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NASA/TM—2015–218756

Development and Short-Range Testing of a 100

kW Side-Illuminated Millimeter-Wave Thermal

Rocket

Alexander Bruccoleri

Izentis LLC, Cambridge, Massachusetts

James Eilers

Ames Research Center, Moffett Field, California

Thomas Lambot

Carnegie Mellon University, Silicon Valley Campus, Moffett Field, California

Kevin Parkin

Carnegie Mellon University, Silicon Valley Campus, Moffett Field, California

April 2015
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NASA/TM—2015–218756

Development and Short-Range Testing of a 100

kW Side-Illuminated Millimeter-Wave Thermal

Rocket

Alexander Bruccoleri

Izentis LLC, Cambridge, Massachusetts

James Eilers

Ames Research Center, Moffett Field, California

Thomas Lambot

Carnegie Mellon University, Silicon Valley Campus, Moffett Field, California

Kevin Parkin

Carnegie Mellon University, Silicon Valley Campus, Moffett Field, California

 Insert conference information, if applicable; otherwise delete

 Click here: Press F1 key (Windows) or Help key (Mac) for help

National Aeronautics and

Space Administration

Ames Research Center

Moffett Field, CA 94035-1000

April 2015
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Acknowledgments

 The authors would like to thank NASA Ames Research Center, Defense

Advanced Research Projects Agency, ASRC Research and Technology Solutions, US Air

Force Research Lab, US Army High Energy Laser System Test Facility, General

Atomics, Carnegie Mellon University, Schaffer Corporation, CENTRA Technology Inc.,

Science Applications International Corporation, Leidos, Lightcraft Technologies Inc.,

Machine Technology Inc, Pratt Hobbies Inc., Balsa Machining Service, AirTanksForSale,

and Top Flight Recovery LLC for their contributions to this project.

Disclaimer

 The views, opinions, and/or findings contained in this article/presentation are

those of the author(s)/presenter(s) and should not be interpreted as representing the

official views or policies of the Department of Defense or the U.S. Government.

Level of Review: This report has been technically reviewed by technical management.

Available from:

This report is available in electronic form at

http:// www.izentis.com
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Development and Short-Range Testing of a 100 kW side-Illuminated

Millimeter-Wave Thermal Rocket

Alexander Bruccoleria, James Eilersb, Thomas Lambotc, Kevin Parkinc

a

Izentis LLC, Cambridge, 02142, United States of America

bNASA Ames Research Center, Moffett Field, 94035, United States of America

cCarnegie Mellon University, Silicon Valley Campus, Moffett Field, 94043, United States of America

Phone: (609) 902-8416, Email: alex@izentis.com

Topic: Microwave Power Beaming and Propulsion

Abstract. The objective of the phase described here of the Millimeter-Wave Thermal Launch

System (MTLS) Project was to launch a small thermal rocket into the air using millimeter waves.

The preliminary results of the first MTLS flight vehicle launches are presented in this work. The

design and construction of a small thermal rocket with a planar ceramic heat exchanger mounted

along the axis of the rocket is described. The heat exchanger was illuminated from the side by a

millimeter-wave beam and fed propellant from above via a small tank containing high pressure

argon or nitrogen. Short-range tests where the rocket was launched, tracked, and heated with the

beam are described. The rockets were approximately 1.5 meters in length and 65 millimeters in

diameter, with a liftoff mass of 1.8 kilograms. The rocket airframes were coated in aluminum

and had a parachute recovery system activated via a timer and Pyrodex. At the rocket heat

exchanger, the beam distance was 40 meters with a peak power intensity of 77 watts per square

centimeter. and a total power of 32 kilowatts in a 30 centimeter diameter circle. An altitude of

approximately 10 meters was achieved. Recommendations for improvements are discussed.

Keywords: Beamed Energy, Propulsion, Millimeter-Waves, Rocket

INTRODUCTION

 Millimeter-wave thermal rockets are a class of beamed energy rocket that utilize

millimeter waves to heat a ceramic heat exchanger, which then convectively heats a working

fluid. The hot working fluid is then expanded out a nozzle generating thrust similar to a

conventional chemical rocket. This system decouples the energy source from the propellant

enabling high heat capacity propellants to be used. For example, pure hydrogen propellant could

achieve a specific impulse of 700 s with a modest temperature of 1800 K [1]. This differs from

the first detailed beamed energy propulsion concept by Kantrowitz in 1972, which used lasers

and ablation [2].

The first laboratory scale microwave thermal heat exchanger was developed and

demonstrated by Parkin, which used a resonant cavity to focus and heat a single heat exchanger

tube [1]. Further laboratory studies with a single tube and resonant cavity were conducted by
2 Approved for Public Release, Distribution Unlimited

Bruccoleri et al. [3], which demonstrated microwave energy could be transferred to a working

fluid. The work described here is the progression from laboratory scale studies to a subscale

demonstrator, representing the first flight demonstration of a millimeter-wave thermal rocket.

The vehicle was built by the Millimeter-Wave Thermal Launch System (MTLS) project,

a project funded by the Defense Advanced Research Projects Agency (DARPA) and the NASA

Ames Research center. The goal of the this phase of the program was to advance the state-of-theart in millimeter-wave thermal propulsion technology and demonstrate a flight test rocket. This

work covers the design, construction and flight of the rocket. The millimeter-wave source was

the Active Denial System 0 (ADS 0) provided via a partnership with the US Air Force Research

Laboratory (AFRL) and the US Army High Energy Laser Test Facility (HELSTF). ADS 0 has a

95 GHz gyrotron with a total power of approximately 100 kW. The beam was focused and

directed via an array of millimeter-wave optics to a target of 30 cm diameter. It was designed and

built by Science Applications International Corporation (SAIC) and Leidos.

DESIGN CONSIDERATIONS

Based on initial assumptions of the beam transmission and absorption efficiency, Lambot

et al. [4] set the basic design for the size and operating characteristics of the millimeter-wave

absorbent heat exchanger to be mounted at the aft end of the rocket. The beam was focused down

onto the heat exchanger at the launch site at a distance of 40 meters with a peak power density of

77 W/cm2 with a total power of 32 kW in a 30 cm-diameter circle. The heat exchanger mass was

estimated to be 0.5 kg. Compressed Argon gas was used as the propellant for simplicity. Argon

is also dense and has a low specific impulse. The low specific impulse was desirable since it

maximized thrust for a fixed power. This can be observed from simplified models of thrust and

power. The power in the rocket exhaust jet Pj is;

2

j mue

2

1

P 1)

where m is the mass flow rate and ue

 is the velocity of the jet which is the specific impulse

multiplied by the gravitational acceleration (u I g e SP ). The thrust T , ignoring pressure effects

is;

e T mu 2)

Equation 2 can be substituted into equation 1 and rearranged to yield;

e

j

u

2P

T 3)

which shows that for a fixed power, thrust is maximized by minimizing the exhaust velocity or

specific impulse. To overcome gravity for a small demo flight, maximizing thrust is important

but it does not use propellant efficiently. For an actual mission to orbit, much higher-power

beams will be used, and the specific impulse will be maximized.

 The rockets needed to be pressurized remotely from the launch pad and then transported,

which meant the tanks and associated high pressure components needed to have large safety

margins for human handling. An aluminum tank, Catalina 9009, was chosen since it is a DOTcertified pressure vessel, making it a good choice for durability. The tank has an internal volume
Approved for Public Release, Distribution Unlimited 3

of 361 cm3 and a mass of 0.47 kg empty. The Catalina tank can hold 130 grams of Argon at

3,000 psi. The targeted rocket mass at liftoff was 1.5 kg due to the limited power of the beam on

the vehicle, which meant the airframe, recovery system and propellant feed system had to be

approximately 0.5 kg.

VEHICLE DESIGN

 The vehicle was designed to have the propellant tank in the airframe with an adapter,

valve, and pressure regulation system upstream of the heat exchanger. Ahead of the tank was a

bulkhead and a bay for the parachute. The timer, batteries, and switches were mounted in a small

bay within the nose cone. This design was modular such that different tanks, airframes and

electronics bays could be swapped and replaced. See Fig. 1 for a computer aided design (CAD)

depiction of the rocket.

Figure 1. CAD drawing of Rocket.

PROPELLANT FEED SYSTEM

The propellant feed system started with the main propellant tank, which had an adapter machined

from aluminum to mate with a ball valve. The adapter had a quick disconnect as a fill valve and a

5,000 psi burst disk for safety. The ball valve was the main release valve for the propellant and

was turned externally by a pneumatic actuator. See Fig. 7 for a picture of the actuator. Following
4 Approved for Public Release, Distribution Unlimited

the ball valve was a sonic orifice to reduce the pressure in the heat exchanger. See Fig. 2 & 3 for

a drawing and picture of the propellant feed system.

Figure 2. Drawing of propellant feed system.

Figure 3. Photograph image of propellant tank, adapter, valve and orifice. Note:

The orifice shown is stainless steel and was used for static testing. An

aluminum orifice was used for the flight vehicle.

 The drawback of using compressed gas as the propellant was a rapid reduction in tank

pressure and hence mass flow rate with time. Qualitatively, the tank is losing propellant mass

and the gas also cools as it does work during the expansion, which leads to a non-linear function

of tank pressure with time. The pressure in the tank can be derived analytically assuming an

isentropic process. This modeling provides good insight into the behavior of the gas; however, a

significant amount of heat will transfer into the gas since the tank is approximately three times as

massive as the propellant, which starts at ambient temperature. The following derivation shows

the basics of the pressure model [5].1

 To start, the change in density with respect to time is;

Vtank

m

dt

d

 4)

where m is the mass flow rate and Vtank is the tank volume. The mass flow rate is the product of

the density *, orifice area A , and gas velocity u ;

m A u 5)

1

 The derivation follows from the Iowa State University, class Ae311L homepage. “Blowdown of a

Pressurized Tank ́.
Approved for Public Release, Distribution Unlimited 5

where the superscript * represents sonic, or critical, conditions at the orifice. The tank pressure

starts at approximately 200 atm and flow can be assumed sonic through the orifice for the

majority of the tank blowdown. The gas velocity at Mach 1 is;

u RT t g 6)

where is the ratio of the specific heat at constant pressure to the specific heat at constant

volume, R is the specific gas constant and T t

g

 is the gas temperature at the sonic orifice.

Observing the relations of density and temperature for an isentropic process;

1

g

g

g g

0

t

T t T 0 7)

where T t

g

 and T 0 g are the gas temperature with time and initial gas temperature respectively,

and t

g

 and 0 g are the gas density with time and initial gas density respectively. The

temperature and density at the sonic throat are also functions of the tank stagnation conditions

and ;

1

2

T t T t

g g

 8)

1

1

g g

1

2

t t 9)

Equation 9 can substituted into equation 5, and equation 7 into 8 and then 6 and back to 5 to have

a function for the mass flow rate with density and the initial tank temperature. A differential

equation is found when substituting the mass flow rate into equation 4 and this can be solved and

substituted into the ideal gas law;

P RT 10)

where P is the pressure. The analytical solution of tank gauge pressure with respect to time is;

g g atm g Patm T t

V

R A

P t P P

1

2

2 1

1

tank

*

0 1

1

2

2

1

0 11)

where (0) Pg is the initial tank gauge pressure and Patm is the ambient atmospheric pressure. This

function drops quickly with time. To reduce the rapid drop in pressure, the tank volume can be

increased or the sonic orifice diameter can be reduced.

 The sonic orifice was used to reduce pressure since it was lightweight and simple to

implement. A proper pressure regulator for approximately 50 g/s argon would have weighed on

the order of 1 kg, far too much to fit into the 0.5 kg mass budget of the airframe, recovery system

and feed system. Assuming an isentropic process, the mass flow into the heat exchanger is

proportional to the tank absolute pressure g abs P (t) and sonic orifice area;

* m Pg

t

abs A 12)
6 Approved for Public Release, Distribution Unlimited

The mass flow leaving the heat exchanger is the same as entering during steady state and is

proportional to the heat exchanger absolute pressure HX abs P t , and the nozzle throat area At

;

HX abs At m P t 13)

The pressure in the heat exchanger is proportional to the tank pressure times the area ratio of the

orifice to the nozzle throat;

t

g abs

HX abs A

P t A

P t

*

 14)

This allows the pressure in the heat exchanger tubes to be reduced by using an orifice smaller

than the throat area. This solution was simple, though it still meant the mass flow rate was

changing with time, which was not suitable for efficient heat exchanger operation. An orifice

diameter of approximately 2.4 mm was used in the final rocket configuration.

AIRFRAME AND RECOVERY SYSTEM

 The main airframe was made from 65 mm diameter wound paper tubing and a plastic

nose cone from Estes, as they were lightweight, easy to work with and readily available. The

propellant tank was held in via plywood centering rings and a plywood bulkhead separated the

propellant tank section from the parachute. For stabilization, 4 fins were mounted on the side of

the heat exchanger with a 15 degree dihedral. The fins were made from balsawood laminated

with fiberglass to make a stiff, lightweight structure.

 A 36 inch 1.1 ounce nylon parachute made by Top Flight Recovery was used for

recovery. It was deployed with 1.3 gram of Pyrodex 33 ́ with a filament set off via an electric

timer, a MicroTimer2 by PerfectFlite. A Nomex cloth along with Estes wadding protected the

parachute from the heat of the Pyrodex. Range safety considerations required that a tether be

attached to the rocket to prevent it from exceeding 200 meters in altitude. The tether was

attached to a switch which fired the Pyrodex charge to deploy the parachute if the rocket

trajectory exceeded the tether length. A switch was used to reduce the strength requirement of

the tether and allowed for the use of Kevlar fishing line material. The canisters to hold the

Pyrodex, made by Pratt Hobbies Inc., had two filaments such that the timer and tether could be

wired independently. The timer and tether both had switches wired with their respective batteries

to turn them on and off right before flight. The timer also used a break-wire to initiate the

countdown. See Fig. 4 for a schematic of the timer and tether.
Approved for Public Release, Distribution Unlimited 7

Figure 4. Wiring diagram for timer and tether.

MILIMETER-WAVE RADIATION SHIELDING

 The rockets needed a reflective mm-wave radiation protection system to shield against

excessive heating in the regions outside of the heat exchanger. Aluminum foil (0.018 mm-thick),

and aluminum tape were used for this purpose. The foil was glued to the fin material during the

fiberglass lamination of the balsawood and glued to the body tubes via fiberglass resin or epoxy.

The tape was used to cover complex contours or holes in the foil. Aluminum¶s low electrical

resistivity enables thin films to be used to reflect almost all of the millimeter waves. A simple

model is presented by Buffler for the power reflected from a microwave susceptor film [6]. The

power reflected Pr

 is;

r 2

1 2r

1

P 15)

where r is the ratio of the film resistance to the free space impedance, Z0 (377 ohms). As an

example, a 500 nm-thick aluminum film, with a resistivity of 2.82E-8 *m [7], would reflect

99.94% of the beam.

 A test was done to test the efficacy of the protective shielding. Both bare and protected

paper tubing and fin material were exposed to the millimeter wave beam. The beam was still
8 Approved for Public Release, Distribution Unlimited

undergoing characterization at the time; however, the intensity was still significant and yielded

good qualitative results. See Table 1 for the beam parameters.

Parameter Value

Beam size 14 cm horizontal, 18 cm vertical

Peak power density 45 W/cm2

Average power density over 20 cm diameter circle 26 W/cm2,

Distance from mirror 24 m

Shot duration 0.67 s

Table 1. Beam parameters for radiation tests.

 The bare tubing and fin material both showed considerable burn marks from the beam,

while the foil protected material showed no signs of heating or damage. This built confidence

that the thin aluminum foil and tape were sufficient for shielding. See Fig. 5 for pictures of the

fin absorption test.

 (a) (b)

Figure 5. (a) Bare fin material before radiation test. (b) Bare fin material after

radiation test with observable burn mark in the center.
Approved for Public Release, Distribution Unlimited 9

FLIGHT VEHICLE AND LAUNCH

 The complete flight

vehicle had a liftoff mass of

1.81 kg and was approximately

1.4 meters in length. It had two

sets of green LED¶s for

utilization by the pointing and

tracking system of the beam.

Twenty feet of 1 inch square

aluminum extrusion from

80/20 Inc. was used as the

launch rail, and the rocket had

two plastic buttons to slide

along the rail. The rail was

supported with a mast system

from BlueSky Mast Inc. A rod

to grip the main ball valve on

the rocket was attached to the

stand and pneumatically

actuated to turn 90 degrees,

and weighted to fall away as

the rocket moved upward. The

actuator was based on similar

design by McCormick et al.

[8]. See Figs. 6 and 7 for

photographic images of the

rocket on the pad with the

pneumatic actuator installed.

 The heat exchanger for

propulsion is described in

detail by Lambot et al. [4], and

static thrust data is presented

here as well. The thrust of the

heat exchanger was measured

with and without the beam to

give quantitative data for the

thrust increase. See Fig. 8 for

the thrust data. The heat exchanger operating purely with cold gas had approximately 72 N peak

thrust and a total impulse of 36.3 N-s. The total impulse is calculated by integrating the thrust

with time curve,

t2

t1

Total Impulse Tdt 16)

where T is the thrust and t1 and t2 are start and finish times of the thrust. The heat exchanger,

with the beam on, had approximately 88 N peak thrust and 41.6 N-s total impulse, representing a

Figure 6. Photograph of rocket on pad.
10 Approved for Public Release, Distribution Unlimited

22% increase in peak thrust and 15% increase in total impulse. The thrust curves were used to

estimate the altitude of the rocket. The acceleration a was calculated by;

m

T mg

a 17)

where T is the thrust, m is the rocket mass, and g is the gravitational constant. The acceleration

was integrated with time to find velocity, which was integrated to calculate altitude. Air drag,

rail friction, actuator imperfections, and the change in rocket mass were all ignored in the

calculation. See Fig. 9 for a plot of predicted altitude versus time. The rocket with the beam on

has a predicted altitude of 15.1 m versus 10.8 meters without the beam.

Figure 7. Photograph of rocket showing the actuator attached to the main ball valve.

 The rocket flight of February 28, 2014 achieved an altitude of approximately 10 meters,

determined by observations from the ground. The rocket configuration was the same as that used

for the static thrust test data shown in Fig. 8. It successfully cleared the rail and was a free flying

rocket albeit with a tether attached. The rocket impacted the top of the rail as it came down

releasing the parachute, which became tangled on the launch rail's guy wires.

 The beam tracking system did not have sufficient control authority to track the rocket. In

lieu of accurate tracking, the heat exchanger was heated for 3 seconds prior to the start of gas
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11

Figure 8. Static thrust vs time for the heat exchanger with and without the beam

on. Load cell data taken February 26.

Figure 9. Predicted altitude versus time from thrust data. Rocket mass of 1.81 kg

used and assumed constant. No effects for drag or friction were used in the

calculation.
12 Approved for Public Release, Distribution Unlimited

flow which still added a significant amount of heat to the fluid. Proper tracking with the beam

would have added to the final altitude of the flight and is necessary for any beamed energy

propulsion rocket system. Rail friction was also a likely cause of a loss in altitude since the

plastic buttons that held the rocket to the launch rail had visible wear after the flight. The ball

valve actuator also perturbed the rocket such that it swung side to side as it climbed up the rail,

further adding to the friction.

IMPROVEMENTS AND CONCLUSION

 The rockets were too heavy to fly in excess of 100 meters for the amount of beam power

available, which is to say that too little power was absorbed by the rocket due to the compound

effects of inefficiencies in the beam director and heat exchanger, all of which are correctable

with further work. For the power actually absorbed, the entire rocket would need to be

downscaled by at least a factor of two. For example, a rocket mass of 0.9 kg has a predicted

altitude of 85 meters. The main propellant tank and high pressure components were all built to be

safe and pressurized with people in close proximity. A remote system to fill the rockets would

have allowed much lighter components. A burst-type valve would also have made the launch of

the rocket smoother and reduced rail friction. A pointing and tracking system with better

acceleration is also required since typically the rocket immediately flew out of the tracker field

of view.

 The rocket flights were a successful demonstrator of a beamed energy propulsion rocket

with millimeter-wave power. Specifically the rockets successfully converted ground-based

millimeter waves into heat within thin ceramic tubes. These tubes in turn heated a propellant,

which measurably increased the thrust of the rocket, increasing its maximum altitude during a

short range flight.

REFERENCES

1. K.L.G. Parkin, The Microwave Thermal Thruster and its Application to the Launch Problem. Doctorate Thesis,

Department of Aeronautics, California Institute of Technology, 2006.

2. A. Kantrowitz, Propulsion to orbit by ground-based lasers. Astronautics and Aeronautics 10, 74-76 (1972).

3. A. R. Bruccoleri, K.L.G. Parkin, and M. Barmatz, Axial temperature behavior of a heat exchanger tube for

microwave thermal rockets. Journal of Propulsion and Power, 23(4), 889-894 (2007).

4. T. Lambot, L. Myrabo, D. Murakami and K.L.G. Parkin, 3Development and Testing of a Refractory MillimeterWave Absorbent Heat Exchanger ́, International High Power Laser Ablation and Beamed Energy Propulsion,

Santa Fe, NM, 21-25 April 2014.

5. Iowa State University, AerE 311L Homepage, Blowdown of a Pressurized Tank, Website, March 2014,

http://www.public.iastate.edu/~aero311l/blowdown.pdf

6. C. R. Buffler, A Simple Approach to the Calculation of Microwave Absorption, Transmission, and Reflection

of Microwaves from a Susceptor Film. Microwave World, 12(3), 5-7 (1991).

7. Hypertextbook, Students of E. Glenn, Resistivity of Aluminum, Website, March 2014,

http://hypertextbook.com/facts/2004/ValPolyakov.shtml

8. A. McCormick, E. Hultgren, M. Lichtman, J. Smith, R .Sneed, and S. Azimi, Design, Optimization, and Launch

of a 3 ́ Diameter N2O/Aluminized Paraffin Rocket, 41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference

& Exhibit, Tuscon Arizona, July 10±13, 2005.