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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
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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
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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
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384 IEEE TRANSACTIONS ON MAGNETICS, VOL. 45, NO. 1, JANUARY 2009
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.
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MCNAB: PROGRESS ON HYPERVELOCITY RAILGUN RESEARCH FOR LAUNCH TO SPACE 385
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
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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.
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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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