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Beam Powered Propulsion Systems

Nishant Agarwal University of Colorado, Boulder, 80309

While chemical have dominated , other forms of propulsion based on nuclear power, electrostatic and magnetic drives, and other principles have been considered from the earliest days of the field with the goal to improve efficiency through higher exhaust velocities, in order to reduce the amount of fuel the rocket needs to carry. However, the gap between technology to reach from surface and from orbit to interplanetary travel still remains open considering both economics of time and money. In addition, methods have been tested over the years to reach the orbit with single stage rocket from ’s surface which will eventually result in drastic cost savings. Reusable SSTO offer the promise of reduced launch expenses by eliminating recurring costs associated with hardware replacement inherent in expendable launch systems. No Earth- launched SSTO launch vehicles have ever been constructed till date. Beam Powered Propulsion has emerged as a promising concept that is capable to fulfill all regimes of space travel. This paper takes a look at these concepts and studies the feasibility of Propulsion for possibility of SSTO vehicle.

Nomenclature At = nozzle throat area C* = characteristic velocity Cf = Coefficient Cp = Specific heat constant Dt, Dexit = Diameter of nozzle throat, Nozzle exit diameter g = acceleration due to gravity Gamma = Specific heat ratio LOX/LH = /Liquid Hydrogen MR = mass ratio Mi,Ms,Mp = total mass, structural mass, mass, payload mass Mpayl = payload mass Mdot = flow rate Isp = Ve = exhaust velocity

I. Introduction n the year 2015 launching to orbit is still done in the same way as was done 6 decades ago. As described by the I rocket equation, this is due partly to the structural limits of existing materials, and partly to the limited specific impulse (Isp) of chemical , which have reached a practical limit of 460 seconds. When considering how much propellant is consumed by launch vehicles, one realizes that present day propulsion systems are the means to, and at the same time a bottle neck to the access to space. Currently, the spectrum of option is bimodal: either large thrust with lower specific impulse, as with chemical propulsion or high specific impulse at the expense of low thrust, as with electric propulsion. The only fully developed option available to get into orbit is Chemical Propulsion, which turns out to be a very costly affair to launch payloads into orbit. Although, this have been realized for many years now, the recent increase in demand in the market and the interest of general community in space access has highlighted this issue like never before resulting in the involvement of both government as well as private companies to explore new options for cost reduction. Innovative economic models are been developed in addition to the development of technology. It is been realized that chemical rockets have nearly reached their technological maturity,

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with improvements mainly being sought in the areas of reliability, cost, diagnostics or controllability. Some exceptions are the potential for active control of combustion instabilities, or the possible miniaturization of these rockets.

Approach towards beyond-orbital space travel has started to change in recent years with emergence of electric propulsion which shows great promise for the futuristic interplanetary travels. Instead of the short, powerful burn and fast acceleration of a chemical engine, such advanced engines burn for long periods of time, providing a continuous gentle nudge that builds up. Most such schemes cannot be used to propel payloads from the surface of the Earth into orbit, but they provide great advantages for interplanetary flight. Since these engines do not use chemical reactions, they do not need to carry an oxidizer like liquid oxygen. This can simplify system plumbing.

Nuclear propulsion has also been a prominent candidate as the source of energy for rocket propulsion been studied for many years. Despite the high Isp of 700–950 seconds, solid core nuclear rockets to date still cannot reach orbit because the nuclear reactor is heavy and makes the thrust-to-weight ratio (T/W) too low. The low T/W means that if the rocket can get off the ground at all, it spends a long time accelerating to orbital velocity. During this extra time, the ascent trajectory accumulates greater drag and gravity losses, which increase the total ΔV of the ascent, decreasing the payload fraction and greatly reducing the advantage of higher Isp.

This calls for a study of other advanced propulsion systems that can fill the voids left by the above stated systems.

II. Theory Beam-powered propulsion is a class of or propulsion mechanism that uses energy beamed to the spacecraft from a remote power plant to provide energy. The beamed energy propulsion concept is similar to nuclear or solar thermal propulsion, where a working fluid is heated in a heat exchanger and then expanded through a nozzle. In the beamed energy concept, the heat exchanger is powered by an electromagnetic beam produced on the ground and propagated through the atmosphere – the power source and its accompanying weight is left on the ground. The major theoretical advantages of this concept are high Isp, high thrust to weight ratios, and the accompanying payload fraction and structural margin increases.[21] The beam would typically either be a beam of or a . are subdivided into either pulsed or continuous beamed. As high-power laser technology continues to mature the possibility of using a laser to generate rocket thrust for propulsion applications grows more feasible. The concept was first introduced by Kantrowitz, more than thirty years ago, and was experimentally demonstrated by Krier and by Myrabo. As with any thermal propulsion system, the efficiency of conversion of laser beam energy into the kinetic energy of propellant gas is a critical figure of merit. In addressing laser thruster performance it is useful to consider both the absorption efficiency as well as the propulsion efficiency. In the context of laser , the propulsion efficiency is a measure of how much absorbed energy appears as kinetic energy of the propellant at the nozzle exit.

In addition, microwaves can be used to heat a suitable heat exchanger, which in turn heats a propellant. This can give a combination of high specific impulse (700–900 seconds) as well as good thrust/weight ratio (50-150).

III. Categories of Beam Powered Propulsion Systems 1) Thermal Systems – A laser thermal system is both a beamed power system and a thermal system. The thermal energy source is a laser, which heats the working fluid through a heat exchanger. The working fluid is then expanded through a nozzle to produce thrust. Depending upon the laser power thrust to weight ratios similar to that of chemical propulsion can be achieved with very high specific impulse. A variant of the energy source is Microwave powered system. The laser propelled heat-exchanger (HX launcher) concept was suggested by Kare (1995). The heat exchanger operates in a laminar regime by analogy to designs used for integrated circuit cooling. The heating channels are 200 μm wide by 2 mm deep and 3 cm long, raising the hydrogen propellant temperature to 1300 K; corresponding to an Isp of ~ 600 s. Kare estimates the heat-exchanger mass to be 125 kg for a 5.4 ton vehicle carrying a 122 kg payload, corresponding to a payload fraction of 2.26%.

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Figure 1 Laser Powered Thermal Propulsion System Concept Design

2) Electrothermal propulsion, as opposed to thermal propulsion, uses the plasmadynamic breakdown of the propellant itself to absorb incoming radiation, thereby heating it to a very high temperature. Early experiments revealed that forming in a propulsive duct has a tendency to propagate toward the source of radiation. This effect was first studied by Raizer (1972)

Figure 2 Laser Powered Electrothermal propulsion

Thus far the majority of electrothermal propulsion work has focused on in-space propulsion, which operates at far lower pressures and mass flow rates than needed to produce a T/W ratio large enough for launch. The scalability of these techniques to the high pressure, high mass flow rate regime need for launch has yet to be demonstrated.

3) Molecular Absorption Propulsion – In contrast to the electron inverse bremsstrahlung absorption of plasma formation, molecular absorption is achieved via excitation of an internal rotation or vibration mode. The molecular absorption approach for laser propulsion was previously identified by Caledonia in 1975. By acting on the propellant itself or a seed molecule, heating of the flow can be achieved in subsonic or supersonic flows without the use of plasmas. Because area variation along the flowpath can be used to offset temperature increases, more gradual addition of energy at lower temperatures could result in propulsion systems with less stringent cooling requirements and lower frozen-flow losses. As the propellant gas flows through and around the stationary plasma high bulk temperatures are sustained which can be in excess of 10,000 K in gases such as argon. Stable LSPs were created and observed by Keefer, who report absorption efficiencies as high as 86%. [1] 3

4) Rectenna Based Concepts - A rectenna is a rectifying antenna, a special type of antenna that is used to convert microwave energy into direct current . They are used in wireless power transmission systems that transmit power by radio waves. Rectennas have been developed over a number of years in large part due to the efforts of Brown (1984; 1992), who demonstrated the flight of an using 2.45 GHz microwaves in 1964. [2] The microwave concept of Myrabo (1995) shown in Fig. 3 uses rectennas to produce a DC current that drives a series of Lorentz force accelerators around the periphery of the craft. The 1400 kg helium-filled lenticular craft is 15 m in diameter and receives a 5.6 GW, 35 GHz microwave beam from above. A portion of the beam is reflected and focused ahead of the craft to form a microwave-induced aerospike that reduces heating and drag on the vehicle.

Figure 3 RF-Powered Lenticular Craft Figure 4 RF-Powered Lenticular Craft

However, rectennas are currently limited to power densities of less than 1 KW/m2 at 2.45 GHz, which together with the high weight of ~1 kg/kW for DC power processing equipment, limits their use on highly energetic vehicles. Myrabo calculates that the microwave lightcraft requires a 35 GHz rectenna to operate at power densities 10,000 to 40,000 times higher than the present state of the art, and suggests a high-pressure helium-cooled silicon carbide rectenna array as a possible solution.

Figure 5 Microwave Powered Lightcraft concept

5) Ablative Laser & Microwave Propulsion - The laser lightcraft shown in Figure 6 has a diameter of 12.2 cm and weighs roughly 50 grams. It is powered by a 10 kW CO2 laser. A parabolic on the underside of the craft focuses the beam into the engine air or propellant. The pulsed laser heats the air, causing it to break down into a plasma. The plasma strongly absorbs the incoming pulse, heating to roughly 18,000 K before exploding from the annular underside region, generating thrust. The ablative microwave lightcraft concept 4

of Myrabo and Benford (1994) shown in Fig. 5 uses microwaves rather than lasers. The concept is airbreathing but switches to an on-board hydrogen supply for the later stages of ascent (in ). In 2003 ablative propulsion using a 1 MW 140 GHz gyrotron was demonstrated. At present, ablative propulsion converts only a few percent of the incident energy to thrust; however, propulsive cycles using double pulses appear to offer efficiencies as high as 40%.

Figure 6 Ablative Laser Powered Lightcraft Figure 7 Arrangement of Propellants

6) Laser/Electric Propulsion - This is for spacecraft with electric thruster [ion, magneto palsmadynamics]. The photovoltaic arrays of the spacecraft are then powered not by , but by a monochromatic laser of much higher efficiency. Higher laser power density saves weight by allowing smaller solar arrays. This application is most meaningful when orbiting laser stations exist.

7) Solar sails powered by lasers- The physicist James Benford gives insight into the potential benefits: The fundamental attraction of high power beams for space is simple: microwaves and lasers can carry energy and (both linear and angular) over great distances with little loss. lose a negligible energy when radiated out of a potential well such as Earth’s. Microwave energy in space is cheap. The point where the sail no longer receives the full output of the laser is termed the thrust run point. After this the intensity of the beam received by the sail will also drop off inversely with distance.

8) Microwave Beam Powered Propulsion [3] Starwisp is a hypothetical unmanned interstellar design proposed by Robert L. Forward in 1985. It is propelled by a microwave sail, similar to a in concept, but powered by microwaves from a man-made source. The probe itself would consist of a mesh of extremely fine carbon wires about 100 m across, with the wires spaced the same distance apart as the 3-mm wavelength of the microwaves that will be used to push it. The spacecraft would be accelerated by a 10–50 GW microwave beam at 1,130 m/s2 or 115 g and would attain a cruise velocity of 60,000 km/s or 0.2c with a specific impulse of around 6,000 s. Forward speculated that such a probe could even be accelerated to near light speed, due to its low mass. Figure 8 shows the concept design. Constructing such a delicate probe would be a significant challenge. One proposed method would be to "paint" the probe and its circuitry onto an enormous sheet of plastic which degrades when exposed to ultraviolet light, and then wait for the sheet to evaporate away under the assault of solar UV after it has been deployed in space.

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Another proposed method noted that the Starwisp probe wires were of the same physical scale as wires and circuit elements on modern computer microchips and could be produced by the same fabrication technologies as those computer chips. The probe would have to be built in sections the size of current chip fabrication silicon wafers and then connected together.

Figure 8 Microwave Propelled Sail

The sail acceleration from momentum produced by a power P on a thin sail of mass m and area A is [4]

IV. Comparison of Laser v/s Microwave Propulsion One approach to beamed energy propulsion uses a solid heat exchanger to absorb energy from a distant source and transfer it to a working fluid. Systems of this type can be designed using either microwave or laser sources. In general, microwave sources have been expected to be less expensive than lasers for a given power, but to be more limited in range and/or energy density. With the development of high power millimeter-wave sources and low-cost diode laser arrays, both assumptions are open to question. The microwave system has a much shorter range and acceleration time than the laser system. This implies higher accelerations and higher loads on the vehicle and payload, but also means the microwave system can launch more payloads in a short period of time, or more total mass over a long period. Both systems have enormous total launch capacity compared to current launchers, so capital cost and payload size are probably more important than launch rate in the near term, but launch rate may be significant for some users. The microwave system scales favorably to larger payloads and higher powers, since a simple scaling of the system increases power and reduces beam divergence (increasing the useful range) while the laser system beam divergence

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does not improve with power. Conversely, the microwave system scales down poorly, either for smaller systems or for development and flight demonstrations. In a study done by Kare and Parker in 2006[5], Microwave- and laser-based beamed energy launchers both have technical unknowns that place them at TRL 2 to 3. For modular laser sources, there is a single major technical issue – scaling unit laser power to ~10 kW or greater. For microwave systems, the clearest technical issue is phase locking of large gyrotron arrays, but the microwave phased array aperture, and specifically phase sensing and control on the required scale, are also major challenges. Laser systems have a modest advantage in end-to-end efficiency, mainly due to difficult-to-eliminate array losses in a steerable-dish phased array. Other losses need further analysis, but appear to be similar for the two systems. The laser system cost is dominated by the lasers themselves. For the microwave system, there is no dominant cost element likely to be affected by technical progress or production scaling, but the greatest gains in system performance vs. cost would come from lowering the cost of large millimeter-wave antennas.

A. Transatmospheric Microwave Beam Propagation Beamed-energy concepts are limited to frequencies at which the atmosphere is transparent. Near total absorption by H2O in a large portion of the far infrared spectrum divides these two viable beamed-energy concepts. Beaming energy sufficient to propel a ton into LEO requires more than 100 MW of energy transmission through the atmosphere. Microwaves have two main advantages: First, at microwave wavelengths, atmospheric turbulence is not the major problem as it is with lasers. Second, commercially available microwave sources are already capable of generating this level power output whereas today’s most powerful lasers are still an order of magnitude weaker. The advent of submillimeter wavelength astronomy has highlighted the existence of locations with particularly low atmospheric water content, opening up new microwave transmission windows from 35 to 300 GHz and sometimes beyond. Microwave frequency determines the maximum beam energy density via the constraint of atmospheric breakdown. Free-space atmospheric breakdown is an electron avalanche process. The beam frequency has a disproportionate effect on the breakdown intensity; for example, a 300 GHz beam can achieve 1000 times the power density of a 3 GHz beam, assuming that it is constrained at the altitude of minimum breakdown intensity. By moving to higher frequency in this way, the energy density can enter the energetic regime needed for .

V. Microwave Propulsion Although, a lot of research and intrastructure development is needed for considering Microwave Propulsion as a viable regular candidate for Propulsion system, like any other leap in technology here are a few reasons that corroborates the effort –

1) Higher Isp 2) Lower Structural Mass – more room for factor of safety – more payload (Regular Cyrogenic Chemical Propulsion systems have a failure rate of 4%.) [23] 3) The highly nonlinear relationship between structural margins and probability of failure raises the possibility of reusability, leaving a vehicle with mass to spare for an increased payload fraction using only a single stage. The combined effect of lower structural cost, greater payload fraction and higher flight rate can profoundly alter the economics of launch, minimizing the need to boost launch demand in order to solve the launch problem.

B.A Microwave Thermal Propulsion System [6] Using an array of high power microwave sources directed at the rocket underside, propellant is heated within hundreds of small channels running through a microwave absorbent coating. Microwave thermal launch is possible due to the recent advent of high power microwave sources in the millimeter wavelength range for which microwave launch is economically feasible. a) Thruster - It is a heat exchanger with microwaves directed from the ground on its base which heats a hydrogen propellant flowing through thousands of small channels underneath by convection and not combustion enabling the use of a single propellant and reduced system complexity. b) Resonant Absorption: Silicon Carbide Thruster

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SiC is a nontoxic, highly effective substance for enhancing microwave-induced heating with a microwave ablation system in vitro. Silicon carbide is used as a heat-enhancing agent in microwave ablation [7] It is also a highly oxidation resistant refractory material used on the re-entry of the . Channel pressure, mass flow rate and the high temperature material properties of resistivity, yield strength, and creep rate all drive the structural design of the heat exchanger channels. Metallic refractories such as W, WC, ZrC, HfC, TaC and TiB2 are of interest for susceptor-based approaches and as load-bearing channel liners for high temperature designs, where ceramics such as silicon carbide (SiC) would soften and cause a rupture in the channel wall. For example, the flexural strength of SiC increases with temperature; however, so does the strain rate, whose variation with temperature is approximated by,

[8] where is the creep strain, C is a constant dependent on the material and the particular creep mechanism, m and b are exponents dependent on the creep mechanism, Q is the activation energy of the creep mechanism, σ is the applied stress, d is the grain size of the material, k is Boltzmann’s constant, and T is the absolute temperature.

[9] Figure 9 Thermal Conductivity and Liner Thermal Expansion for SiC c) Microwave sources- [5] In principle, any number of gyrotron oscillators can be phase-locked to produce a common frequency. Phase locking of microwave oscillators has been used since the 1940s and has reached powers exceeding 1 GW. Phase locking of gyrotron oscillators has been analyzed, but it has been demonstrated only at low power and with small numbers of oscillators. Gyrotrons have a small advantage in DC power supply efficiency because they operate at high voltage; ohmic losses and semiconductor voltage drops are more significant in low-voltage, high-current diode laser supplies.

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Figure 10 RUZA: 34 GHz, 1MW pulse, 0.01 duty gyroklystron-based semi-active phased array radar [11]

BC. Practical Applications and Experiments The concept of Microwave propulsion have been in discussion for years. The applications and experimentations are not retricted to the Aerospace Industry. 1) A lightcraft test in 2000 at White Sands Missile Range in New Mexico. (Figure 11) 2) Laser airspike experiments are underway in Brazil at the Henry T. Nagamatsu Laboratory of Hypersonics and Aerothermodynamics (Figure 12)

Figure 11 Figure 12 3) Stability of Lightcraft Testing at Air Force Research Laboratory (Figure 13) 4) Chevrolet Chaparral 2X Vision concept at LA 2014 (Figure 14)

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Figure 13 Figure 14 5) NASA Beam Propulsion Challenge winners, LaserMotive LLC was awarded $900,000 in the 2009 Power Beaming Challenge (Figure 15) 6) The demonstration of a Gyrotron - essentially a - at the Naka Fusion Institute of the Japan Atomic Energy Agency (Figure 16)

Figure 15

Figure 16

VI. Microwave Propulsion v/s From the 1950s to the 1970s a series of over 30 nuclear thermal thruster tests, conducted as part of the ROVER, KIWI and NERVA programs, demonstrated hydrogen heat exchangers approaching an Isp of 825 seconds experimentally, operating at power levels exceeding 1 GW (for high thrust), and for durations of longer than an hour. Later particle bed designs constructed as part of the Timberwind program demonstrated heat exchanger exit temperatures of 3200 K, corresponding to an Isp of around 900 seconds. There are presently two performance-limiting factors for nuclear 10

thermal rockets: The first is the thrust-to-weight ratio, which is limited by the need to carry uranium fuel, shields, reflectors and moderators. The second is the neutronic properties and hydrogen compatibility of refractory materials. For the microwave thermal thruster, materials are constrained by microwave properties and hydrogen compatibility. Furthermore, the microwave thermal thruster uses no pressurized control drum, so the channel wall materials themselves must hold the propellant pressure against vacuum on the other side. Microwave thermal thrusters inherently operate at peak microwave energy absorption. Unlike the breakdown of a high power laser mirror or the reactor meltdown, many types of malfunction in the microwave thermal thruster should only serve to decrease the microwave absorption, and slow or stop the progression of a fault. [18]

A. Potential constraints in development of this technology There are two specific engineering reasons why a microwave thermal launch system might not be possible: The first is that a suitable way to construct a refractory hydrogen heat exchanger may not be found. Past experience with nuclear thermal rockets suggests otherwise, but in this case the pressure difference across channel walls is far greater. The second reason is that a suitable way to phase lock gyrotrons may not be found, but experience with every other class of vacuum-type microwave source suggests otherwise.

VII. Mission Analysis for Microwave Propulsion System - SSTO

A. Single Stage to Orbit – SSTO SSTO vehicles have some potential benefits over multi-stage launch systems. By combining the system into one vehicle, SSTOs are more operationally flexible because they do not require the assembly of multiple vehicle components. Also, SSTO vehicles may have smaller wetted areas, and thus a lower amount of maintenance-demanding TPS, reducing the maintenance hours required to turn around a RLV after returning from orbit. While these benefits make SSTO RLVs (Single Stage To Orbit – Reusable Launch Vehicles) appealing, multi-stage launch vehicles have been standard for over 6 decades. With the advent of new propulsion technologies SSTO seems to be viable in the near future.

Figure 17 Payload Mass Fraction v/s Structural Mass Fraction 11

Table 1. Key SSTO system requirements.

- Intact abort any time during flight - Rapid, low-cost turnaround - Medium payloads deployed and/or retrieved

- Manned and/or unmanned operation - Rocket propulsion as the prime mover

B. Proposed SSTO Vehicles

Figure 18 X-33/Ventura Star (Lockheed Martin) - Project cancelled in 2001. - Propellant: LOX/LH - T/W: 3.01 - mpayl/m: 0.02

Figure 19 (Reaction Engines Ltd.) - Status: Under Development - Propellant: LOX/LH - T/W: 1.2 – 3 (at burnout) - mpayl/m: 0.04

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Figure 20 Delta Clipper Advanced (McDonnell Douglous) - Status: Last flown in 1996 - Propellant: LOX/LH - T/W: 3.17

Figure 21 Roton ATV ( Company) - Status: Closed 2001 - Propellant: LOX/Kerosene - T/W: <150 - mpayl/m: 0.018

Figure 22 Kankoh-Maru (Japanese Rocket Society) - Status: Proposed in 1995/No current status known - Propellant: LOX/LH2 - T/W: 1.56 - mpayl/m: 0.09

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Figure 23 (DRDO/ISRO) - Status: Under Development - Propellant: LOX/LH2

- mpayl/m: 0.04

C. Using Hydrogen as a Propellant On the technical side there are significant and fundamental limitations in chemical propulsion–the energy density of propellants is limited by the energy of chemical bonds, and hence the energy density of chemical reaction known. For all intents and purposes that is the H2/O2 reaction, which at high pressure has an Isp of 460 seconds. For H2/O2 systems at high pressure. The highest chemical Isp achieved was 523 sec in vacuum using a Li/H2/F2 tripropellant combination, but this and other combinations have all proven to be highly impractical, volatile, and economically infeasible for mass production. Since Isp is related to the exhaust velocity by Isp gVe, this can be substituted into the enthalpy conservation equation and rearranged using well known perfect gas relations to give,

The Isp is proportional to the square root of the propellant temperature Tc prior to nozzle expansion. Hydrogen is the optimum propellant for a propulsion system whose peak temperature is limited by the melting and softening points of materials, because the enthalpy at any given temperature is higher than any other gases and liquids that could be used for propulsion. That is to say it has the highest enthalpy at the maximum material temperature of the heat exchanger, and therefore the most energy per unit mass to convert into kinetic energy, hence exhaust velocity, hence Isp.

D. Designing SSTO Vehicle With the recent increase in the demand and interest of governments and industry in small payloads and to be sent to LEO and GEO, it is almost becoming an immediate necessity to find ways to lower the cost of such misisons. In this section, a rough attempt is made to design a SSTO vehicle with an aim to put a payload of 100 kg in the Lower Earth Orbit. For simplicity, many assumptions have been made under reasonable limits. Assumptions: 1) T/W ratio required is 1.3 at lift off (This is taken as a standard for designing the system) 2) Specific Impulse of 900s for Microwave and Nuclear Propulsion Systems, 450 s for Chemical Propulsion System 3) Liquid Hydrogen is used as a fuel so mass estimations have been made for cryogenic engine 4) Delta V required to reach LEO is 9.60 km/s, including Drag and Gravity losses. 5) Heat Exchanger used in Microwave Propulsion is made out of SiC 6) Propulsion System Components

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Table 2. Chemical Propulsion Microwave Propulsion Nuclear Propulsion . Propellant Tanks (LOX/LH2) . Propellant Tank (LH2) . Propellant Tanks (LH2) . Thrusters (Combustion Chamber . Thrusters (Nozzle) . Thrusters (Nozzle) & Nozzle) . Plumbing . Plumbing . Plumbing . Pressure Feed System . Feed System . Feed System () (TurboPumps/ Pressure Fed) (TurboPumps/Pressure Fed) . Thrust Control System *optional . Thrust Control System . Gimbling System . Thrust Control System . Gimbling System . General Structure . Gimbling System . General Structure . Avionics . General Structure . Avionics . Avionics . Nuclear Reactor . Heat Exchanger . Shielding (Internal & External)

(* Other Assumptions if any have been stated below)

a) Microwave Propulsion System -

Chamber Pressure: 20 bars

Gamma = 1.4054 M= 2 G= 9.81m/s2 R=8314 J/mol/K For Isp = 900, Tc=2704.56K

DeltaV=Isp*g0*ln(MR) MR=2.966 MR=(Ms+Mp+Mpayl)/(Ms+Mpayl) Heat Exchagers SiC density= 3.21g/cm3. Heat Conductivity of about = 25 W/m/K Heat Conductivity for Hydrogen = 0.1805 W·m−1·K−1

Ms= mass of heat exchanger = 200 kg mass of propellant tank = 10% of mp mass of instrumentation = 20 kg mass of other structural parts = 100 kg Mpayl=100 kg Mp= 1027.78 kg

Mi= Ms+Mp+Mpayl = 1550.56kg

T = g0*Mdot*Isp Thrust required = 19774.27N Initial Acceleration = 12.75 m/s2 Final Acceleration = 37.82 m/s2 Mdot = 2.24 kg/s

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At = 0.0055 m2 Dt = 0.084 m

C or Ve= 8829 m/s C*= 4910.71 m/s Cf= 1.8

For this value of Cf at gamma=1.4054, expansion ration is infinite. We consider a very high expansion ratio of say, 100. Area of nozzle at exhaust = 0.55 m2 , Dexit= 0.84 m.

Jet Power of 87.29 MW. Assuming, thrust efficiency of 0.80, Input power will be 109.12 MW. Further assuming that the efficiency of heat transfer from Heat Exchanger to the Propellant is about 0.60, we need the available power of about 181.86MW. At 30MW/m2 for SiC, we need roughly 2m * 3m of heat exchanger surface. (Very close to our assumed value for the Heat Exchanger) (However, with currently available technology, it means that we need to provide about 2 times this energy from ground.)

b) Chemical Propulsion - Chamber Pressure: 10 bars MR=8.80 Ms= mass of turbo pumps and plumbing = 150 kg mass of propellant tank = 10% of mp mass of instrumentation = 20 kg mass of other structural parts = 5% mp (Payload mass is about 30-35% of the structural mass for small bipropellant rockets) (*spaceflight101) mpayl=100 kg mp= 12388.24 kg mi= ms+mp+mpayl = 14556.97 kg

Thrust Required = 185645.04 N Initial Acceleration = 12.75 m/s Final Acceleration = 85.60 m/s mdot = 42.05 kg/s At= 0.132m2 Dt= 0.4m C or Ve=4414.5m/s 16

C*=3152.21m/s Cf=1.4

Nozzle Exit Area = 13.2 m2, Dexit= 4.1 m

Jet Power of 409.76 MW Assuming 80% of thrust efficiency, available power should be = 512.21 MW

c) Nuclear Propulsion -

MR=2.966 Assumption: (Structural mass for a nuclear rocket has been derived by scaling down the NERVA rocket) NERVA – Reactor Mass= 5500kg, Other Systems =2500kg, Internal Shielding = 1500kg, External Shielding = 4500 kg; Power = 2500MW) [19] From our previous calculations for Microwave Powered Rocket, we take the required power to be 110 MW. Assuming 50% efficiency of the nuclear power plant, the aount of power to be generated should be around 220 MW. The mass of the system for this power generation will then be = 1232 kg. ms= mass of nuclear plant system = 1232 kg mass of propellant tank = 10% of mp mass of instrumentation = 100 kg mass of other structural parts = 100 kg mpayl=100 kg mp= 3748.96 kg

mi= ms+mp+mpayl = 5655.96 kg

Thrust required = 72130.42 N Initial Acceleration = 12.75 m/s2 Final Acceleration = 37.82 m/s2 mdot = 8.17kg/s At = 0.02 m2 Dt = 0.16 m C or Ve= 8829 m/s C*= 4895.96 m/s Cf= 1.8 For this value of Cf at gamma=1.4054, expansion ration is infinite. We consider a very high expansion ratio of say, 100. Area of nozzle at exhaust = 2.0m2 , Dexit= 1.60m.

Jet Power of 318.42 MW. But the input power is only 110 MW (Available Power of 220 MW). On iterating the values for power input and power output, it is concluded that we can not design a nuclear propulsion engine with T/W>1 that would be capable of liftoff from earth with the current technology.

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Table 3. Mass and Power Requirements for the system Type Isp Ve Mdot System Thrust Jet Power System Power Specific Specific (s) (m/s) (kg/s) Mass (kg) Required (N) (MW) Required Power Energy (MW) (MW/kg) (MJ/kg) MicroWave 900 8829 2.24 1550.56 19774.27 87.29 181.86 8.5 39 Chemical 450 4414.5 42.05 14556.97 185645.04 409.76 512.21 28.0 13

Table 4. Mass Budget and Mass Ratios Type System Propellant Dry Propellant Structure Payload Mass (kg) Mass (kg) Mass Mass Ratio Mass Ratio Mass Ratio (kg) MicroWave 1550.56 1027.78 522.78 0.66 0.33 0.065 Chemical 14556.97 12388.24 2168.73 0.85 0.15 0.007

Table 5. Size comparison for the designed system Type Nozzle Size Propellant Tank Throat Area Exit Area Propellant Tank Volume Tank (m2) (m2) (m3) Pressure MicroWave 0.055 0.55 14.51 2 MPa Chemical 0.132 13.20 9.05(LOX) 29.14(LH2) 1 MPa

Table 6. Mass properties typical cryogenic launcher stages [20] Vehicle Stage Dry mass Propellant mass Total mass Dry mass fraction2 Dry Mass/Propellant [ton]1 [ton] [ton] [-] Mass

Saturn V S-II 38 427 465 0.082 0.09

Ariane 5 H155 12.6 156.2 168.8 0.075 0.08

Saturn V S-IVB 9.9 104.4 114.3 0.087 0.09

H2 Stage 1 11.9 86.2 98.1 0.121 0.14

Titan IV Centaur 3 23 26 0.115 0.13

H2 Stage 2 3 16.7 19.7 0.152 0.18

Ariane 4 H10 1.2 10.8 12 0.100 0.11

VIII. Conclusion Out of the technologies currently available to us, Beam Powered Propulsion System seems to be a highly potential candidate for space propulsion. It covers a broad regimes of mission capabilities including Earth to LEO, LEO to GEO and especially interplanetary travel. Although, most of the technology discussed in this paper have been in concept stage, there is no reason why these concepts can not be converted to practicle technology. Ofcourse a giant leap is to be taken from the decades old inertial of Chemical Propulsion to Beam Powered Propulsion. As for the present scenario, Beam Powered Propulsion may hold the key to the future of space exploration and might be the milestone in the field of Rocket Propulsion.

Acknowledgment I would like to thank Dr Lakshmi Kantha of University of Colorado at Boulder Department of Aerospace Engineering for his assistance with this report. Also, I would like to acknowledge the authors of various Research Papers, Thesis, Dr. Parker, whose reports were refered for this paper. 18

References

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