. HRSI tiles (black in color)
provide protection against temperatures up to 1,260 °C (2,300 °F). There are 20,548 HRSI
tiles which cover the landing gear doors and the rest of the orbiters under surfaces.
HRSI is used in conjunction with stronger, waterproof materials in the space shuttle heat
shielding to give a balance of strength and resistance to the high re-entry temperatures
experienced in earth's upper atmosphere.
Figure 5 HRSI Tile used in Landing Gear doors and some parts of orbiter under
surface Credits: Space Shuttle Tiles NASA
3.2.3. Low temperature reusable surface insulation LRSI tile
A white coating was used on the LRSI tiles which had the optical properties
required to maintain on-orbit temperatures for vehicle thermal control purposes. LRSI is
white in color, these covered the upper wing near the leading edge. They were also used
in selected areas of the forward, mid, and aft fuselage, vertical tail. The white color was
by design and helped to manage heat on orbit when the orbiter was exposed to direct
sunlight. These tiles protect areas where reentry temperatures are below 1,200 °F (649
°C). These tiles were reusable for up to 100 missions with refurbishment.
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Figure 6 LRSI Tile used in some upper surface of space shuttle Credits: Orbiter
TPS, NASA
3.2.4. High-Efficiency Tantalum-based ceramics (HETC) composite
NASA has developed High-Efficiency Tantalum-based ceramics (HETC) composite
structures that are suitable for use in thermal protection systems. These composite
structures have high-efficiency surfaces (low catalytic efficiency and high-emittance),
thereby reducing heat flux to a spacecraft during planetary re-entry. These low catalytic
efficiency and high-emittance ceramic materials were developed in order to increase the
capability of a Toughened Uni-Piece Fibrous Insulation (TUFI)-like thermal protection
system, with its high-impact resistance, to temperatures above 3000 ℉ (1650℃). These
ceramics have been applied to various aerodynamic configurations, such as wedge, wingleading segment and conventional tile shapes used on high-speed atmospheric entry
vehicles. In addition, this family of tantalum-based ceramics exhibits low catalytic
efficiency to atom recombination during exposure to high-energy dissociated hypersonic
flow.
Benefits of High-Efficiency Tantalum-based ceramics (HETC)
– Survives high heat fluxes 1650℃ and above
– Light weight
– Low cost to fabricate and maintain
– Performs at high-efficiency during hypersonic earth atmospheric entry
– Easier to design
– Resistant to erosion and/or impact damage
– Provides a composite insulating structure
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3.2.5. Phenolic Impregnated Carbon Ablator, PICA
Phenolic Impregnated Carbon Ablator, PICA, a carbon fiber is applied in phenolic resin. It
is a modern thermal protection system. PICA is a carbon fiber is applied in phenolic resin.
It is a modern thermal protection system. Phenolic Impregnated Carbon Ablator (PICA) is
a member of the family of Lightweight Ceramic Ablators (LCAs) and was developed at
NASA Ames Research Center as a thermal system (TPS) material for the Stardust mission
probe that entered the Earth’s atmosphere faster than any other probe or vehicle to date
It has a main advantage of low density that is much lighter than the carbon phenolic. PICA
has low thermal conductivity, when compared to other high heat flux ablative material
like convectional carbon phenolic.
Figure 7 PICA formation process Credits: [3]
Properties of Phenolic Impregnated Carbon Ablator PICA
– Much lower density than other carbon/phenolic systems
– PICA has low thermal conductivity, when compared to other high heat flux ablative
material like convectional carbon phenolic.
– Design optimization using composite response surfaces
– Improved robustness
– Applied without having an effect on the final tolerance
– Optimization of surface densification as needed
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Figure 8 Mars Science Laboratory (MSL) 4.5m PICA Heat Shield
3.2.6. Tantalum Carbide and Hafnium Carbide:
These materials are refractory chemical compound/ceramics where they can able to
withstand high temperature. During re-entry high temperature is generated due to
friction. Tantalum carbide TaC and Hafnium carbide HfC can able to resist extreme heat
generated, so it is an efficient material to re-entry vehicle. The re-entry temperature
during re-entry is of approximate 1649℃. So, we need to withstand this high amount of
temperature. Tantalum carbide and Hafnium carbide can able to withstand approximately
4000℃. So, we can ensure this material for hypersonic space vehicle or re-entry vehicles.
Mainly these materials are used for nose cap in re-entry vehicle (since temperature effect
is high for nose part in re-entry vehicles). Their melting point is high enough than any
other material (approximate 4000℃ for HfC and TaC), make it as a good quality material
to used for re-entry vehicle manufacture.
➢ Properties of Tantalum Carbide
– High strength
– High hardness (between 11 to 26 GPa)
– Wear resistance, fracture toughness
– Resistance to chemical attack
– TaC as high melting point (3880˚C),
– High elastic modulus (up to 550 GPa)
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– Low electrical resistivity
– Good corrosion resistance, durability
– High-temperature strength
– Oxidation resistance
So, TaC suggest that this material could be used in tool steels, wear-resistant parts,
diffusion barriers, hypersonic vehicles (leading edges and nosecaps), propulsion
components (rocket nozzles), supersonic re-entry vehicles, scramjet components, hard
coatings, conducting films, electrical contacts, and electronic applications
➢ Properties of Hafnium Carbide
– HfC has excellent chemical stability
– High oxidation resistance
– High hardness (up to 33 GPa)
– High electrical and thermal conductivity
– High Young’s modulus (up to 434GPa).
HfC is used mainly in aerospace applications due to its high melting point and low diffusion
coefficients at high temperatures (thermal protection materials in both re-entry and
hypersonic vehicles).
It is used in coatings for ultrahigh-temperature environments due to its high hardness,
excellent wear resistance, good resistance to corrosion, and low thermal conductivity.
It is also used in cutting tools, high temperature shielding, field emitter tips, and arrays
HfC has the lowest work function of all transition metal carbides.
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4. Challenges and Opportunities
4.1 Challenges during Re-entry
•
Atmospheric Drag
Atmospheric drag is the atmospheric force (friction) acting opposite to the relative
motion of an object. Atmospheric drag is affected by the density of the atmosphere—
a denser atmosphere will cause more atmospheric drag than a lower density one. At
higher altitudes, the atmosphere is less dense. The atmosphere is created by the
earth’s magnetic field which emerges from the earth’s interior. This creates a gravity
field that shapes the atmosphere.
•
Aerodynamic Heating
Aerodynamic heating is the heating of a solid body produced by its high-speed passage
through air (or by the passage of air past a static body), whereby its kinetic energy is
converted to heat by adiabatic heating, and by skin friction on the surface of the object
at a rate that depends on the viscosity and speed of the air. In science and engineering,
it is most frequently a concern regarding meteors, atmospheric reentry of spacecraft,
and the design of high-speed aircraft.
Heating caused by the very high reentry speeds (greater than Mach 20) is sufficient to
destroy the vehicle unless special techniques are used. The early space capsules such
as used on Mercury, Gemini, and Apollo were given blunt shapes to produce a standoff bow shock, allowing most of the heat to dissipate into the surrounding air.
Additionally, these vehicles had ablative material that sublimates into a gas at high
temperature.
Aerodynamic heating mainly depends on
▪
Entry speed
▪
Shape of vehicle
▪
TPS material composition
▪
Properties of surface
▪
Atmospheric condition
▪
Trajectory path
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4.1.1 Challenges in Selection of Materials
The task of materials selection and applications in Re-entry and space systems is very
challenging. Since it must take into account many different and often conflicting factors
that affect their choice, Materials and processes considered for space applications
generally fall into two categories:
•
Existing materials and processes: These materials processes are already developed
and proven either in aerospace or in non-space applications but proposed for use in
new space applications
•
New and advanced materials and processes in various stages of development, for
aerospace and non-aerospace applications, but with limited or no experience.
Materials Challenges in Re-entry vehicle or space system
•
Materials performance: The goal here is to minimize the mass of the space systems
since the cost per pound of launching payloads is extremely high. At the same time
the designer must maintain adequate safety margins, to minimize the risk of
mission failure.
•
Environmental factors: Materials must withstand the harsh environment of
propulsion systems and the space environment itself. The space environment
typically is hard vacuum with galactic cosmic radiation that can be harmful to the
materials. Lack of stability of materials in these environments can be life limiting.
•
Manufacturability: Any material selected for space system must be capable of
being formed into shapes and joined to build components and subsystems. Special
equipment or facilities are often required. The assembled hardware should be
capable of being inspected nondestructively.
•
Affordability: The materials should be affordable to keep cost under control.
•
Availability: The materials should be readily available in order to procure them in
sufficient quantities to build flight hardware. Space business is typically a small
volume business and often unsteady and reliable suppliers of key materials may
be hard to find.
Technological challenges in materials processing, particularly in the area of hightemperature should be viewed as opportunities to innovate and create value.
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6. Conclusion and Suggestions for Future Work
6.1 Conclusions
In the analysis of materials used in re-entry vehicle it is found that the material plays a
vital role in re-entry vehicle. Because material is one which acts as covering layers of reentry vehicle. The materials must be temperature resistance, affordable, easy to
manufacture and implement to re-entry vehicle. From the study conducted it is
concluded that the material must have following properties to overcome difficulties
during re-entry.
•
Low thermal conductivity
•
High creep and rupture strength
•
Low density
•
Able to withstand Space environment
•
High strength
•
Corrosion resistance
•
High abrasion resistance
•
Outstanding fatigue strength
•
Higher Melting point
•
Oxidation resistance and thermal
•
Reusable
resistance etc.
Also, from the study carried it can be concluded that High efficiency Tantalum based
ceramics composites, Tantalum Carbide and Hafnium Carbide shows excellent properties
during high temperature conditions. So, it can be used as the coatings for re-entry vehicle.
6.2 Suggestions for Future Works
The new vision for space exploration encompasses a broad range of human and robotic
missions to the Moon. Mars and beyond. Extended human space travel requires high
reliability and high-performance systems for propulsion, vehicle structures, thermal and
radiation protection for re-entry vehicle, crew habitats and vehicle health monitoring, all
of vehicle are dependent on advanced materials.
Materials challenges for building such systems include the development of high strength
lightweight metallic alloys, ceramics and composite materials, which can withstand the
harsh environment atmospheric re-entry. The development of new high temperature
materials is necessary. There is lot of research is being done on high temperature
materials.
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REFERENCES
1. Ahana Arshad S, Dinu Vijay, Sajith R, Jincy J C (2019) Analysis of materials used for reentry vehicle, (ISSN-2349-5162)
2. Sylvia M. Johnson (2015) Thermal Protection Materials and Systems: Past and Future
40th International Conference and Exposition on Advanced Ceramics and Composites
3. Ryan Oakes “Space Shuttle Ceramic Tiles”.
https://depts.washington.edu/matseed/mse_resources/Webpage/Space%20Shuttle
%20Tiles/Space%20Shuttle%20Tiles.htm
4. J. Thornton, W. Fan “PICA Variants with Improved Mechanical Properties”.
5. G. Pulci, J. Tirillo “Carbon–phenolic ablative materials for re-entry space vehicles:
Manufacturing and properties”. (doi:10.1016/j.compositesa.2010.06.010)
6. Alvaro Rodriguez, Cooper Snapp “Orbiter Thermal Protection System”.
7. Sylvia M. Johnson, (2001) "Thermal Protection Materials for Reentry and Planetary
Applications".
8. Ali Asgar, Sarath Raj N.S, (2019) “Ablative Heating Technology in Hypersonic Re-entry
Vehicles and Cruise Aircrafts”.
9. H K M Al-Jothery, M A Abdullah (2019) “A review of ultra-high temperature materials
for thermal protection system”.
10. Osama Gaballa, (2012) “Processing development of 4TaC-HfC and related carbides and
borides for extreme environments”.
11. Dr. Biliyar N. Bhat, “Materials Challenges in Space Exploration”.
12. Material and coatings, High Efficiency Tantalum based Ceramic Composite (HETC)
Structures, (TOP2-153), NASA Ames Research center
13. Materials Used in Space Shuttle Thermal Protection Systems – Alessandro Pirolini (10th
Oct 2014) https://www.azom.com/article.aspx?ArticleID=11443
14. https://technology.nasa.gov/patent/TOP2-183
15. https://en.wikipedia.org/wiki/Atmospheric_entry
16. https://technology.nasa.gov/patent/TOP2-241
17. Thermal
protection
system
design-engineering-reentry-vehicles,
https://peachyessay.com/sample-essay/thermal-protection-system-designengineering-reentry-vehicles/
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