A Whitepaper on

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A Whitepaper on

Responsive and Affordable Nano
-
Satellite Launcher

NanoLaunch, LLC

November 30,
2010


Abstract

Despite more than 45 years of experience, launching payloads into earth orbit is still an expensive and
ris
ky enterprise. Achieving Low Earth Orbit (LEO) with medium to large launchers costs in the range of
$10,000
-
$20,000 per kilogram. The situation for the small satellites is even worst. The lowest price for
any single orbital launch mission is around $7.5 mi
llion, resulting in costs exceeding $100,000 per
kilogram for small satellites. The high costs and the poor responsiveness of the existing orbital launch
vehicles have hindered the emergence of many business opportunities and prevented the implementation
o
f a number of DoD or NASA missions that rely on small cost effective satellites.


Recent advances in electronics allow for the smaller and inexpensive satellites (in the micro or nano class)
to perform most of the functionality of the larger systems. The d
ramatic changes in the electronics
combined with the advent of small affordable ground stations, and the wide application of the Internet
means that constellations of small LEO satellites have the potential to be a transformative technology. At
the very le
ast, there is a growing potential for niche applications in telecommunications, sensor data
collection, on
-
orbit servicing, on
-
orbit re
-
supply, and navigation markets. These factors combine to
create a latent need for launches of small military, NASA and
commercial payloads. Despite the emerging
opportunities, there are currently no vehicles dedicated to launch nano or micro satellites into LEO.
Instead small satellite producers and operators, including NASA, have relied on piggy
-
back opportunities
for whi
ch the small satellite is carried as a secondary payload in a large vehicle such as Minotaur, Shuttle
or the Russian system Dnepr. Piggy
-
backing approach severely limits the size, schedule, orbit choice and
the equipment that can be included in the small s
atellite. For example active propulsion systems are
typically not incorporated in the secondary payload in order to limit the potential risks to the primary
mission. These limitations significantly hinder the capability of a small satellite launched as a p
iggy
-
back,
compromising its overall utility and economic viability.


In response, NanoLaunch, LLC proposes to develop a devoted, responsive and cost effective
vehicle to launch Nano
-
satellites (1
-
10 kg) and Micro
-
satellites (10
-
100 kg) into

Low Earth Orb
it (LEO) using the American
high performance

F
-
15 Eagle as the air platform.
NanoLaunch enterprise, a joint venture of four small entrepreneurial companies, is in a unique
position to achieve this challenging objective, since the venture is formed specific
ally to combine
the enabling elements of 1) a new high performance hybrid rocket technology, 2) a broad
operational experience with the jet aircraft and 3) an experienced team involved in earlier launch
vehicle development programs.

Key Elements of the Nan
oLaunch Approach:

NanoLaunch’s approach for the small orbital launcher development has the following key and
mostly unique elements which are believed to be essential in the success of the proposed
program.

1) Propulsion System:

It is widely accepted that
the propulsion system is the most critical component of an orbital
launch vehicle. It typically constitutes 90% of the vehicle’s mass and propulsion system related
issues often dominate all other failure modes of a launch system, contributing to its reliab
ility
significantly. One of the key aspects that differentiates NanoLaunch’s orbital launch concept
from the previous attempts is the use of a new high performance hybrid rocket technology which
is currently being developed by SPG Inc. (a partner in the Na
noLaunch venture) under an Air
Force contract.

Paraffin
-
based/LOX Hybrid System:

The classical hybrid rocket systems developed to date have suffered from two major
shortcomings: 1) complex multiport fuel grains as a result of the poor regression rate
perfo
rmance of the classical polymeric fuels and 2) low frequency instabilities. In the past, the
mitigation methods for these problem areas have introduced significant complexity to the motor
design, compromising the simplicity advantage of hybrids. For exampl
e, the 250 klb motor
developed by American Rocket Company (AMROC) was based on a complex 15 port wagon
wheel configuration (resulting in poor fuel utilization and expensive fabrication) and the motor
stability was achieved by the continuous injection of a
hazardous pyrophoric substance,
triethylaluminum (TEA). SPG’s paraffin
-
based/LOX hybrid rocket technology, which has an
inherently high fuel regression rate, allows for the use of a simple single circular port fuel grain
design approach. SPG has also devel
oped a unique proprietary technology to eliminate the low
frequency instabilities in LOX
-
based hybrids without resorting to external heat or pyrophoric
liquid addition at the fore end of the motor. These two technological advancements are crucial in
keepin
g the hybrid concept cost effective, simple and safe compared to the state of the art liquid
and solid rocket systems. The key virtues of the paraffin
-
based/LOX motor technology and its
impact on the NanoLaunch’s launch vehicle design are summarized in Tab
le 1.


Table 1:
Enabling propulsion system technologies for the NanoLaunch concept.

Virtue

Enabling Key Technology

Impact

Single Circular Port Design

Paraffin
-
based fuels

Simple, inexpensive

grain

High fuel utilization

Adjustable fuel regression rate

Fuel formulation tailored to
mission

Mission flexibility

Stable and efficient
combustion with no external
heat or TEA addition

SPG proprietary
injector/fore end design

High performance without
compromising systems
simplicity

High Isp and impulse density

Paraffin
-
based fuels and
LOX

Two stage ERV to orbit is
feasible

Low cost, readily available
propellants

Paraffin
-
based fuels and
LOX

Reduced development and
recurring costs

Simple motor design with no
exotic materials

Advanced internal ballistic
design

and testing

Reduced costs and high
reliability

Low environmental impact

Paraffin
-
based fuels and
LOX

Simplified operations and
reduced costs

Safety

Zero TNT equivalency

Low fire hazard

Ideal for aircraft operations
Reduced development costs

Throttling

SPG proprietary throttling
valve developed and tested

Mission flexibility

Efficient gas phase
combustion

Tailored fuel grain
technology

Effective use of the
pressurant as propellant

Thrust vector control (TVC)

Liquid Injection Thrust
Vector Control*

Sim
ple TVC capability

* Whittinghill Aerospace technology

C
omparison to Other Chemical Systems:

Most of the previous small air launch concepts have utilized solid rocket propulsion for all stages
of the ERV. The use of paraffin
-
based/LOX hybrids is believed
to have significant advantages
over the solid rocket systems. The important advantages can be listed as:



Cost savings in motor development, manufacturing and launch operations. This cost
saving stems from the inherent safety of the hybrid system and its fa
brication
simplicity which can be done in typical industrial areas. Shipping is simple because of
the non
-
hazmat classification of a paraffin
-
based motor. Moreover the ingredients of
the paraffin
-
based/LOX system propellants are much less expensive than th
e
ingredients of a typical solid rocket propellant.



Inherent safety in manufacturing, transportation, storage, testing and launch operation
phases. Zero TNT equivalency of the hybrid is especially critical during the launch
operation which is carried out
with a piloted aircraft. Elimination of the storage
hazard is also critical in allowing operations in a wide range of airports with minimal
infrastructure investment. The other important advantage of the zero TNT
equivalency and motor shutdown capability i
s the simplification of the range safety
requirements which is known to be a major cost driver for launch systems.



Higher fault tolerance. Hybrids are not sensitive to cracks or debonding, eliminating
the requirement for expensive quality control operatio
ns such as x
-
ray examination of
the motor.



Environmental friendliness. Paraffin, its additives and LOX are all green and non
-
hazardous materials. Most high performance solid rockets use ammonium perchlorate
for oxidizer, a substance which presents hazard t
o the environment and shown to have
adverse effects on human health.



Paraffin
-
based/LOX hybrid has a significant delivered Isp advantage (~35 seconds)
over the solid rocket.



Hybrid can be throttled or can be shut down. This virtue introduces mission flexi
bility
and improves the insertion accuracy.


The primary disadvantage of the hybrid system compared to the solid rocket is its lower impulse
density which results in a volume limited design approach for applications with strict
geometrical envelope require
ments.

For air launched orbital systems, paraffin
-
based/LOX hybrids have also significant advantages
over liquid rockets.




Mechanical simplicity. Hybrids have significantly simpler liquid storage/feed systems
and injectors (one liquid as opposed to two).

Moreover no active cooling of the
hybrid chamber is necessary, since it is protected by the fuel grain.



Structural mass fraction of the paraffin
-
based/LOX system is expected to be better
than a pressure fed RP
-
1/LOX liquid rocket. Note that pump fed liqu
ids are not
expected to be cost effective for small launch vehicles.



Hybrids are fault tolerant compared to liquids. The tolerance requirements on the
machined parts can be much more relaxed in hybrids.



Hybrids have reduced fire hazard compared to liquids
.


Paraffin
-
based/LOX Hybrid Propulsion System Development:

Since July 2005, SPG has been on a SBIR Ph III contract from the Air Force Propulsion Lab
tasked to develop and test the paraffin
-
based/LOX hybrid systems. The following is a summary
of the impor
tant progress made in this ongoing program:



Formulation and characterization of paraffin
-
based fuels with a wide range of ballistic
properties. This effort included the development of mechanical and regression rate
testing techniques for candidate fuel for
mulations. The ability to tailor the fuel
formulation for a given set of mission requirements has also been achieved.



Fuel processing technologies has been developed. SPG has now capability to cast
grains up to 36 inches in diameter. Two 22 inch OD paraffi
n fuel grains (700 kg each)
have been successfully fabricated using the existing facilities (see Figure 1).


Figure 1:

22 inch paraffin
-
based fuel and casting facility.




A total of 40 motor tests have been conducted using the 11 inch LOX motor, in the
7
,000 lb thrust class (see Figure 2).



A unique proprietary technology has been developed to eliminate the low frequency
instabilities without resorting to heat addition or pyrophoric liquid injection at the
fore end.



Methods have been developed and tested
to mitigate the acoustic instabilities in
hybrid systems.



95% c* efficiency has been demonstrated in the 11 inch LOX motors.



Figure 2:

11 inch Paraffin
-
based/LOX motor testing in Butte, Montana.


The following progress is expected in CY 201
1
.




The 10 in
ch motor will be converted to a carbon composite case flight weight system.
The design of the flight weight system has been completed and the testing shall start
in June 201
1
. Note that a cluster of 10 inch flight weight motors shall be used in the
second
stage of the NanoLaunch’s ERV.



Testing of the 22 inch LOX motor (35,000 lb thrust class)
commenced in

2010. The
test stand for these firings
has been constructed
near Butte, Montana
.

Two 22 inch
motors configured parallel will be used in the first stage of

the ERV system.



SPG has recently acquired a filament winding system that can be used to produce
motors as large as 36 inches in diameter.


2)
Atmospheric
Launch

In the NanoLaunch concept, an air launch approach has been adapted due to its advantages over

the conventional ground launch:



The ambient air density during the flight of an air launched ERV is significantly reduced
compared to a ground launched vehicle. For example the air density at 35,000 ft is
approximately 30 percent of the density at sea lev
el. Launching the rocket propelled ERV at
a higher altitude results in lower dynamic pressure loads. This allows the vehicle to fly a
flatter, more efficient trajectory. The depressed trajectory into the target orbit results in a
decrease in gravity losses
.




Low atmospheric pressure allows for higher nozzle area ratios and, consequently, better
performance for a given combustion chamber pressure.




The delta V required from the ERV is reduced since the vehicle starts its flight with more
energy due to its al
titude and nonzero velocity. This reduction in mission delta V can be as
high as 10 percent. This decrease in required delta V translates into a significant reduction in
the launch vehicle gross mass (for a given payload mass) due to the exponential nature

of the
rocket equation.




Using a mobile launch platform increases flexibility for the launch angle and launch latitude
hence increasing the range of achievable orbital inclinations.




Air launch simplifies range safety and abort modes. Launch over the oce
an allows hazardous
operations to occur well separated from population centers.


F
-
15 Feasibility

as Multi
-
Stage, Orbital N
-
MSLV Atmospheric Launch Platform



In order to keep the development costs low and to reduce the programmatic risks,
an existing
hig
h performance aircraft from among those now operation
al
, will be used.
The American
high
performance
F
-
15 Eagle
is the platform of choice, due to its suite of characteristics:




High payload mass capability and large
undercarriage dimensions
.



Ability to
launch the ERV at high altitudes, high Mach numbers and large trajectory
angles.



Indigenous
US built aircraft. A large inventory of F
-
15’s exists in the US.



Current operational status

available spare parts, maintenance flight
expertise.



Safety: Twin engi
ne safety. Fourth generation technology. Modern safety protocols.



Proven ability: the 1985 ASAT was launched from an F
-
15.



The baseline ERV does not require major modifications to the existing aircraft







Figure 3.
F
-
15 in two views




The F
-
15, a fourth generation long
-
range air superiority fighter, was introduced in 1976 with
approximately 410 F
-
15A/B model aircraft produced. In 1979 U.S. production proceeded to
more advanced models, with foreign production cont
inuing until the present time. A total of
1,286 F
-
15s were produced for the United States. They still operate within the USAF and ANG
today

as well as in selected foreign nations
.


The F
-
15 A/B Eagle, with F100
-
PW
-
220 Engines and without conformal or ext
ernal fuel tanks
has the following flight characteristics and limitations

A.

Length: 19.43 m (63.75 ft)

B.

Wingspan: 13.05 m (42.81 ft)

C.

Height: 5.63 m (18.46 ft)

D.

Maximum Mach: 2.2

E.

Maximum Altitude: 18,181 m (60,000 ft) (Cabin Pressure Control System Limi
tation)

F.

Basic Operating Weight: 12,387 kg (27,500 lb)

G.

Full Internal Fuel Weight: 5,338 kg (11,850 lb).

H.

Aircraft Gross Weight = 17,725 kg (39, 350 lb) with full internal fuel

I.

Maximum Allowable Gross Weight: 25,225 kg (56,000 lb)

J.

Maximum Payload: 7,2
60 kg (16,000 lb)

K.

Forward Center of Gravity Limit: 22% MAC (gear down)

L.

Aft Center of Gravity Limit: 29.9% MAC

M.

Centerline Store Acceleration Force Factor Limit (Store Weight x Maximum G Allowed) [Example uses
weight of full centerline fuel tank multiplied

by G limit]:

4581 lbs. x 6.5 G = 29, 776 lbs (13, 412 Kg).

N.

Centerline Store Maximum Dimensions:

a.

Undercarriage width limit:
1.45 m (4 ft, 10 in.)

b.

Undercarriage vertical clearance:

1.35m (
4.5 f
t)

c.

Undercarriage length limit: 8.1 m (27.0 ft)

{Nose

gear door to tail hook
housing
}




A study by Air Force Research Laboratory, Kirtland AFB, New Mexico in 2003 analyzed the F
-
15 capability in carrying a 4550 kg (10,100 lb) vehicle with a 100 kg (222 lb) payloa
d to a target
orbit of 225 km. [9] This centerline load was found to meet the center of gravity requirements of
the aircraft. [9] The pylon used was a modified SUU
-
60. [13] The study concluded that both the
aircraft
-
to
-
pylon and the pylon
-
to
-
vehicle interf
ace were safe to a factor of 2.0 for these in
-
flight
maneuvers: Symmetric G loads from
-
1G to 5G; half
-
stick rolls from 1G to 3G; and half
-
rudder
pedal sideslips. [9] A Boeing Company simulation predicted directional stability exceeding
standard loads up
to Mach 1.6 and acceptable directional stability above Mach 1.6 for the MSLV
identified in the study. A safe release to light
-
off time of 2 seconds was identified in the study.
[13]


F
-
15 Simulator tests in 2003 proved its capability to a payload of 93 kg

(206 lb.)
, using launch
parameters of FL 380, Mach 1.7, and a 60 degree flight path angle. That payload weight exceeds
the requirement for a NanoLaunch
1 atmospheric
satellite launch. [9]


The preliminary design for the NanoLaunch
1

N
-
MSLV is

A.

7.62 m (
25.0

ft
)
in length

B.

3,378 kg (
7,500 lbs
)

in weight

C.

Vertical Diameter
1.02 m (
3.33 ft
)


D.

Positive Acceleration Limit of 3.9 G. on F
-
15 centerline station

E.

Estimated Drag Index with orbital rocket mounted centerline station: 20.0


A rocket of those metrics can
be carried on the centerline station of an F
-
15 aircraft and be
within aircraft gross weight limits, centerline store weight limits, and external store drag limits.


Refined parameters will be calculated for each orbital launch but will fall within these
parameters of which the F
-
15 is capable:


Mach: 0.8


1.40


Pitch: 0
-
8
0 degrees


Altitude: 20,000


50,000 feet


It must be said that NanoLaunch does not currently possess an F
-
15. One must be acquired via
federal legislation or purchase from a non
-
U.S
. owner. Acquisition via a Government Furnished
Equipment (GFE) agreement during the R&D phase is possible. Possession of three or more
would constitute a preferred operational nucleus. Other internationally available aircraft with
adequate characterist
ics are available, should no F
-
15 be acquired to meet the need.


3)
Experience in High Performance Aircraft

Premier Space Systems (PSS) will manage the atmospheric launch portion of NanoLaunch. The
foundation of this contribution is based on more than 25
years of international commercial and
high performance military jet aircraft operations. The PSS complement of pilots, maintenance
technicians, and engineers includes experience in R,D,T & E of both civil and high performance
tactical military aircraft

th
rough their fourth generation
. This experience includes the
integration, qualification, and employment of external stores.


The
Reusable

Launch
Vehicles

(RLVs)
…the high performance aircraft…
are extant technology,
which require minimal modifications for su
ccessful employment in the atmospheric launch
mission. A “Captive on Bottom” [12] carry system allows the utilization of existing and cost
-
effective external stores mounting and release mechanisms. Required but simple modifications to
the pylon and release

mechanism have been designed previously. [13]

The NanoLaunch operational concept involves the following elements:

A.

2
-
Stage, F
-
15 Launched, Nano
-

and Microsatellite Launch Vehicle (N
-
MSLV)

B.

Operating base near over
-
water launch range

C.

Low population route

to over ocean launch point

D.

Sub
-

or supersonic air launch at pre
-
determined altitude, speed, pitch, and orbital
inclination. Launch vehicle recovery and recycling.

E.

Dedicated, Responsive, Mobile, Cost
-
Effective
, Safe

Launch


NanoLaunch

possesses the high p
erformance aircraft necessary for initial testing of a single
stage, single motor, hybrid rocket system produced by SPG, NanoLaunch1. This nano
-
satellite
launch vehicle (NSLV) is intended for sub
-
orbital flight only. Development of a NSLV capable
of sub
-
orbital satellite launch is a useful step en route to the ultimate goal of establishing a low
earth orbit MSLV. Achieving operational status in sub
-
orbital launch validates many of the
technical concepts required to proceed to the orbital operational capa
bility.

It also establishes a
salable space R&D system for sub
-
orbital space researchers.


MiG
-
21 Feasibility as
a
Single
-
Stage, Sub
-
Orbital, Atmospheric Launch Platform


Early testing of individual components of the NanoLaunch1 Rocket will be flown using

a MiG
-
21 as the atmospheric test launch platform.
Premier Space Systems possesses three MiG
-
21
aircraft variants. These aircraft are adequate to the task of initial single
-
stage, single motor (10
inch class/150,000+ lb.
-
sec class) NSLV testing and, even
tually, for sub
-
orbital atmospheric
launches.


Romanian computer simulations by G. Savu in 2005 [4]

demonstrated

the technical feasibility of
launching an 800 kg
(1,764 lb.)

rocket with a 10 kg
(22 lb.)

satellite as payload via MiG
-
21 and
centerline moun
ted single stage rocket into 116 km low earth orbit. The rocket size, weight,
aerodynamics, and thrust metrics in this study were less favorable than the parameters in the
preliminary design for the NanoLaunch1 sub
-
orbital rocket. Thus, the NanoLaunch obj
ective of a
sub
-
orbital capability with a more optimally designed NSLV and from the same launch vehicle
as used in the Romanian study, is feasible.


The Mikoyan
-
Gurevich MiG
-
21 is a Russian
-
designed aircraft, originally dedicated to short
-
range point air d
efense. It was introduced in 1959 and continued in production until 1985. Over
11,000 were built, making it the most
-
produced supersonic jet aircraft in aviation history. Later
versions are considered third generation fighters.






Figure 4.

Airborne MiG
-
21 with Centerline Fuel Tank


The MiG
-
21UM has the following characteristics



Length


15.8 m (51 ft 8 in)



Wingspan


7.1 m (23ft 5in)



Height


4.1 m (13ft 5in)



Engine

Tumansky R11F2S
-
300



Maximum

Air Speed


540 KIAS (1200 Km/Hr, above 5000 m)



Maximum Mach Number


2.05 Clean, 1.6 with external fuel tank(s) or sub
-
orbital rocket



G
-
Loading
--

+6.0 with external fuel tank(s)



Maximum Altitude


50,000 feet



Maximum Takeoff

Weight


9,500 kg (20,94
4 lb)



Maximum Landing Weight


7,300 kg (16,094 lb)



Fuel Capacity

o

Internal


2600 liters (645 US Gal.)

o

Centerline External Fuel Tank


490 liters (130 US Gal.) {395 kg, or 870 lb when full}



Centerline Ejector Rack: VYROBCE JE6, Modified



Centerline Lug sep
aration:
64 cm (
29.3 inches
)



Centerline Store Maximum Dimensions:

o

Undercarriage Length:
5.1 meters (
17 feet
)

o

Undercarriage Width:
6.7 cm (
17 inches
)

o

Undercarriage Height:
12.2 cm (
31 inches
)


Test Flight R,D,T&E will proceed from captive carry to un
-
p
owered ballistic drop, to powered
launches. The operational concepts for sub
-
orbital atmospheric tests and launches will lay the
groundwork for the larger and more complex launches to follow:


The external dimensions of th
e single
-
stage, sub
-
orbital

Nan
oLaunch1
rocket
are

4.3 cm (
11
inches
)

in diameter,
4.8 m (
16 f
t)

in length, and with a weight of 338 kg (750 lbs)
.

Refined
launch parameters for sub
-
orbital launches will be calculated for each launch but will fall within
these parameters of which the
MiG
-
21 is capable:


Mach: 0.
8



1.3


Pitch: 0
-

60 degrees


Altitude: 7575


15,000 m (25,000


50,000 ft)


Initial flight testing using a mock
-
up of a single motor (10 inch core) flight weight rocket will
commence in the s
pring of 2011.

These will i
nclude flights with a full
-
scale rocket mock
-
up
(dimensions, shape, aerodynamic profile, center of gravity equal to the flight weight
NanoLaunch1 rocket). This Phase 1 flight test program will consist of graduated release
parameters, culminating in a rock
et mock
-
up release at the anticipated altitude, Mach number,
and pitch angle of a sub
-
orbital launch. These validation flight tests will be conducted with the
test rocket carried beneath a MiG
-
21UM aircraft. This baseline data will undergird the first su
b
-
orbital power rocket launches in
late summer
2011.


Phase II development in 2011 will involve power testing of a flight weight NSLV with graduated
milestones leading to the declaration of sub
-
orbital launch capability for satellites of less than 45
kg.




4)
Systems Integration and Testing
:

Spath Engineering. Terry Spath

graduated from the University of Washington in Aeronautics
and Astronautics in 1974, and has participated in numerous aerospace development and test
programs.


Notable past projects i
nclude aircraft flight tests at Learjet, Inc. and liquid bi
-
propellant rocket development work as a contractor to the Air Force Research Laboratory. Spath
has participated in all hot fire tests of the SPG Paraffin
-
based/LOX hybrid rockets. Spath
Engineerin
g contracted with the Defense Advanced Research Projects Agency (DARPA) for
development of supersonic jet engines. Terry Spath holds an FAA ATP and A&P with
Inspection Authorization.

Portland State Aerospace Society is affiliated with Portland State Univer
sity of Portland, Oregon.
It is an association of academic and professional engineers dedicated to rocketry and space flight.
Their focus and expertise is electronic miniaturization, telemetry, downlink, photographic
documentation, and flight control… F
or the NanoLaunch project this organization will manage
the rocket systems integration.

NanoLaunch I Vehicle

The preliminary design of a two stage orbital ERV (NanoLaunch I) has been completed. The
general characteristics of the NanoLaunch I system can be

summarized as

1)

Air launched from an F
-
15

2)

A two stage Expendable Rocket Vehicle (ERV) dropped from the airplane

3)

Paraffin
-
based/LOX hybrid rockets on both ERV stages

4)

First stage 22” diameter hybrid rocket currently under development

5)

Second stage 10” diamet
er hybrid rockets currently under development

6)

Re
-
startable second stage

7)

Upper stage steering differentially throttled by three parallel motors

8)

First stage uses LITVC


The design and optimization of the launch vehicle has been conducted using an in
-
house
computer code. The predicted payload mass delivered to a 400 kilometer polar orbit is 30

kilograms excluding the fairing and payload adapter ring.





Figure 5
:
F
-
15 ERV Integrated Vehicle


The leading particulars
of the proposed air launched ERV are as summarized in Table 2.


Table 2:

Properties of the NanoLaunch I launch vehicle.

Gross weight of ERV stacked with payload hanging on F
-
15

3,008 kg.

Payload to 400 kilometer polar orbit

30 kg.

F
-
15 Launch altitude

35
,000 Feet

F
-
15 Launch angle (flight path angle)

30
0

F
-
15 Launch velocity (standard atmosphere)

485 Knots (true airspeed)


嘠灲潶楤敤⁢y‱
st

and 2
nd

stage


a摤楴楶攠瑯⁡楲i牡晴

㠬㜳㈠浥瑥牳 ⁳rc潮o

䵡x業畭⁅o嘠scce汥牡t楯渠摵物湧⁡sce湴

9.7 g’s

䕒嘠b
st

stage mass

2,771 kg.

ERV 2
nd

stage mass (excluding payload)

207 kg

Mass contingency carried during performance calculations

15%


Summary

NanoLaunch
has been formed by Premier
Space Systems,
Space Propulsion Group (SPG),
Portland State Aerospace
Society
, and Spath Engineering. NanoLaunch proposes a four year
development program which is structured to leverage the diverse and applicable experience of
the NanoLaunch component companies for achieving operational orbital capability. For example,
SPG i
s one of the world’s leaders in hybrid rocket propulsion field and has developed a unique
high performance hybrid rocket technology. Premier
Space Systems
has wide experience in high
performance jet operation; including the F
-
15. Spath
has
deep
propulsion
engineering expertise
.
PSAS have launched rockets and occupy a leading edge of design and manufacture of miniature
electronic components.


NanoLaunch

proposes to leverage the engineering advances in electronic miniaturization with
its

own propulsion and
launch system to create a
devoted,
responsive, cost
-
effective method of
launching nano
-

and micro
-

satellites, potentially to a weight of 90 kg (200 lb.). It proposes to
create the capability to do so at a fraction of
the
current client cost to do so.

R
ocket motor tests
have occurred for the past five years. The first flight tests will occur in 2011.


END

3D Graphic of F15 with
rocket will be inserted here


R
eferences
:


MiG
-
21


1.

Mikoyan
-
Gurevich OKB (Aircraft Design Bureau) MiG
-
21UM Pilots Flight Operating Instructions 1975
(English Translation by MiG M
asters)


2.

Gordon, Y.,
Mikoyan MiG
-
21 (Famous Russian Aircraft)
,

Hinckley: Midland, 2007. ISBN 1857802578


3.

Davies, S., “
Red Eagles: America’s Secret MiGs.


Osprey Publishing 2008. ISBN 1846033780


4.

G. Savu. “Micro, nano, and pico satellites launched
from the Romanian

Territory,” 2005 Elsevier Ltd.
doi:10.1016/j.actaastro.2005.07.004.


F
-
15


5.

T.O. 1F
-
15A
-
1 (USAF F
-
15 Flight Manual)


6.
Gunston, B
.
, ed.
The Encyclopedia of Modern Warplanes
. NY: Barnes & Noble, 1995, p
p

176
-
177,



McDonnel
l

Douglas F
-
15 A/C Eagle.


7.
Munro, B
.,

and Chant, C
.,

Jane’s Combat Aircraft.

Glasgow: Harper Collins Publishers, 1995,



p
p
.
1
54
-
155, McDonnel
l

Douglas F
-
15C Eagle.


8.
AF.mil Fact Sheet: F
-
15 Eagle


9. Rothman, J
.

and Siegenthaler, E
.,

“Re
sponsive Space Launch


The F
-
14 Microsatellite Launch


Vehicle
,”
AIAA
-
LA Section/SSTC 2003
-
9002
,
1
st

Responsive Space Conference,
April 1
-
3, 2003, Redondo

Beach, CA


10. Chen, T.
,

Ferguson, P.
,

Deamer, D
.,
Hensley, J.
,

The Boeing Company.


“Responsive Air Launch Using F
-
15 Global Strike Eagle,” 4
th

Responsive Space Conference, April 24
-
27,


2006. RS4
-
2006
-
2001


11. Socher, A
., a
nd Gany, A
.,

6
th

Responsive Space Conference, April 28
-
May1, 2008, Los Angeles,


CA. AIAA
-
RS6
-
2008
-
5
003


12. Sarigul
-
Kliijn, M
.

and Sarigul
-
Kliijn, N
,

Mechanical and Aeronautical Engineering Dept.,


University of California, Davis, CA 95616. “A Study of Air Launch Methods for RLVs,” AIAA 2001
-


4619, 2001.


13. Hague, N., Siegenthaler, E.,

and Rothman, J., “Enabling Responsive Space: F
-
15 Microsatellite Launch


Vehicle”, Proceedings of the Aerospace Conference, 2003. 2003 IEEE Volume 6, March 8
-
15, 2003, pp


6_2703


6_2708, IEEEAC paper #1102.