solar thermal propulsion full report
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SOLAR THERMAL PROPULSION
ABSTRACT
Solar thermal propulsion is a form of space craft propulsion. Space craft propulsion is used to change the velocity of space craft and artificial satellites. There are many methods for space craft propulsion. Each method has draw backs and advantages, and space craft propulsion is an active area of research. Solar thermal propulsion conceived in 1956 by Kraft Echrike. Solar thermal propulsion is an excellent choice because it requires only one propellant gas and combines moderate thrust with moderate propellant efficiency. Solar thermal propulsion effectively bridges the performance gap between chemicals and electric propulsion by potentially offering higher specific impulse (800 to 1000 seconds) than chemical propulsion (300 to 500 seconds). Typically hydrogen is used as the propellant due to its low molecular weight corresponding to a high specific impulse.
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.
1. INTRODUCTION
Solar thermal propulsion is a form of space craft propulsion. Space craft propulsion is used to change the velocity of space craft and artificial satellites. There are many methods for space craft propulsion. Each method has draw backs and advantages, and space craft propulsion is an active area of research. Solar thermal propulsion conceived in 1956 by Kraft Echrike. Solar thermal propulsion is an excellent choice because it requires only one propellant gas and combines moderate thrust with moderate propellant efficiency. Solar thermal propulsion effectively bridges the performance gap between chemicals and electric propulsion by potentially offering higher specific impulse (800 to 1000 seconds) than chemical propulsion (300 to 500 seconds). Typically hydrogen is used as the propellant due to its low molecular weight corresponding to a high specific impulse.
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.
Fig: 1 Solar thermal thruster
3. COMPONENTS OF AN STP SPACE CRAFT
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 and 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 thatâ„¢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.
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. Upon deployment, the torus and support structure would have nickel carbonyl introduce. Solar radiation exposure heats the inflatable, 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. GIMBALING RECEIVER ASSEMBLY:
The gimbaling receiver-assembly is made up 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.
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.
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. INFATABLE CONCENTRATORS
Each solar thermal propulsion vehicle will have two pre-molded, inflatable solar concentrators made almost entirely of a new polyimide material developed by the NASA Langley Research Center, Hampton, Virginia. The LaRC-CP1TM polyimide is a clear, lightweight material with a large thermal operating range. It is ideal for this aerospace application because it effectively forms compound curved shapes; it is resistant to UV radiation, stable in a space vacuum and lightweight compared to glass or metal optics.
The solar concentrators must be designed so that their 9 x 13-foot reflectors achieve a precise surface slope when inflated in orbit. A precise shape is needed to focus an optimal amount of the Sunâ„¢s energy on a heat exchanger engine that heats hydrogen gas. The expanding gas provides enough thrust to transfer the satellite to the higher orbit. Figure shows an inflated solar concentrator
Fig: 2 inflatable solar concentrator
Inflatable space systems invariably require less packaged volume, are lower in weight and cheaper through both development and production phases than competing mechanically erected systems. The potentially harmful effects of the space environment, including that of micrometeoroids, are much less than originally anticipated since large inflatable concentrators require very low inflation pressure; gas lost through leaks can be easily replaced from a small supply of reserve gas. Inflatable deploy and function
very well in space, where the absence of gravity creates extremely low loads. High surface accuracy is obtained due to the constant force provided by the inflatant. The ultimate system will require two reflectors, each having an elliptical rim with a 40m major axis, to provide 40 lbs of thrust to the two engines of the rocket. Under the present project and implimentation, a one-fourth scale, 9*7m off axis concentrator has been under development as a pilot for the full scale flight unit. The reflector component consists of a reflective membrane made of specially designed gores and a geometrically identical transparent canopy. The two forms together an inflatable lens like structure which, upon inflation, assumes the accurate paraboloidal shape. This inflatable structure is supported along its rim by a strong, bending-resistant torus. Figure illustrates the concentrator showing its main parts.
.
Fig: 3. Parts of concentrator
It consists of two geometrically identical thin membranes forming together a pillow like structure which, upon inflation, assumes the accurate paraboloidal shape. The two membranes are attached together along the plane, elliptical rim. One membrane has a reflective coating and serves as a concentrating mirror; the other is transparent and serves as a canopy. As mentioned earlier, the dimensions of the prototype reflector presently under development are 7 x 9 m. The inflatable pillow is mounted on a strong, bending-resistant rim support by means of connecting tandems and straps. The reflector
membrane and the canopy are made of very thin films of plastic material. For the current development, 0.25 mil Mylar has been used. The materials for the flight unit will be selected to withstand the damaging effects of the space environment. Each membrane consists of specially designed flat gores seamed together along the edges. When pressurized, the gores undergo an elastic deformation so as to assume the design paraboloidal shape. Note that the uninflected shape is not a paraboloidal. A design algorithm has been developed to calculate the shape of the gores as a function of the properties of the material and the inflation pressure. The details of the design process are described in the next section. It should be noted that the inflation pressure required is very low, on the order of 0.0002 psi, and is inversely proportional to the size of the reflector. The pressure acting over the membrane surface creates a sizable force which must be taken up by the rim support at the edges of the membrane. The rim support contemplated for the present concentrator is an elliptical torus with a circular cross section. Two alternatives have been considered for the torus aigu: one IS a IUII- mtlatable toroidal tube; the other is a rigidized structure with either a solid or an annular cross-section. The former is preferable from the standpoint of easy deployment and low weight; there is, however, a risk of puncture by micrometeoroids and space debris which could cause significant loss of inflatant. Note that the torus is different in this respect from the reflector, the former requiring much greater pressure than the latter. The members of the truss connecting the concentrator assembly to the rocket are made of the thin, lightweight aluminium shell. The truss is packaged prior to deployment along with the rest of the concentrator. When inflated, the aluminum shell undergoes plastic deformation which rigidizes it in place. Loss of inflatant from the shell later on will not affect the dimensional stability of the truss. Changes in length of the truss members due to thermal expansion and contractions may be compensated for by the jackscrew. The concentrator is equipped with a guidance and control system design to maintain its accurate paraboloidal shape and point it towards the sun, when required, at the correct position with respect to the rocket engine. The former task is performed by sensing the distance between the membrane and adding or venting inflatant, as needed; the later task is performed by the jackscrews and turntable. Note that the concentrators are in operation only through a fraction of the mission time. The solar rocket path from LEO to geo involves a series of orbits around the earth, with relatively short apogee and perigee firings of the engine.
5. 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.
THERMAL 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.
6. 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 flow 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
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. In direct propellant heating.
7. 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 produce thrust
Fig:6. Deployed view
8. 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 thrusterâ„¢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;
3. 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 ofâ„¢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.
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
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.
FIG:11; SOLAR RECEIVER
9. 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.
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
10. 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
11. 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 todayâ„¢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
12. REFERENCE
en.wikipediawiki/solar_thermal_rocket
osti.gov/energycitations/product.biblio.jsposti_id=7112464
vectorsitetrarokt2.html
inspacepropulsiontech/solar_therm.html
highway2space.com
grc.nasa.gov/www/RT2001/5000/5490wong2./html
Jet and Rocket Propulsion , ML Madhur and RP Sharma
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.
CONTENTS
1. INTRODUCTION 1
2. BASIC PRINCIPLE 2
3. COMPONENTS OF AN STP SPACE CRAFT 3
4. INFATABLE CONCENTRATORS 6
5. SOLAR THERMAL PROPULSION CONCEPTS 9
6. METHODS FOR HEATING PROPELLANT 10
7. WORKING OF SOLAR THERMAL SPACE CRAFT 12
8. SOLAR THERMAL PROPULSION FOR A SMALL SPACE CRAFT 13
9. SPECIFICATIONS OF STP SPACE CRAFT 18
10. BENEFITS AND LIMITATIONS 19
11. CONCLUSION 20
12. REFERENCE 21
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