solar thermal propulsion
TRANSCRIPT
MOHAMMED AHSAN SHARIEF
SOLAR THERMAL PROPULSION
ABSTRACT
Solar thermal propulsion is a form of spacecraft propulsion. Spacecraft propulsion is used to change
the velocity of spacecraft and artificial satellites. There are many methods for spacecraft propulsion. Solar
thermal propulsion is an excellent choice because it requires only one propellant and combines moderate
thrust with moderate propellant efficiency. A solar thermal rocket has to carry only the means of capturing
solar energy such as concentrators and mirrors. Instead of converging that solar energy to electrical power as
in the case of photovoltaic systems where a solar thermal propulsion system uses the solar energy directly as
heat. The heated propellant is fed through a conventional rocket nozzle to produce thrust. Typically
hydrogen is used as the propellant due to its low molecular weight corresponding to a high specific impulse
Solar thermal propulsion effectively bridges the performance between the chemical and electrical
propulsion. Solar thermal propulsion system provides long duration and long distances are suitable for inter
orbit transfer and maneuvering missions. In this system the engine thrust is directly related to the surface
area of the solar collector and to the local intensity of the solar radiation. Now solar thermal propulsion is an
active area of research. This technology development has continued to be advanced under air force research
laboratory [AFRL] over the last 20 years. this paper focuses on a low earth orbit LEO to geosynchronous
equatorial orbit GEO transportation system using a solar thermal system.
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1. INTRODUCTION
A solar thermal rocket has to carry only the means of capturing solar energy such as concentrators
and mirrors. Instead of converging solar energy to electric power as like a photovoltaic system, a solar
thermal propulsion system uses the solar energy directly as heat. The heated propellant is fed through a
conventional rocket nozzle to produce thrust. The engine thrust is directly related to the surface area of the
solar collector and to the local intensity of the solar radiation.
2. BASIC PRINCIPLE
The propulsion system of a solar thermal powered space craft consist of three basic elements.
1. Concentrator
2. Thruster/Absorber
3. Propellant system
Concentrator focuses and directs incident solar radiation to an absorber/thruster which receives solar
energy, heats and expands propellant (hydrogen) to produce thrust. A propellant system which stores
cryogenic propellant extended periods and passively feeds it to the thruster/absorber. Figure1 shows the
basic principle of the solar thermal propulsion system.
The basic principle of solar thermal propulsion is to utilize the solar light to heat up a propellant and
providing thrust by expanding the resulting hot gas through a conventional rocket nozzle. Therefore, the
light is collected by parabolic reflectors and focused into a black-body cavity. Inside the cavity the high
temperatures in the focal area are radiated to its walls where the heat is absorbed and transferred to the
propellant flowing around the cavity. The propellant heats up to temperatures above 2000 K and is expanded
through the nozzle, thereby generating the thrust. The best propulsive performance can be achieved with
hydrogen (lowest molar mass) preferably stored in the liquid phase.
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Fig: 1 Solar thermal thruster
3. COMPONENTS OF AN STP SPACE CRAFT
3.1 SOLAR CONCENTRATORS:
Solar concentrators for use in space have received growing attention in the past few years in view of
their many potential applications. Among those, perhaps the most important ones are space power
generation and solar thermal propulsion. In the former, the concentrator is used to focus solar radiation on a
conversion device, e.g, a photovoltaic array or the high temperature and of a dynamic engine; in the latter,
concentrated solar radiation is used to heat a low molecular weight gas, thereby providing thrust to a solar
rocket.
In this propulsion scheme, solar energy is reflected by the large parabolic reflectors towards the
rocket body, where hydrogen fuel is heated to a very high temperature and exhausted through a nozzle.
Another application of space borne solar concentrators is for power generation. Future mission in space will
require abundant power for use on satellites. While conventional photovoltaic have been used in the past and
provide a reliable source of power, they do have several drawbacks. Their low efficiencies make it necessary
to use large areas of cells, requiring extendible hard structures for support. These large structures make for a
complex deployment scheme as well as a high system weight. Another drawback is that the large area
required for the low efficiency cells will create significant drag for satellites, especially in low earth orbit.
Solar dynamic power systems [SDPS] offer a viable alternative to photovoltaic, with lower system weight
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MOHAMMED AHSAN SHARIEFand drag area. These power systems typically consist of large parabolic reflectors that focus solar radiation
into a receiver where high intensity heat is collected. This heat is then used to generate mechanical power
using a Brayton, Rankine, or Stirling cycle engine. The lower system weight and area is mainly due to the
higher efficiency of dynamic power systems; for a given area of collector surface more energy is generated
with the dynamic power system than with photovoltaic.
A solar concentrator uses lenses called Fresnel lenses, which take a large area of sunlight and directs
it towards a specific spot by bending the rays of light and focusing them. Fresnel lenses uses like a dart
board, with concentric rings of prisms around a lens that’s a magnifying glass. All these features let them
focus scattered light from the sun in to a tight beam. Solar concentrators put one of these lenses on top of
every solar cell. This makes much focused light come to e ach solar cell, making the cells vastly more
efficient.
Two concentrator designs, rigid or inflatable were originally being evaluated under two different
contracts. However, these two different programs have since been merged, with the inflatable concentrator
design taking lead as the primary technology. An inflatable solar concentrator offers significant advantages
in comparison to state-of-the-art rigid panel concentrators, including low weight, low stowage volume, and
simple gas deployment.
3.2 TORUS AND SUPPORT STRUCTURE:
The reflector is mounted on the torus and support structure such that the mirror focuses solar
radiation into the receiver to the solar energy absorber. An inflatable torus and support structure can be
fabricated with kevlar-weave teflon laminate materials. Solar radiation exposure heats the inflatable torus,
causing pyrolitic deposition of nickel metal on the inside of the inflatable, rigidizing it to produce load-
heaving capacity, high-rigidity and high pointing accuracy.
3.3 GIMBALING RECEIVER ASSEMBLY:
The gimbaling receiver-assembly is made of the receiver housing, the reflector mounting ring
rotation systems, and the rotation system that mates from the receiver housing to the spacecraft. The receiver
mechanically points the reflectors to maintain solar energy focus on the solar energy absorber.
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MOHAMMED AHSAN SHARIEF3.4 SOLAR ENERGY ABSORBER
The solar energy absorber produces superheated hydrogen with the heat from the absorption of
focused solar energy. Small capillary metal-matrix heat transfer elements may be useful in the construction
of solar energy absorbers. In the operation of a solar thermal engine, the absorber configuration as a heat
exchanger. Transport of high intensity solar flux from the concentrator to the solar receiver via optical fiber
cable the solar receiver core is made of graphite cylinder because of high solar absorbtivity [.7-.9] ,excellent
thermal mechanical stability and ease of fabrication The gas was injected tangentially in to the graphite
cylinder and flows out through the molybdenum tube. The graphite core is surrounded by the molybdenum
radiation shields. Achievement of high temperature via radiative heat transfer.
3.5 POINTING AND NAVIGATION SYSTEM
In order for the reflectors to remain focused on the solar energy absorber at all times, the navigation
and sun sensing and pointing systems must be integrated in real-time. Upon change in attitude to the sun the
receiver mechanism will make suitable adjustments to maintain solar radiation pointing accuracy
4. SOLAR THERMAL PROPULSION CONCEPTS
Two system level approaches for STP are currently being explored. Direct gain approach and
thermal storage concept. That determines the amount of rotation required from the concentrator pointing
mechanism.
DIRECT GAIN CONCEPT
In the direct gain concept the concentrator continuously tracks the sun during the burn while the
space craft remain pointed along the desired orbital trajectory. This requires that the concentrator be able to
rotate up to 180 degrees while the space craft rolls 180 degrees. The direct gain concept will eventually
require that the concentrator be mounted on a turn-table capable of the large deflections. The absorber
configuration is a windowless heat exchanger having a delivered specific impulse of 800-960 seconds.
Volumetric absorber concepts can potentially provide performance levels approaches 1100 seconds.
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MOHAMMED AHSAN SHARIEFTHERMAL STORAGE CONCEPT
The second design approach involves the incorporation of a thermal storage medium in which solar
energy is required and stored during the coast period of the orbit and when a propulsive burn is required,
propellant flows through the thermal storage medium to provide thrust. The storage of solar energy enables
a higher thrust than the direct gain concept with smaller concentrators. For efficient operation, the burns of
this engine concept should be performed in the eclipse portion of the orbit. This greatly simplifies the sun
tracking and thrust orientation compared with the direct gain concept since the system does not have to be
"on sun" during the burn. In the current design concept, which uses rhenium coated graphite as the thermal
storage medium, a delivered specific impulse of 700 to 900 sec is predicted dependent on the thermal
storage temperature. Once the vehicle is in orbit, the concept can also provide on orbit power using the
concentrators and thermionic elements to generate electricity. To achieve the desired long life for the power
system, the concept typically incorporates a rigid concentrator.
5. METHODS FOR HEATING PROPELLANT
There are two methods for heating the propellant. They are direct method and indirect method.
DIRECT METHOD
In the direct method, the propellant flows through sandy material within the heat exchange cavity.
We put holes in the pipes or walls of the indirect heat exchanger so that the gas flows directly into the heat
cavity, which requires a window, as pictured below: Direct solar radiation absorption (steam goes into
windowed heating chamber In the direct concept, the cylindrical heating chamber rotates so that the
centrifugal force keeps the sand, or "seeds", along the chamber wall, which is porous to let the gas in. The
seeds are chosen for stability at high temperature and heat transfer properties. (Tantalum carbide and
hafnium carbide are popular.)Heat transfer is more efficient in the direct concept, i.e., it's more compact, but
clouding of the window or eventual leakage around and other seals are serious concerns. The rotating
chamber is considerably more complex
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Fig: 4. Direct propellant heating
IN DIRECT METHOD
Indirect solar radiation has the propellant flow through only pipes or passages in the wall of a
windowless heating cavity as shown below. Then this gas passes through a nozzle.
Fig: 5. Indirect propellant heating.
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6. WORKING OF SOLAR THERMAL SPACE CRAFT
The concentrator and the absorber/thruster are optically coupled with the absorber located at the
concentrator focus. Due to large size inflated concentrators and non rigid support structure, the optically
coupled concentrator absorber configuration can be sensitive to structural deformations caused by
concentrator sub system rotation or acceleration. The optical wave guide transmission line is the key
component to integrate the concentrator system with the solar thermal receiver. The cable inlet interfaces
with the concentrator system and the outlet interfaces with the solar thermal absorber. The propellant was
injected tangentially in to the graphite core, which contain channels for heating the propellant Hydrogen is
expanded and produces thrust.
Fig:6. Deployed view
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7. SOLAR THERMAL PROPULSION FOR A SMALL SPACE CRAFT
Fig:7. The off axis inflated concentrator STP system
The Boeing Company is developing an innovative solar thermal propulsion system for application to
small solar thermal propulsion system for application to small space craft with funding support by the Air
Force Research Laboratory. In this system, as schematically presented in Fig.7, solar radiation is collected
by the concentrator which transfers the concentrated solar radiation to the optical waveguide transmission
line consisting of low-loss optical fibers. The optical waveguide cable transmits the high intensity solar
radiation to the thermal receiver for efficient, high performance thrust generation. Part of the solar radiation
can be switched to attitude control thruster as necessary. The features of the proposed system are:
l. Highly concentrated solar radiation (I03 suns) can be transmitted via flexible optical waveguide
transmission line to the thruster’s absorber cavity;
2. The flexible optical waveguide linkage de-couples the thruster from the concentrator to provide
freedom from the constraints imposed on previous solar propulsion system designs;
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MOHAMMED AHSAN SHARIEF3. The configuration of the solar receiver can be optimized for efficient heat transfer with minimal re-
radiation loss;
4. Aiming and tracking for the concentrator become significantly easier by moving the termination of
the optical fiber cable to follow the focal point of the primary concentrator
5. High intensity solar radiation can be switched to different receivers to deploy several them1a1
thrusters as necessary.
Fig: 8. Solar thermal propulsion system for small space craft
The experimental facility consists of two solar tracking units each with two 50 cm parabolic
concentrators. The two concentrators are mounted on a rotating frame to track the sun. The optical fiber
cable placed at the focal point of the concentrator transmits the concentrated solar radiation to the solar
receiver located at the center of facility. The optical fiber cable (4 m long) consists of’37 fused silica fibers
(1.2-mm dia). The four optical fiber cables deliver about 200 W of solar power into the receiver.
The solar receiver is located at the center with four optical fiber cables connecting it to
four concentrators. The configuration of this experimental setup simulates the solar thermal propulsion
system described in Fig.8.
The hardware components that we developed in this program include: optical waveguide
transmission line; interface optical components; and the solar thermal receiver.
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OPTICAL WAVEGUIDE TRANSMISSION LINE
The optical waveguide transmission line is the key component to integrate the concentrator system
with the solar thermal receiver. The cable inlet interfaces with the concentrator system and the cable outlet
interfaces with the solar thermal receiver. The cable inlet design we used in this program is based on our
heritage: the quartz secondary concentrator collecting the solar radiation and injecting it to the optical fibers.
Figure 9 shows the inlet portion of the four optical fiber cables used for this program. All four cables are 4 m
long and each consists of 37 high numerical apertures. The fiber has an excellent off-axis transmission up to
25 degrees. The design of the cable outlet was developed for optimum interface with the high temperature
solar receiver. A photo of the fiber cable outlet is given in Fig. 10. The 37 optical fibers transfer the solar
radiation to the 10 mm quartz rod. The quartz rod, by the principle of total internal reflection, transfers the
solar radiation to the thermal receiver. The tip of the quartz rod is placed close h the receiver high
temperature heat exchanger in order to deliver the solar power directly to the receiver.
Fig. 9: inlet of optical fiber cable
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Fig 10: optical fiber cable out let made of quartz rod.
Solar receiver
One of the important objectives of this program was to demonstrate the basic solar receiver heat
transfer mechanisms:
Transport of high intensity solar flux from the concentrator to the solar receiver via optical fiber
cable;
Efficient delivery of high intensity solar flux to the solar receiver heating element;
Achievement of high temperature via radiative heat transfer; and .
Viability of optical components.
A schematic of the solar thermal receiver is given in Fig. 11.
The solar receiver core is made of graphite cylinder (diameter = 1.75 cm; height = 2.54 cm), because
of (i) high solar absorptivity (a= 0.7-0.9), (ii) excellent thermal-mechanical stability, and (iii) ease of
fabrication. The gas was injected tangentia1ly into the graphite cylinder and flows out through the
molybdenum tube. The graphite core is surrounded by the molybdenum radiation shields. Solar power (200
W) was delivered to the graphite core by four quartz rods (dia. = I cm).
The solar receiver housing with four optical fiber cables is shown in Fig.11. The construction of this
housing was similar to the materials processing experiment conducted in the previous NASA Program. The
propellant gas flows from the bottom of the housing, flows through the heat exchanger, and flows out of the
housing.
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FIG:11; SOLAR RECEIVER
8. SPECIFICATIONS OF STP SPACE CRAFT
The SOTV Space Experiment will be a turn-of-the-century demonstration of the operation and
performance of an advanced solar thermal propulsion and power engine. The SOTV engine offers the
potential for a revolutionary increase in specific impulse at moderate thrust levels that allow operation of
LEO-to-GEO transfers in 30 days or less. The technologies being developed for the SOTV in this AFRL
program have a wide range of applications including improved payload performance on expendable boosters
and reusable launch vehicles, power systems for high-power satellites, satellite servicing and repositioning,
and planetary injection for NASA probes. Ultimately, this technology can enable a fully reusable orbit
transfer vehicle capable of making routine a wide range of space operations at substantially lower cost than
current systems. SOTV is the direct successor of another AFRL program, the Integrated Solar Upper Stage
(ISUS) Engine Ground Demonstration (EGD) which was carried out in a large vacuum facility at NASA-
Lewis Research Center in the summer of 1997. EGD validated system level feasibility for the SOTV solar
thermal propulsion mode. The Space Experiment is the next logical step towards fielding an operational
SOTV.
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.
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Space Experiment Operational Vehicle
Propulsion Mode LH2 LH2 and/or storable
Max. Temperature up to 2300K up to 2400K
Chamber Pressure: 20-25 psia 20-50 psia
Nozzle Area Ratio: 100:1 fixed 100:1 - 200:1 fixed
Thrust: 1.6 lbf 10-50 lbf
Specific Impulse: 750 sec 800 sec+
Power Mode Space Experiment Operational Vehicle
Electric Power: 50 We 500 We to 50 KWe
Voltage: ~ 1 Vdc 28 to 70 Vdc
Mission Life : ~ 1 year 5 to 15 years
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9. BENEFITS AND LIMITATIONS
BENEFITS OF SOLAR THERMAL PROPULSION
High efficiency at potentially low cost
Higher payload fraction than chemical
Solar derived electric power
Concentrator & high-gain antenna or aero assist system
Higher Isp (> 700 s) than chemical options (300 -500 s)
Higher thrust-to-weight ratios than electric systems
Space solar power
Synthetic Aperture radar
Sunshield for space telescopes
High temperature materials
LIMITATIONS OF SOLAR THERMAL PROPULSION
It would not be very useful where places of intensity of sunlight is low
This propulsion system generates relatively low thrust necessitating 20-30 days to travel from LEO
TO GEO
Difficulty of ground level testing
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10. CONCLUSION
In the distant future, low cost propulsion will be needed for interplanetary travel and unmanned
exploration. NASA forces solar thermal propulsion as a way to boost future payloads from a low earth orbit
to a geosynchronous earth or high orbit. For more distant travel, a solar thermal engine using this propulsion
would acts like a simple, efficient tugboat in space. Solar thermal propulsion systems would be less
expensive, much simpler and more efficient than today’s rocket engines. A large liquid hydrogen tank with a
innovative feed system was tested at Marshall to simulate a 30 day solar thermal mission. Data gathered
from the tests would have applications for missions to the moon and mars, as well as boosting payloads to
higher orbits. Solar absorber, thruster, and inflated concentrator technology development have continued to
be advanced under Air force research laboratory [AFRL] over the last 2 years. Small scale hardware has
been designed and fabricated AFRL for ground level evaluation. Therefore solar thermal propulsion can be
literally defined as the future of space explorations
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11. REFERENCE
en.wikipedia.org/wiki/solar_thermal_rocket
www.osti.gov/energycitations/product.biblio.jsp?osti_id=7112464
www.vectorsite.net/trarokt2.html
www.inspacepropulsion.com/tech/solar_therm.html
www.highway2space.com
www.grc.nasa.gov/www/RT2001/5000/5490wong2./html
Jet and Rocket Propulsion , ML Madhur and RP Sharma
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CONTENTS
1. INTRODUCTION
2. BASIC PRINCIPLE
3. COMPONENTS OF AN STP SPACE CRAFT
4. SOLAR THERMAL PROPULSION CONCEPTS
5. METHODS FOR HEATING PROPELLANT
6. WORKING OF SOLAR THERMAL SPACE CRAFT
7. SOLAR THERMAL PROPULSION FOR A SMALL SPACE CRAFT
8. SPECIFICATIONS OF STP SPACE CRAFT
9. BENEFITS AND LIMITATIONS
10. CONCLUSION
11. REFERENCE
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