solar thermal propulsion

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


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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http://www.grc.nasa.gov/www/RT2001/5000/5490wong2./html

http://www.highway2space.com/

http://www.inspacepropulsion.com/tech/solar_therm.html

http://www.vectorsite.net/trarokt2.html

http://www.osti.gov/energycitations/product.biblio.jsp?osti_id=7112464


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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solar thermal propulsion

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