progress on hyper velocity railgun research for launch to space

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IEEE TRANSACTIONS ON MAGNETICS, VOL. 45, NO. 1, JANUARY 2009 381

Progress on Hypervelocity Railgun Research for Launch to Space

Ian R. McNab , Fellow, IEEE

Institute for Advanced Technology, The University of Texas, Austin TX 78712 USA

The Universities of Texas, Minnesota, and New Orleans, and Texas Tech University are undertaking research supported by the Air

Force Office of Scientific Research on critical issues for a launch to space from a railgun carried on an airborne platform. The Universityof Texas at Austin is studying techniques to achieve hypervelocity with a goal of 7 km/s: So far, 5.2 km/s has been achieved in a 7-maugmented railgun using a preinjected plasma armature. Texas Tech University is studying distributed power feed concepts that willimprove the efficiency of launch for a long railgun: So far, 11 km/s has been achieved with a plasma arc in a five-stage system. TheUniversities of Minnesota and New Orleans are investigating the aerothermal behavior of a 10-kg projectile for flight from a high-altitudelaunch into orbit: So far, the results show that an acceptable amount ( 15 mm) of nosetip ablation will occur. This paper provides anoverview of progress in these areas; more details on specific topics are provided in companion papers.

Index Terms—Aerothermal, high velocity, railgun, space.

I. INTRODUCTION

OVER the last half century, thousands of vehicles have

been launched into space using well-established rockettechnology based on liquid fuels and solid propellant boosters.

The advantage of this approach is that the rocket starts slowly

from the surface of the Earth with a full fuel load and builds up

speed gradually as the fuel is burned off. This minimizes aerody-

namic and aerothermal loads while providing relatively modest

accelerations that can be tolerated by humans and delicate pay-

loads. However, this comes at the cost of the need for very large

vehicles with payload ratios of only a few percent and launch

costs up to $20 000 per kilogram. With advances in technology

over the last decade, the desire to put many additional satel-

lites into space exists—but the high cost of launching limits the

ability to achieve this.One alternative for putting small (1–10 kg) satellites into

space could be the use of electromagnetic (EM) launch tech-

nology to replace chemical propulsion. EM launch to space

has been an appealing concept since the first demonstration of

hypervelocity launch in the 1960s and 1970s [1]–[4]. It turns

out that the cost of “fuel”—i.e., electricity—to do this job is

remarkably low. For example, 1 kg launched to 8 km/s has a

kinetic energy of 32 MJ. The cost of electrical energy to achieve

this with an assumed electrical system efficiency of only 30%

(as can be achieved now in the laboratory)—that is, an input

energy of 107 MJ—is only about $2.50 for a typical utility

electricity cost of $0.08/kWh. Of course, this ignores the cap-

ital cost of building an EM launcher as well as the operationalcosts, both of which have yet to be determined. Nevertheless,

early estimates are that moderate costs could be achieved when

amortized over a reasonable number of launches [5].

As part of the Multidisciplinary University Research Initia-

tive (MURI) supported by the U.S. Air Force Office of Scien-

tific Research (AFOSR), the Institute for Advanced Technology

Manuscript received September 26, 2008. Current version published January30, 2009. Corresponding author: I. McNab (e-mail: [email protected]).

Color versions of one or more of the figures in this paper are available onlineat http://ieeexplore.ieee.org.

Digital Object Identifier 10.1109/TMAG.2008.2008601

Fig. 1. IAT airborne launch-to-space concept.

(IAT) at The University of Texas at Austin (UT) is working

with researchers at other universities to develop a hypervelocity

EM launcher that could form the basis of a high-altitude launch

system such as that shown in Fig. 1.

In addition to IAT, the MURI team consists of the Center for

Pulsed Power and Power Electronics at Texas Tech University

(TTU), the University of New Orleans (UNO), and the Univer-

sity of Minnesota (UMN). To demonstrate proof of principle,

the IAT is developing an EM launcher capable of accelerating a

small (5–10 g) projectile to 6–7 km/s. TTU is developing a dis-

tributed-energy power-supply configuration that will improve

high-velocity launcher performance and efficiency. UMN and

UNO are evaluating the aerothermal loads that a projectile trav-

eling 7 km/s will encounter upon exiting a high-altitude EM

launcher.

An introduction to this research effort was provided in ear-

lier papers [6]–[11]. This paper provides an overview of recentprogress by the MURI researchers; more detailed discussions

are provided in companion papers.

II. BACKGROUND

When modern railgun research began in the 1970s, it was

believed that railguns in which solid projectiles are driven by

plasma armatures should be able to attain velocities as high as

50 km/s, because similar velocities had been observed when arcs

alone had been studied. However, by the mid-1980s, a velocity

ceiling of 6 km/s had been observed by several researchers in

experiments when solid payloads were used. The explanation is

that this velocity ceiling for plasma-armature-driven payloads

0018-9464/$25.00 © 2009 IEEE



382 IEEE TRANSACTIONS ON MAGNETICS, VOL. 45, NO. 1, JANUARY 2009

Fig. 2. Plasma formation in the railgun bore [12].

is a direct consequence of ablation from the low-cost G-10 bore

insulators that were used in those experiments. Radiation from

the plasma armature (at a level of ) caused an

ablation of the epoxy from these insulators, which caused the

bore to fill with a hot dense neutral gas [12], [13]. This gas does

not affect the performance of the railgun until, at high veloc-

ities and low pressures, the voltage across the railgun breech

increases to the point where conditions for high-voltage break-

down are met. When this occurs, additional plasma armatures,

called restrike or secondary arcs, form behind the main arma-

ture. These secondary armatures are retarded by viscous drag

as they push the ablation products created in the launcher bore.

This drag prevents the restrike arcs from catching up to the main

armature, causing a significant fraction of the applied current to

be diverted into the restrike arc. This process, which prevents

further acceleration of the main payload, is shown in Fig. 2.

The research being conducted at the IAT under this MURI is

focused on preventing ablation from the bore walls so that thevelocity-limiting effect of restrike arcs can be eliminated. The

IAT research philosophy developed for controlling bore ablation

follows from that developed by Stefani et al. [14] and uses a

multifaceted approach that includes the following.

1) Magnetic augmentation is used to reduce power dissipation

in the plasma.

2) High-purity alumina insulators are used to raise the abla-

tion resistance of the bore.

3) A preinjection of the payload is used to prevent ablation of

the bore materials at low velocity.

Magnetic augmentation allows the current transferred

through the plasma armature to be reduced while the magneticfield inside the bore is kept at a high level to maintain the EM

accelerating force on the armature. This reduces the heat flux

radiated to the bore insulators to a value that can be sustained

without insulator or rail ablation. Because plasma armatures

generate a high heat flux, insulator materials that can withstand

this without ablating must be chosen, and alumina

was chosen for this reason. Finally, the heat flux on the bore

components from the plasma armature is increased substan-

tially when the plasma armature is moving slowly at start-up.

For this reason, the projectile must be preinjected into the barrel

at a velocity of 1 km/s before the plasma armature is created

behind it. Each of these three approaches required a separate

subsystem to be designed, constructed, and tested. This wascompleted, and the three subsystems were integrated, leading

Fig. 3. IAT’s modified MCL core. (Left) Solid model. (Right) Assembledcore.

TABLE IOPERATING PARAMETERS

to recent successful commissioning tests, as described in the

following.

A fourth aspect of the approach to reducing restrike arcs is to

use a distributed-energy power supply rather than a traditional

breech-fed system. A distributed-energy power supply reduces

the effective rail length driven by each power supply to a region

near the armature, thereby preventing voltage application to the

breech of the railgun once the armature has accelerated down

the barrel. In a shorter launcher, like that being used presently

at IAT, the need for a distributed energy feed is less critical.

This aspect of the future EM launch-to-space system is there-

fore being developed separately by TTU. For a future full-scale

system that could launch 10-kg microsatellites, the launcher will

be substantially longer—probably 50 m, depending on the ac-

ceptable acceleration levels for the projectile and payload com-

ponents. This will emphasize the need for the distributed-energy

configuration for system efficiency reasons so that the power

system mass can be minimized.Once these launcher and power-supply approaches are vali-

dated and exit velocities of 6–7 km/s are achieved, the projectile

can be expected to encounter very high aerothermal loads when

exiting the railgun muzzle and flying to orbit. Researchers at

UMN and UNO are currently evaluating how to overcome the

effect of aerothermal ablation under these conditions. These re-

search efforts are discussed in the following.

III. THE UNIVERSITY OF TEXAS AT AUSTIN

This section provides an overview of experimental work un-

dertaken at the IAT—more details are provided in a companionpaper at this conference [15]. Due to limited funding, the IAT



MCNAB: PROGRESS ON HYPERVELOCITY RAILGUN RESEARCH FOR LAUNCH TO SPACE 383

TABLE IIINTEGRATED SYSTEM SHOT SUMMARY

Fig. 4. Plasma preinjector chamber and barrel. (Left) Solid model. (Right)Constructed hardware.

was unable to design a new railgun for the MURI studies. Be-

cause of this, an existing launcher already in use at the IAT—the

medium-caliber launcher (MCL)—was chosen for these exper-

iments. It was judged to be best suited to reach the full velocity

goal (7 km/s) with low current and acceleration loads. The core

designed and utilized for the MCL is shown in Fig. 3.

Two of IAT’s three approaches to overcoming bore ablation

were implemented in this modified MCL core—namely, the use

of magnetic augmentation and alumina insulators. These ap-

proaches were implemented by creating a two-turn indepen-

dently augmented railgun. The outer railgun core had a bore

area of 40 40 mm bounded by rails and insulators. Theserails conducted the augmenting current of 800 kA and set up a

large magnetic field inside the inner core railgun. Because there

was no armature to conduct the return current in the augmenting

rails, a crossover was located at the muzzle end of the rails to

transfer current from the forward rail to the return rail.

The inner core structure formed a 17 17 mm bore, inside

which the plasma armature was accelerated. Roughly 160 kA

was conducted in the inner rails and through the plasma arma-

ture. The rail insulators for the inner core were composed of

high-purity (99.5%) alumina because of its high thermal abla-

tion threshold (12 ). The entire inner core was

evacuated to 20–30 torr.Since an unacceptable amount of bore damage would occur if

the projectile were accelerated from rest using a plasma arma-

ture, a plasma-driven preinjection system was designed, con-

structed, and successfully tested. A plasma-driven injector was

chosen to limit the amount of gas injected into the railgun bore

behind the projectile since past experiments have shown that ex-

cessive gas from a light-gas gun preinjector can cause restrike

arcs at the railgun breech [16]. The plasma preinjector (Fig. 4

and [17]) consisted of a polyethylene liner contained within a

steel pressure vessel through which an arc discharge was initi-

ated by discharging a current pulse into an aluminum wire. After

the wire exploded, a plasma arc between the cathode and anode

ablated a controlled amount of polyethylene from the liner andrapidly heated it to a high temperature. The resulting gas, at a

Fig. 5. Current waveforms (07082904).

Fig. 6. Velocity and position versus time (07082904).

pressure of 100–200 MPa, accelerated the projectile at the en-

trance to the barrel.

Measurements confirmed that the electrical conductivity of

the plasma was high enough to adequately conduct the main

plasma current; therefore, three capacitor modules were dis-

charged into the primary rail breech as soon as the projectilearrived.

Commissioning shots that integrated all three of the subsys-

tems just discussed were performed in 2007. Table I summa-

rizes the basic operating parameters of these experiments. The

first tests used a 3.2-m-long gun, while the 7-m gun was being

built. All shots are summarized in Table II.

The current waveforms for the third shot are shown in Fig. 5,

while the position–time data derived from the B-dots are shown

in Fig. 6. Data showed that an average velocity of 5.2 km/s was

achieved between the positions of 3.74 and 5.29 m, although a

decrease in velocity was observed beyond 5.29 m.

Based on the last two B-dots, it is believed that the armature

separated from the projectile due to the decreasing propulsiveforce. The absence of any secondary arc following the projectile




Fig. 7. 3.2-mm-thick steel target plate.

Fig. 8. Distributed power input concept.

in the last few B-dot traces is the most important result of these

experiments since achieving a velocity above 5 km/s without the

formation of any secondary arcs is a major step forward from

results obtained in earlier decades.

Since hypervelocities were achieved in these experiments,

special precautions were taken to safely stop the projectileat railgun exit. A 3.2-mm-thick steel plate was set up 1 m

from the exit of the gun to slow down the projectile, followed

by more substantial plates beyond that in the catch tank. The

punchthrough observed on the front plate in this shot is shown

in Fig. 7 and is essentially identical to that on the two earlier

shots.

IV. TEXAS TECH UNIVERSITY

A future railgun capable of launching useful projectiles into

space from a high-flying aircraft, like that shown in Fig. 1, will

need to be tens of meters long to ensure that the acceleration

forces can be tolerated by the projectile and payload.It is impractical and inefficient to power such a barrel only

from the breech since the combination of resistive losses in the

rails and unused inductive magnetic energy that is dissipated as

resistive heating will result in a low overall launcher efficiency.

By distributing the power input to the railgun in multiple small

power stages along the barrel, as shown conceptually in Fig. 8,

current flow can be localized in a short region near the armature.

This has the dual benefit of increasing the launch efficiency and

reducing the probability of secondary arc formation [18].

This aspect of the MURI research program is being investi-

gated by researchers at TTU, who have built and tested a dis-

tributed-feed free-running arc railgun [19]–[23]. The free-run-

ning arc railgun allows for realistic armature velocities (5–10km/s) with existing capacitive storage. The major modification

Fig. 9. Plasma velocity in a breech-fed free-running arc railgun—10 kA andvarious pressures.

to a previously developed solid-armature railgun to create the

free-running arc railgun involved operation in a low-pressure

environment ( 5–50 torr) and generation of the initial plasma to

be accelerated. Once constructed, the railgun was tested in three

configurations: breech feed, asynchronous distributed feed, and

synchronous distributed feed. Details are provided in [23].

All of these configurations were tested with G-10 insulators,

and restrike arcs were observed in all cases. Alumina insulators

have only been tested in the breech-fed configuration so far, and

restrike was not observed in that case. Fig. 9 shows the effect

of ablation and restrike on the arc velocity for the breech-fed

system using alumina and G-10 as the insulators.

As expected, using the alumina resulted in higher arc veloc-

ities as a result of low ablation and no restrike. Heavy ablation

with the G-10 resulted in a velocity reduction, most notably at

pressures of 5–10 torr, where restrike drastically reduced the arc

velocity. Increasing pressure slowed down the arc for both cases

because more gas was swept up and added to the plasma mass.

Accompanying the two waveforms is a third data set calculated

from an equation which describes the plasma velocity assumingno ablation. Calculations made using this equation correspond

reasonably well to experiments using the low-ablating alumina.

V. UNIVERSITY OF MINNESOTA

In the experiments undertaken at UT, only simple polycar-

bonate slugs were launched to study the fundamental aspects

of plasma armature propulsion. Of course, any practical system

will need to launch real projectiles that have the following ele-

ments.

1) A hypervelocity aeroshell that can traverse the ambient at-

mosphere.

2) A guidance, navigation, and control capability, togetherwith some rocket propulsion capability that will ensure or-

bital insertion and/or terminal rendezvous.

3) A payload capability, probably involving microelectronics,

to accomplish the desired mission. The required payload

is a function of the mission requirements. Nanosatellites

or possibly even critical space station or satellite resupply

components could be included in the payload.

To accomplish these tasks successfully, it will be necessary

for the launcher to provide an environment that is acceptable

for the launch package survival during launch and gun egress.

Primarily, this will require acceptably low axial launch accelera-

tions, but also important are lateral accelerations and balloting,

as well as control of the muzzle blast during egress from thebarrel.




Probably, the most critical issue that the launch package will

face is transit through the ambient atmosphere at very high

velocities immediately after launch. It is therefore necessary

to consider either ablative materials—such as a carbon–carbon

composite thermal protection system (TPS)—or actively cooled

concepts. Under this MURI program, researchers at two col-

laborating universities, UMN and UNO, have used existing andnewly developed codes to model the aerothermodynamics of

ablating TPS nosetips for slender high-beta flight bodies [24],

[25].

The flight profile of a projectile launched from an airborne

EML resembles a reverse re-entry with the velocity being largest

where the air density is highest, in contrast with re-entry flights

where the majority of the velocity is lost in the upper atmos-

phere before entering denser air. A notable difference between

this flight profile and those of planetary re-entries like the space

shuttle’s, which last several minutes, is that the flight time for the

EM-launched projectile through the thinner atmosphere above

the launch point is only a few seconds, which provides optimism

that the projectile can survive.

A. CFD Solver

To evaluate the severity of the aerothermal conditions en-

countered by the projectile, UMN has developed a combination

of a simple trajectory solver coupled with an adaptation of a

2-D axisymmetric computational fluid dynamics (CFD) solver

to simulate the physical environment experienced by a notional

10-kg projectile during the entire flight from airborne launch

into orbit [26]. The CFD solver couples the simulation of the air

flowing around the projectile, surface interactions between the

air and the solid heat shield, and the conduction of heat into the

heat shield subsurface. The trajectory of the launch vehicle wasfound by specifying its initial conditions and integrating in time

until the projectile leaves the atmosphere ( 60 km) or reaches

a specified orbital height.

The main area of concern for the projectile is thermal protec-

tion against the high heating rate on the nose. The large heating

loads require the use of carbon–carbon for the projectile heat

shield, and the work by Keenan [27]–[29] was used as the basis

of the CFD equations for modeling surface reactions and the

thermal response of the heat shield. The CFD approach con-

sisted of a solver for the fluid flowing around the projectile, a

thermal response model for the solid heat shield, and a set of

equations that coupled the two domains together. The couplingequations also accounted for the reaction of the fluid flow with

the solid surface and calculated the ablation of the TPS.

The solver for the fluid domain was adapted from a two-tem-

perature finite rate chemistry solver developed by Nompelis [30]

which included the following species: , , NO, , ,

, CO, CN, N, O, and C. Twenty-four reactions were used for

the chemical kinetics model, as suggested by Keenan, and chem-

ical equilibrium rates were found by curve-fitting data tables

[31] over the range of temperatures expected in the flow field

(up to 20 000 K). The surface chemistry of the carbon–carbon

involved three phenomena: surface catalysis of oxygen, abla-

tion of the heat shield due to oxidation, and ablation due to the

sublimation of carbon—all of which act to erode the carbon ab-lator surface. By summing the mass production terms over all

Fig. 10. Fuel mass fraction required for injection rocket versus pre-injectionprojectile velocity angle. Each line represents the velocity of projectile beforeinjection.

species, the surface recession was obtained. Few experiments

have been conducted at flight conditions similar to those beingstudied here. The most appropriate were those conducted under

the passive nosetip technology (PANT) program [32], which

were undertaken at a velocity of 5.48 km/s. A comparison shows

that the UMN code matches the PANT data fairly well (within

50%) for the stagnation temperature and recession rate at the

pressures of interest ( 125 atm), thereby validating the present

approach.

B. Trajectory Estimation

In this study, it was assumed that the desired final orbit is

a circular low-Earth orbit. The circular shape requires that the

projectile velocity vector has an angle of 0 with respectto the horizon as it reaches the orbit altitude. Any other angle

will require a rocket to provide an appropriately directed extra

velocity increment to inject the projectile into its final

orbit. Depending on the angle of the projectile velocity vector

, this additional correction can be quite large. Assuming

that a solid rocket is used for injection with an of 250 s, Fig.

10 shows that, unless this injection angle is small, a large mass

fraction of rocket fuel will be required for orbit insertion. This

indicates that the flight dynamics of the projectile will be very

important: A large mass penalty will be paid if the projectile

cannot fly in a trajectory that will give it a small value for

before orbital injection.The purpose of this part of the study was to find trajectories

that reduced thermal loading on the projectile heat shield while

minimizing rocket fuel mass due to the requirement for large

injection velocities. The angle of attack of the projectile was

varied from 0 to 5 to create a lift to turn the velocity vector of

the projectile inline with the smallest possible before injection

to orbit. The projectile geometries used are listed in Table III.

Trajectories for these geometries were found for launches from

15 km.

These trajectory estimations showedthat some type of turning

maneuver was necessary for the launch to orbit to minimize

rocket fuel mass. This turning maneuver cannot be simulated

in the CFD solver due to its axisymmetric nature. As a compro-mise, initial launch angles of 20 and 45 were studied, and it




TABLE IIIPROJECTILE GEOMETRY

TABLE IVDESIGN STUDY PARAMETERS

Fig. 11. Stagnation point recession over trajectory for 7-km/s case.

was assumed that the projectile would follow a ballistic trajec-

tory out of the thicker atmosphere and, after a certain height,

perform a turning maneuver where a lift would make the injec-

tion angle small. The main focus of this study was to understand

the aerothermal issues of the airborne EM launch concept, not

to find exact design solutions, which will follow later.

C. CFD Trajectory Simulations

For this section, only one of the studies undertaken so far

is presented here to illustrate the results: More details can be

found in [26]. This example results for a 45 angle launch from

a 15-km altitude. The geometry and launch conditions werevaried, as shown in Table IV for launch velocities of 7 and 9

km/s and several TPS thicknesses. Only the details of the 7-km/s

launch are shown here.

The maximum stagnation point surface recession during

flight along the trajectory is given in Fig. 11 while Fig. 12

shows the final surface profile. The most important result is that

none of the cases evaluated showed the TPS failing due to burn

through—that is, the total calculated recession was always less

than the thickness of the TPS.

The next most important performance factor is therefore the

back wall temperature, which must be low enough to ensure that

the payload and other important subsystems do not overheat.

Fig. 13 shows the back wall temperature of the TPS at the stag-nation line. For the 7-km/s case, the 3-cm-thick TPS has a back

Fig.12. Final surface location for 7-km/s launch with 2-cm nose radius geom-etry.

Fig. 13. Stagnation back wall temperature over trajectory for 7-km/s cases.

Fig. 14. Projectile velocity over trajectory for 7-km/s cases.

wall temperature of 1000 K, which is probably still too high.

(Interestingly, the 9-km/s case (not shown here) has a similar

temperature contour and back wall temperature.)

Fig. 14 shows that the velocity loss through the trajectory

for the 2-cm nose radius case is 0.57 km/s and, for the 3-cm

nose radius case, is 0.86 km/s for an initial velocity of 7 km/s,

corresponding to a required rocket motor fuel mass fraction

of 42% and 49%, respectively. Since the aerothermal penalty

for launching at 9 versus 7 km/s is small compared to the re-

quired fuel penalty, the larger launch velocity is more attractive

from this point of view—although this is more stressing for the

launcher. Further studies will qualify the ability of launches atshallower angles and lower launch velocities.




Fig. 15. Final ablated profiles for r = 2 c m , V = 5 , 6, 7, 8, and 9 km/s,laminar flow, l a u n c h a l t i t u d e = 1 6 k m , and l a u n c h a n g l e = 4 5 .

VI. UNIVERSITY OF NEW ORLEANS

The primary goal of the UNO effort was to adapt the ABRES

Shape Change Code (ASCC) to the conditions of interest for

airborne EM launch and demonstrate that ASCC can provide

recession histories to guide preliminary designs of a TPS. The

designs can then be validated and further refined with the codes

currently under development by UMN. In the studies undertaken

so far, launch trajectories and sphere–cone geometric param-

eters were chosen to match the UMN work so that a quanti-

tative comparison of recession histories predicted by two ap-

proaches could be made. As a first step, the UNO effort focused

on adapting and applying ASCC to compute projectile trajecto-

ries and total ablation for relevant geometry configurations and

launch conditions.

The ASCC computations served two purposes. First, ASCC

is a well-documented code that has historically agreed well with

flight data for sphere–cone geometries and, therefore, can pro-vide validation data for the code development work under way at

UMN. Second, ASCC typically takes less than 5 min to run on a

desktop or laptop computer; therefore, parameter studies can be

efficiently performed to screen potential projectile designs and

then the preliminary designs refined with the more sophisticated

methods under development at UMN.

The preliminary studies adopted a sphere cone geometry of

length , cone angle , launch mass ,

and nose radii , 2, and 3 cm. The nosetip material was

graphite with density . For all the computa-

tional results presented in this paper, the launch angle was held

constant at . For each nose radius, trajectories werecomputed for 5, 6, 7, 8, and 9 km/s. The trajectory computa-

tions were terminated at 60 km, where previous computations

indicated that the surface recession had abated and the nose tip

had begun to cool. The initial temperature of the projectile was

assumed to be 300 K. As an example, the final ablated profile

computed for is shown in Fig. 15 for a launch al-

titude of 16 km assuming laminar flow. This can be compared

with the UMN calculations (Fig. 12). The lateral recession on

the conical section of the projectile was small.

The over the trajectories were 423, 891, and 1375 km/s

from 16 to 60 km for the -, 2-, and 3-cm nose radii,

respectively. Clearly, that additional velocity decrement for

the larger nose radii would require additional propellant mass;therefore, it appears that smaller nose radii are preferable as

long as the volume remains large enough to hold the payload

mass and stagnation temperatures are acceptable.

A preliminary comparison with the calculations by UMN

indicated that, in all cases considered, ABRES overpredicted

the total stagnation-point recession, compared to the results of

Gosse [26]. Some of the differences are attributable to slightly

different assumptions, but, even after resolving these, therewere still differences in the range of 20%–50%. The resolution

of these remaining differences will be one of the objectives of

future studies.

VII. SUMMARY

The case for launching a projectile to space using an EM

launcher in a high-altitude aircraft has been shown to be plau-

sible from a thermodynamics view. It should be possible to se-

lect a heat shield configuration that will successfully protect the

payload from the heating experienced during exit from the at-

mosphere and not consume a large fraction of the projectile’s

mass budget. In parallel, experiments with the UT approach toachieving hypervelocities appears promising, with velocities of

5.2 km/s achieved, while TTU has confirmed the benefits of

alumina ceramics as an insulator choice with distributed power

feed experiments using free-running arcs at over 11 km/s.

Clearly, further work is needed in all areas to ensure future

success. It is hoped that this work will be conducted under

MURI funding over the next two years.

ACKNOWLEDGMENT

This work was supported by the United States AFOSR under

MURI Award FA9550-05-1-0341, under the direction of Dr. M.Birkan. The authors would like to thank all the members of the

MURI team, and the coinvestigators and team members for their

contributions, particularly including Dr. J. Parker (SAIC and

consultant to IAT), F. Stefani, Dr. D. Wetz, and D. Motes (IAT),

Prof. J. Mankowski, R. Karhi (TTU), Prof. G. Candler and Dr.

R. Gosse (UMN), and Prof. M. Guillot (UNO), without whom

this paper would not have been possible.

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progress on hyper velocity railgun research for launch to space

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