AIAA 2012-0584

50th AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition
09 - 12 January 2012, Nashville, Tennessee
AIAA 2012-0584
Lockheed Martin’s Samarai Nano Air Vehicle:
Challenges, Research, and Realization
Steve Jameson1 and Dr. Kingsley Fregene 2
Lockheed Martin Advanced Technology Laboratories, Cherry Hill, New Jersey, 08002
Ming Chang3
Lockheed Martin Advanced Development Projects, Palmdale, California, 93599
Dr. Ned Allen4
Lockheed Martin Advanced Development Projects, Bethesda, Maryland, 20817
Harold Youngren5
Aerocraft Consulting, Portsmouth, Maine, 04103
and
Joe Scroggins6
North Carolina State University, Raleigh, North Carolina, Zip Code
Renewed interest in centimeter-scale flight vehicles for short-range urban surveillance
and sampling missions led to the Defense Advanced Research Projects Agency’s Nano Air
Vehicle (NAV) program (2006-2008). This effort explored technology for fixed wing, rotary
wing, or flapping wing vehicles of 7.5 centimeter length and 10 grams total mass that can fly
for 20 minutes with a range of 1 kilometer. Under this program the authors participated in
a Lockheed Martin-led development of a unique self-propelled monowing rotorcraft,
Samarai, inspired by a maple seed [1]. Following the end of the NAV program, continued
research and development led by Lockheed Martin has resulting in successful demonstration
of flight by Samarai vehicles at multiple scales. 7
Nomenclature
A
a
Cp
Cx
Cy
c
dt
Fx
=
=
=
=
=
=
=
=
amplitude of oscillation
cylinder diameter
pressure coefficient
force coefficient in the x direction
force coefficient in the y direction
chord
time step
X component of the resultant pressure force acting on the vehicle
1
Lockheed Martin Fellow, Advanced Technology Laboratories, 3 Executive Campus, Suite 600, Non Members
Principal Research Scientist, Advanced Technology Laboratories, 3 Executive Campus, Suite 600, Senior AIAA
Member
3
Lockheed Martin Fellow, Vehicle Systems and Sciences, 1011 Lockheed Way/MS 1100, Professional AIAA
Member
4
Lockheed Martin Senior Fellow, Professional AIAA Member
5
Consultant, Aerocraft Consulting, 129 Pitt St., AIAA Member
6
PhD Candidate, Aerospace Engineering
7
The views expressed are those of the authors and do not reflect the official policy or position of the Department of
Defense or the U.S. Government.
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2
Copyright © 2012 by Lockheed Martin Corporation. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.
Fy
f, g
h
i
j
K
=
=
=
=
=
=
Y component of the resultant pressure force acting on the vehicle
generic functions
height
time index during navigation
waypoint index
trailing-edge (TE) nondimensional angular deflection rate
I. Introduction
Renewed interest in centimeter-scale flight vehicles for short-range urban surveillance and sampling missions led
to the Defense Advanced Research Projects Agency’s Nano Air Vehicle (NAV) program (2006-2008). This effort
explored technology for fixed wing, rotary wing, or flapping wing vehicles of 7.5 centimeter length and 10 grams
total mass that can fly for 20 minutes with a range of 1 kilometer. Under this program the authors participated in a
Lockheed Martin-led development of a unique self-propelled monowing rotorcraft, Samarai, inspired by a maple
seed [1]. Following the end of the NAV program, continued research and development led by Lockheed Martin has
resulted in successful demonstration of flight by Samarai vehicles at multiple scales.
II. Samarai Concept
In 2005, DARPA’s Defense Sciences Office (DSO) initiated the Nano Air Vehicle (NAV) program to investigate
technologies for vehicles on the scale of 7.5cm length and 10g total mass (including 2g payload) that can fly for up
to 20 minutes with a range of 1km. To meet these requirements Lockheed Martin (LM) proposed a novel rotary
wing vehicle, the Samarai, inspired by the samara (the winged seed of many plants as exemplified by the maple
seed). The single-bladed rotorcraft concept (Figure 1) has an overall length of 7cm and weight of 10g.
Samarai
Samara
Figure 1 Biologically Inspired Samarai Nano Air Vehicle
Though similar in some aspects to the autorotating maple seed, the Samarai is self-propelled for hover and forward
flight by a tip jet engine and includes a cyclic/collective flap for flight control. It is mechanically simple with few
moving parts and includes a central fuel tank for propellant and electronics for communication, sensing and control.
The 2g payload is located on the underside of the fuel tank and can be dropped as part of the mission. The Samarai
rotates at roughly 12,000 RPM for hover and forward flight.
During the execution of the program the LM team quickly discovered a number of challenges associated with the
size and operating environment that is unique to the Samarai design. For the most part typical flight vehicle
development programs entail a demonstration of subscale components or vehicles for proof of concept which then
progresses to the full scale vehicle development. The Samarai concept is just the opposite. The full size vehicle is
actually a hand size vehicle where components and subcomponents are not readily available for integration.
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Therefore the challenges that surface included miniaturization of component parts, durability and operability of the
miniaturized components and operation in the micro scale environment.
III. NAV PROGRAM ACCOMPLISHMENTS
Lockheed Martin’s biologically inspired maple seed flyer concept, although simple in design and application proved
to be challenging in its development and deployment due to the nature of the environment it operates in. Despite
these challenges, the Samarai program made a number of significant technical accomplishments,
• Validated fundamental aerodynamics of the Samarai through extensive Computation Fluid Dynamics
(CFD) analysis, wind tunnel testing, 6-DOF simulation and large-scale prototype vehicle testing.
• Validated feasible propulsion technique with extensive trade analysis and experimentation of multiple
propulsive schemes. Significant experimentation was conducted on ramjets and pulsejets such that the
micro pulsejet showed sufficient technical viability for incorporation onto the final vehicle concept.
• Validated feasible lift modulation technique through extensive computational analysis and
experimentation. Trade studies include CFD and wind tunnel evaluation of mechanical, fluidic and
plasma lift modulation schemes and MEMS, piezoelectric and conventional electromechanical flap
actuation schemes. This resulted in the selection of mechanical flap with piezoelectric actuation as the
preferred system.
• Developed a concept for GNC from rotating non-inertial platform. Initial trade studies yielded a
hybrid optical/magnetic approach for sensing vehicle state which provides information for operator
guidance. Algorithm development and simulation validated this technique.
• Developed flyable configuration for Samarai air vehicle. At end of phase 1, demonstrated technologies
in propulsion, sensors, electronics and battery at miniature scale.
From these accomplishments LM completed the Phase 1 effort by adapting existing technologies in propulsion,
actuation, guidance, navigation and control and sensor imaging into a “feasible configuration” to demonstrate a low
risk design. Figure 2 reflects our design evolution ending in a feasible design.
1
Solid Fuel
Multiple Wings
2
Mass and
inertial balance
3
Liquid fuel Baselined
Blade Element Model
Design Evolution
Engine
Redefinition
5
Design for
Mission profile
(20 Minutes)
6
Spheroid fuel tank
Straight wing
7
Pulse Jet
Longer wing reduces
required thrust
Figure 2 Evolution of Samarai Nano Air Vehicle Design
To further reduce risks for the NAV program two large-scale flying prototypes were developed to assess hover
power requirements and cyclic/collective control capability utilizing trailing edge flap actuation. Flight testing of the
pair (Figure 3) showed both stable hover and adequate vertical flight control via collective flap. Through concurrent
testing on a tether, the prototypes also showed efficacy of monowing cyclic control through phased actuation of the
trailing edge flap.
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Figure 3 -- 500g RC prototype showed stable hover and collective climb in flight. Flight of 100g monowing
prototype (right) shows flat coning angle in hover.
A. Overview
The Samarai maple seed flyer is inherently a stable and resistant to upset. Its great appeal is simplicity in design and
operation over concepts such as flapping wing designs or rotorcraft designs. The insights that drove the design of the
Samarai vehicle are summarized in Table 1
Table 1 Key insights driving Samarai concept
Observation
The mechanical complexity of conventional hovering vehicles
(rotorcraft) will not have sufficient robustness for battlefield operation
when scaled down to centimeter scale
Nature has demonstrated the required aerodynamic principles in an
existing system of exactly the right scale – the maple seed
Helicopter flight principles show us how to achieve forward flight from a
rotating lift-producing airfoil
Conventional mechanically-scanned sensors will not scale to fit on a
centimeter-scale vehicle.
Response
Unitary body structure and minimal
moving parts
Rotating monowing
Controlled lift modulation using trailing
edge flap
Electrically controlled camera for 360degree controlled coverage
In carrying the concept forward during Phase 1 the Samarai trade-space explored design choices that were made on
the basis of requirements for a design mission profile. That profile encompassed indoor and outdoor flight from
hover to forward flight and speeds up to 0.5 m/s for a range of 1 km. Figure 4 shows the Samarai mission profile
with accompanying segment descriptions.
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5,000 ft Altitude
6
7
3
5
4
2
8
1
Phase
Description
1. Takeoff and Checkout
60 s Hover
2. Ingress to Target
1.0 km dash at 8 m/s
3. Navigate Through Bldg.
55 s of 0.5 m/s flight
4. Hover to Deploy Pld.
60 s of Hover to deploy Payload
5. Deploy Pld.
Mandate payload release, 2 g mass reduction.
6. Navigate Through Bldg.
55 s of 0.5 m/s flight
7. Egress From Bldg.
1.0 km dash at 6 m/s
8. Hover Reserves
60 s Hover
Figure 4 --Samarai design mission profile
This mission profile drove the evolution of the design as depicted in Figure 5, trading off parameters such as fuel
(energy density), engine concept (thrust), controls and sensor choices to mature the Samarai into a design that can
autonomously fly into and out of a building, through window openings with no loss in communication. For each of
these configurations careful balancing of system components was done to achieve the characteristics that play a
large role in the vehicle’s natural stability making it progressively more capable to perform the functions describe
above.
B. Propulsion Development
The feasibility of valveless micro pulsejets for use with micro aerial vehicles was examined under this project.
Pulsejets, acoustic resonant devices whose natural frequency is proportional to the inverse of their lengths, were first
developed in the early 1900s and achieved notoriety as the propulsion system for the German V-1 cruise missile in
World War II. Lacking a mechanical compression mechanism like that utilized in a gas turbine engine, pulsejets
suffer from relatively poor performance at the large scale when compared to modern air-breathing propulsion
systems. Though this low pressure ratio is detrimentally disadvantageous on the large scale, the simplicity of the
system compared to complex turbomachinery makes the pulsejet a potentially attractive option for scaling down for
micro aerial vehicle propulsion.
Design Concept
Key Choices
Design Concept
Key Choices
Solid Fuel
Explored multiple wings for
thermal mitigation
Redesign hub based on results
from 6-DOF simulator
Expanded hub to meet center
of gravity requirements
Expand fuel tank based on need
to meet 20-minute mission
Propane fuel
Developed wing based on
blade-element model
Explored counter-rotating
electric option
GN&C system change
• Removed Accelerometer &
Magnetometer
• Added second image sensor
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Defined vehicle endurance
based on mission model
Adopted Bulge for balancing
moments of inertia
Conceptual engine design
Mount payload beneath hub
Baseline straight engine
Consider 90-degree bent
engine
Shape hub to accommodate
GN&C requirements
Adopt cylindrical fuel tank
Straight wing significantly
reduces inertial loads on wing
flap.
Spheroidal fuel tank
significantly reduces mass of
fuel tank vs.
cylindrical tank
Flyable configuration based on technology state at end of Phase 1:
• Longer wing reduces thrust requirement from engine
• Spherical fuel tank accommodates required amount of hydrogen fuel
• Reshaped hub accommodates existing COTS battery that meets
requirements
Figure 5 -- Major Samarai Nano Air Vehicle design iterations
In an effort to investigate feasibility for a range of micro pulsejets, several pulsejet lengths were examined. Both
experimental and numerical approaches were used in developing an understanding of valveless micro pulsejets.
Pulsejets traditionally operate in a valved mode, but this work focused on valveless models due to the high operating
frequency expected with micro pulsejets and the potential unreliability of valved micro pulsejets due to the
likelihood of valve damage during operation. Three pulsejet scales were investigated: 15 cm, 8 cm, and 4 cm. The
results of the 15 cm and 8 cm studies were reported in References [8] and [9]. The 15 cm pulsejet was a scaled
model of a 50 cm Bailey Machining Services pulsejet, while the 8 cm pulsejet was a half-scale version of the 15 cm
except for the combustion chamber volume, which was held constant. This section will highlight many of the
important results from these works, as well as some performance metrics not reported. The 4 cm pulsejet was
operated as a proof of concept, but thrust and other performance metrics were not measured.
1. Computational Modeling of Valveless Pulsejets
A computational model of these valveless pulsejets was developed in order to gain a better understanding of the
operation of the system. The details of the computational models, carried out in CFX, can be found in References [8]
and [9]. The operation of the pulsejet is described by these models in the following 10 steps.
1. Combustion event begins when hydrogen and air mix and are brought to their auto-ignition temperature
through mixing with residual hot products from the previous cycle. The pressure and temperature begin to
increase in the combustion chamber. Air continues entering the combustion chamber through the inlet with
reduced velocity.
2. Combustion continues, and peak pressure and temperature are reached in the combustion chamber.
Compression waves are generated and propagate into the inlet and the exhaust tube. When the pressure of
the hot gases becomes equal to the pressure of the cold air, the velocity goes to zero at the interface of these
two gases.
3. Expansion waves are generated at the inlet and decrease pressure in the combustion chamber. A positive,
increasing velocity characterizes the flow at the exhaust duct exit; while at the inlet, the hot products are
expelled with an increasing (negative) velocity.
4. Expansion waves are generated at the exhaust duct exit and travel back to the combustion chamber. Pressure
decreases in the combustion chamber and the gas velocity out of the inlet and the exit reach their
maximum. Most of hydrogen is burned by this stage.
5. The pressure in the combustion chamber continues decreasing. Temperature increases in the inlet.
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6. Expansion waves from the exit enter the combustion chamber and further decrease the pressure in the
combustion chamber. The outgoing (negative) velocity at the inlet decreases to zero.
7. The combustion chamber pressure decreases below atmospheric, causing air to enter the combustion chamber
through the inlet. Hot products continue to be expelled from the exhaust duct exit but the velocity
continuously decreases.
8. The pressure and temperature in the combustion chamber continue decreasing while the inlet velocity
continues increasing. The product velocity at the exhaust duct exit goes to zero and then actually reverses.
This backflow causes a temperature drop due to entrainment of ambient air up the exhaust duct.
9. Cold air from the inlet continually enters the combustion chamber. Hot gas in the exhaust duct is pushed back
to the combustion chamber. The pressure in the combustion chamber continues to increase.
10. Backflow continues, but its negative velocity becomes smaller. When the pressure in the combustion
chamber approaches atmospheric pressure and air from the inlet mixes with hydrogen in the combustion
chamber, the next cycle begins.
The computational models developed in this work were validated against experimental work for pulsejets of both
sizes. Several parameters were investigated, including pulsejet operating frequency as a function of inlet geometry,
effect of inlet orientation, and pressure profiles within the pulsejet chamber. Additional experimental work
investigated the operability range for a pulsejet length for a given inlet length and time resolved thrust on the 8 cm
scale. Several conclusions were drawn that contribute to the general understanding of pulsejet operation on the
micro scale. The conclusions from these studies are enumerated below.
2.
Fifteen cm pulsejet conclusions
The numerical model of the 15cm pulsejet [8], validated by the experimental results, shows that the expansion
waves from the outlet are able to enter the combustion chamber before the next cycle. In each cycle, the combustion
chamber pressure is decreased by the expansion waves from the outlet and the inlet as well. The backflow from the
exhaust duct cannot travel to the combustion chamber, and the oxygen for the reaction is provided by the airflow
from the inlet only. The exhaust duct may contain hot gases generated by several successive combustion events.
This behavior is not observed in the valved pulsejet. The cold air entering the combustion chamber creates a strong
vortex that greatly accelerates the mixing process.
The model shows that exhaust duct length and inlet length are directly coupled in the valveless pulsejet. Pulsejet
operation at minimum overall lengths was achieved with the shortest inlet lengths. Pulsejet operation at longer jet
lengths was achieved with longer inlet lengths and smaller inlet area ratios Maximum exhaust duct lengths exist for
corresponding inlet lengths and area ratios. Results suggest that for given a valveless configuration, there exists a
critical length above which pulsejet operation cannot be achieved. Minimum exhaust duct length can be further
reduced (from a constant area geometry) by about 16% on average with the addition of a diverging exit nozzle.
Pulsejet operation was achieved from 90°, 135°, and 180° inlet orientations with respect to the conventional forward
facing valveless design.
Preliminary tests indicate that valveless pulsejets with multiple opposed facing inlets present similar behavioral
characteristics as those of conventional design. It also may be concluded that the consolidated inlet area ratio of
multiple inlets is comparable to that of a single rearward-facing inlet, taking into consideration inlet placement and
internal direction of inlet flow.
The two main parameters determining the success of a pulsejet to operate are chemical kinetic time versus jet length
and inlet area to combustor area ratio. At shorter lengths, the chemical kinetic reaction rate (combustion time)
becomes challenged by the period of fluid mechanic oscillations. This would explain why fuels with longer chemical
time scales such as propane did not permit pulsejet operation in smaller jet sizes.
3.
Eight cm Pulsejet Conclusions
Computational modeling of the 8 cm pulsejet [9] provided physical insight into the pulsejet operation. The simulated
operation frequency and peak pressure matched with experimental data. It was observed that for each operational
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cycle, combustion consumes most of the oxygen in the combustion chamber, and the oxygen comes from the inlet
only. Acoustics and fluid mechanics are both important in determining the operating characteristics of these engines.
In the traditional valved inlet, the operating frequency is solely a function of the jet length. However, in valveless
mode, the operating frequency is also a function of inlet length, but does not act as a 1/4 wave tube. Rather, the
frequency scales with the inlet length raised to negative 0.22 power.
The operating frequency and peak pressure rise are a function of fuel mass flux. At low fuel mass flux, both
frequency and pressure are low and increase with increasing fuel mass flux. With the forward-facing inlet, the
frequency and pressure both have a maximum, whereas with the rearward-facing inlet, the frequency continues to
increase until the maximum fuel mass flux is reached. The pressure reaches a maximum at lower fuel mass flux, but
does not decrease as fuel mass flux continues to increase. With forward-facing inlets, the net thrust is very low as
expected. With rearward-facing inlets, the net thrust improves to approximately 1 N, resulting in a TSFC of 0.02
kg/N-h.
4. Experimental Performance Assessment
Further work was done beyond that reported in References [8] and [9] to characterize the performance of 8 cm
pulsejets at a range of potential flight conditions. This includes studying varying inlet lengths and diameters over a
range of fuel flow rates and simulated flight conditions. While performance metrics were measured for forwardfacing inlets, pulsejets with rearward-facing inlets and hybrid inlets with both rearward and forward inlets were also
tested. Because of the expected flight velocity of the system and the limitations of the experimental set-up, the
rearward and hybrid inlets were only shown as proofs of concept and not tested for performance. While the
combustion chamber volume and exhaust length were held constant at 3x10-6 m3 and 51 mm in length respectively,
the forward-facing inlets varied in length from 0.8 to 1.8 cm in length and 0.25 cm to 0.35 cm in diameter. The
exhaust tube was tested in both flared and unflared geometries. Figure 6 shows a sample 8 cm pulsejet.
Figure 6 -- 8 cm pulsejet with flared exhaust and forward-facing inlet
The left chart in Figure 7 shows the thrust and specific impulse produced by an 8 cm pulsejet with one of the
experimental inlets at a simulated free-stream velocity of 50 m/s over a range of fuel flow rates. The right chart
shows the effect of free-stream velocity on the pulsejet thrust and specific impulse for a given inlet at a constant fuel
flow rate. As evident in the figure, the pulsejet performance improves with increasing velocity, though this particular
inlet peaks at 40 m/s and then degrades at 50 m/s.
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Figure 7 -- Thrust and specific impulse for a valveless 8 cm pulsejet as a function of fuel mass flux (left) and
simulated forward flight speed (right).
In addition to the 15 cm and 8 cm pulsejets, a 4 cm pulsejet was designed and tested. This pulsejet, a half-scale
model of the 8 cm pulsejet, was run on hydrogen fuel and had an operating frequency of 2800 Hz with an inlet that
was 0.5 cm in length and 0.16 cm in diameter. Performance data other than the frequency are not available.
5. Pulsejet Fuel Assessment
Fuel operability was also an important topic of interest to this work. All reported performance data are from
pulsejets utilizing hydrogen fuel, an unattractive candidate for micro aerial vehicle use due to the storage
requirements. Acetylene, another logistically-unfriendly fuel, was also shown to work on the 8 cm pulsejet. Room
temperature propane and MAPP gas, two fuels with very attractive storage attributes (in that they liquefy at
reasonable pressures), did not permit operation. The concept of fuel mixtures and fuel preheating were also
investigated. Fuel mixtures of propane and hydrogen, up to 14:1 on a mass basis, were shown to operate on the 8 cm
pulsejet when preheated to approximately 400°C. Because of the abundance of heat available in the walls of the
pulsejet (as seen by the red glow in Figure 6), fuel preheating appears a promising development. Figure 8 shows
how this concept could be implemented, where the fuel injector runs the length of the exhaust prior to reaching the
combustion chamber.
Figure 8 - Fuel regenerative heating concept.
The feasibility of small-scale pulsejets was thus established under the NAV program, with thrust and specific
impulse produced by the 8 cm pulsejet near the range required. Subsequent work was performed later that
investigated the possibility of an ejector-style thrust augmenter on the 8 cm pulsejet to improve performance [10].
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This research showed that the exhaust thrust of an 8 cm pulsejet could be more than doubled at static conditions,
much like a thrust augmenter on the 50 cm scale. The work also numerically investigated thrust augmentation at
flight speed using a modified ejector. It was shown that a 5 cm valveless pulsejet could produce an additional 70%
of thrust at a flight speed of 50 m/s, further establishing the valveless micro pulsejet as a viable NAV propulsion
system. In addition, the possibility of operating micro pulsejets in a valved mode was investigated. None of the
proposed designs resulted in operable valve systems on the small scale however both computational and
experimental tools were developed that should aid future work in the development of valves for micro pulsejets [11].
C. Aerodynamic Feasibility
The aerodynamic design features for the Samarai NAV are defined by the performance that drives system sizing and
requirements for the rotor lift, cyclic lift modulation and controls. Aerodynamic design for the Samarai included the
rotor layout, rotor blade analysis and sizing of the flap for collective and cyclic lift control.
As a consequence of its small size, the Samarai rotor operates at relatively low chord Reynolds numbers (Re) of 15K
to 40K. At these Re, airfoils typically operate with a large region of laminar flow inducing a thicker boundary layer
than a turbulent one which results in higher drag penalties. In addition at these low speeds the boundary layer
buildup can be substantial causing local separation bubbles that result in nonlinearities in the aerodynamic data. The
AG38 airfoil was selected for the rotor based on its performance within the range of 20K to 40K Re. This airfoil
section, shown in Figure 9 with the 10% chord TE flap, was developed by Dr. Mark Drela (MIT) for use on model
aircraft and has a thickness ratio of 7.04% t/c.
Figure 9 -- AG38 airfoil selected for rotor (with 10% chord flap)
In order to develop the correct disc loading for the Samarai design, representative rotor sectional aerodynamic data
at low Re (15K-40K) was needed for use with our blade-element analysis tool. As part of the Phase 1 program wind
tunnel testing was done at NASA Langley in the 2x3 boundary layer wind tunnel shown in Figure 10 to obtain
aerodynamic characteristics for the rotor section and lift control flap.
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Figure 10 -- Samarai AG38 airfoil model in NASA 2x3 wind tunnel
This tunnel was select due to its low turbulence levels (<0.011% at speeds to 30 ft/s). Tests were done using a 2D
wall-to-wall model of the AG38 airfoil equipped with a full-span 10% chord flap. The test used pressure taps on the
wing and wake rakes to obtain force, moment and drag data at Re from 15K to 60K. Further details of the testing
and results for airfoil and flap characteristics at these low Re are presented in reference [4]. Some of these results
are presented to show basic aerodynamics for the rotor sections. Figure 11 shows lift, drag and pitching moment for
the AG38 tested with turbulator strip at Re from 15K to 40K.
Figure 11 -- Aerodynamic characteristics (CL,CD, CM) for AG38 rotor airfoil with turbulator at Re 15K-40K
The AG38 airfoil was tested in both clean condition and with a turbulator developed through systematic testing to
improve lift, drag and flap effectiveness (linearity of flap lift changes) at the low Re required for the Samarai rotor.
A 0.015 inch thick trip strip placed at 15% x/c was found to give the best results for improving lift and flap linearity
with the lowest drag in the 20K – 40K Re range.
Aerodynamic characteristics for the 10% chord flap are shown in Figure 12. The results show lift as a function of
angle of attack for the flap at -12o, -5o, 0o, +5o and +12o for Re from 15K to 40K. Note that flap effectiveness was
improved by the turbulator and gives adequate lift control with flap deflection at Re from 20K to 40K but still shows
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reduced linearity at an Re of 15K. Analysis and wind tunnel testing confirmed that this flap is capable of providing
adequate lift modulation for forward flight.
LS2 WT Test, AG38 Plain flap, 15K T3B turbulator
LS2 WT Test, AG38 Plain flap, 20K T3B turbulator
1
1.2
run81 turb Flap-12
run82 turb Flap-12
run112 turb Flap-5
run113 turb Flap-5
0.8
run99 turb noflap
run98 turb noflap
run116 turb Flap+5
1
run117 turb Flap+5
ru85 turb Flap+12
ru86 turb Flap+12
0.6
0.8
0.6
c la p
c la p
0.4
CL
CL
0.4
0.2
0.2
-6.00
-4.00
-2.00
0
0.00
2.00
4.00
6.00
8.00
10.00
-6.00
-0.2
-4.00
-2.00
-0.4
0
0.00
-0.2
8.00
10.00
LS2 WT Test, AG38 Plain flap, 40K T3B turbulator
1.2
1.2
run83 turb Flap-12
run84 turb Flap-12
run114 turb Flap-5
run115 turb Flap-5
1
run101 turb noflap
1
run119 turb Flap+5
run118 turb Flap+5
0.8
run87 turb Flap+12
run88 turb Flap+12
c la p
c la p
CL
CL
0.4
0.4
0.2
0.2
-2.00
0
0.00
-0.2
0.8
0.6
0.6
-4.00
6.00
angle
Alpha
(deg)
LS2 WT Test, AG38 Plain flap, 30K T3B turbulator
-6.00
4.00
-0.4
angle(deg)
Alpha
run100 turb noflap
2.00
2.00
4.00
6.00
8.00
10.00
-6.00
-4.00
-2.00
0
0.00
-0.2
2.00
4.00
6.00
8.00
10.00
-0.4
-0.4
angle(deg)
Alpha
angle (deg)
Alpha
Figure 12 -- Lift curves showing flap effectiveness for AG38 with turbulator at Re 15K-40K with -12o, -5o, 0o,
+5o and +12o flap deflection
D. Flight Control
Simulation played a key role in the Samarai development in terms of system trade-offs and control system design.
Results from the simulation effort led to a better understanding of the effects of the vehicle inertia properties for
stabilizing the monowing and for engine and lift control requirements for forward flight.
6-DOF Rigid-Body Simulation
A 6-DOF rigid-body simulation of the Samarai vehicle dynamics, based on reference [5] was developed [8] and
extended for use on Samarai by including a tip thruster (engine model) to provide torque for active rotation and with
linear inflow models for hover and forward flight. An enhanced blade-element rotor model calculated aerodynamic
forces on the rotating wing using low-Re aerodynamic data obtained from our NASA wind tunnel test, described in
the previous section.
The simulator was developed in MATLAB and was used for trim state analysis or time domain simulation with open
or closed loop control of the engine thrust and lift flap. Its base is a CAD model of the vehicle for rotor geometry
for the blade elements and used tables of aerodynamic data for the rotor sections. A convenient Graphical User
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Interface (Figure 13) allowed entry of model data, including inertias and simulator parameters, and visualization of
simulation output.
Figure 13 -- The Graphical user interface for the SAMARAI 6-DOF simulator provides control over
simulation parameters and an interface to the CAD model representation.
Control Development
To determine the optimal control scheme a trade study into the level of human-in-the-loop control for the vehicle
was undertaken. This trade evaluated options for human control in:
1. Direct human-in the loop: Human does everything: directly commanding actuators
2. Body level commands: pitch-roll rate, thrust
3. Human provides disk level commands: velocity commands: x, y, z
4. Human gives position commands: automatic navigation system (nav) controls velocity and relative
position
5. Human inputs abs position: automatic navigation system (nav) performs global registration (e.g.,
Simultaneous Localization and Mapping -- SLAM) and controls to position
The trade study selected option 3, as the first two were deemed infeasible for even an experienced human pilot and
the latter two were deemed to require too great a degree of onboard sensing and processing.
Having the flight controls based on the control of the rotor disk in the inertial frame simplifies the equations of
motion of the Samarai to the disk motion of the center of mass and disk orientation, as shown in Figure 14. This
disk model can further be simplified when placed into the world frame, producing a relatively intuitive control
model, shown in Figure 15.
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Figure 14 -- Fixed disk Equation of motion
Figure 15 -- Samarai approximate control model
This simple proportional-differential feedback controller was implemented for RPM control (with engine thrust) for
vertical velocity control and longitudinal and lateral velocity control using the lift control flap. A simple “mission”
using velocity control inputs for vertical, forward and lateral velocities was computed using the 6-DOF simulator.
The results for the trajectories of the CG and the tip are plotted in Figure 16, demonstrating basic three-axis control
of the Samarai. With this model, integration of better sensors or alternate control schemes for relative position
control can easily be evaluated.
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Figure 16 -- Trajectory from "mission" using simulator with closed loop velocity control
E. Sensing
The Samarai vehicle combines elements of onboard and off board control to reduce operator task saturation. Off
board control consists of commands provided by a remote operator for vehicle velocities and direction while
onboard control maintains vehicle stability and motion in accordance with those commands obtained through
sensing of vehicle state and control of vehicle actuation. Samarai’s GNC system therefore contains a single image
sensor, magnetometer, processor and various other supporting components that enable the operator to remotely fly
the vehicle.
Sensors for Vehicle State and Operator View
Samarai’s sensing is driven primarily by its extreme size and power constraints. Due to these constraints it was
necessary to take a minimalist attitude towards onboard sensing functionality. As a result multiple functionalities
were extracted from sensors used onboard the vehicle. As an example, we determined a method for calculating
relative vehicle motion through a well-known optical flow technique called cross-correlation. For the Samarai
vehicle, cross-correlation can be used both for image stabilization and for optical flow.
The Samarai design leverages the inherent rotation of the vehicle to multiplex the use of a single sensor into several
“virtual sensors”, each of which point in a direction 90 degrees from the previous as shown in Figure 17. This is
accomplished by capturing an image at exactly the same point in rotation, for four different locations.
For this approach, we required an image sensor with a very high frame rate and an extremely short integration time
so that images could be captured with minimum blur. A research group at Kodak had developed a small, grayscale
image sensor, the KAC-9630 with a high frame rate and short integration time. This sensor, developed for a
commercial product, met the specifications required for the Samarai control approach.
To ensure that images from this sensor could be utilized for the calculation of vehicle state, we performed
experiments and collected imagery from within a simulation that we processed in real time. In simulation, we
constructed a vehicle model with realistic dynamics that had simulated camera feeds, each 90 degrees from the
previous. Our software polled those feeds and sent them to a Marvel processor (baseline processor for the Samarai
vehicle) and processed the images in real-time to determine pitch, roll and yaw. Our experiments indicate an
accuracy of better than 0.5 pixel in estimating pitch, roll and yaw (about 0.6 degrees). This accuracy can be
improved by allowing larger word size in the software. Timing estimates indicated the software can run in real-time
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on a 120 MHz, fully loaded computer (full operating system, background networking and processes, etc). On a
processor dedicated to these operations, about half this processor speed would be required.
Figure 17 -- Optical flow using a camera capable of sensing multiple frames per rotation enables both disk
state sensing
Operator View Management
A key problem for the operator control is the ability to stabilize the images between frames to ensure that the
operator has sufficient situational awareness to navigate the Samarai vehicle. To perform this, we developed an
algorithm to do the following:
•
•
•
•
Integrate over multiple frames to increase signal
Process to improve contrast and increase detail
Correct for blur due to yaw, pitch, roll and translation
Compress by a factor of 8 for transmission over radio link
We validated the performance of this algorithm using the Kodak sensor in our laboratory through a series of
experiments. A difficult challenge that we had to overcome in constructing an experimental environment was that
the sensor, as it was received in our development kit, required an image capture card, a large lens and a large
mounting board. This prohibited us from spinning the sensor as it would be onboard the vehicle, however we
wanted to verify that it could capture images as the environment was spinning. To solve this, we reversed the
environment by keeping the sensor stationary and spinning the environment that we wanted to capture through
sensing. We created a spinning test rig that could rotate a platform at up to 70Hz. This test rig was extremely
valuable and was used for multiple experiments when we wished to verify that components could sense the
environment while spinning.
To the spinning platform atop the rig, we affixed a label that contained the alphabet so that the letters would face
outward as the platform was spinning. The Kodak sensor was placed a few feet away from the rig and positioned so
that the spinning letters were within the field of view. The test rig was spun to 70Hz as the camera was adjusted
through software so that the frame rate matched the speed of the test rig. Figure 18 shows the images collected and
processed with our software. The first image is a single frame showing white noise and low signal due to the high
frame rate. The second image shows a clearer, but blurred, image as we integrate across 14 images to increase the
signal of the image. Finally, the third and final image shows a yaw corrected version that integrates signal from all
14 images, removes most of the white noise and produces an image that is very clear with detail easily discernable
by an operator.
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Images collected as test rig was spinning at 70 Hz
Single frame
14 frames
Yaw corrected
Figure 18 -- Software techniques were utilized to obtain clear images from a spinning platform
The result of this series of experiments was very encouraging and showed that our algorithm could perform the
necessary corrections and produce an image that was clear to the eye. What remained, was to determine whether an
optimized version of the algorithm could be produced that could operate in real-time on the actual hardware that was
selected. In addition, we also wanted to ensure that an operator could fly a vehicle given this type of image.
Operator Usability
In order to assess whether an operator could navigate a vehicle with imagery similar to what was being produced by
the Kodak sensor, we conducted a study that evaluated the operator’s ability to control a moving vehicle. One
hypothesis that we had was that an operator might perform poorly with solely a video stream. To test this
hypothesis we evaluated the usefulness of providing range to the operator. In this study, we mounted an image
sensor on a laboratory robot, and wrote software to process the image data so that it would have the same
characteristics (resolution, frame rate, contrast, etc.) as the Kodak sensor. We conducted a set of trials with
multiple participants to determine their ability to successfully maneuver the robot through an indoor course to locate
a goal. Trials were conducted under a number of conditions, including differing frame rate, and with and without a
range sensor. The results of the study (Table ) indicate that availability of a range sensor significantly increases the
ability of an operator to maneuver, above all other factors.
Frame rate
Experience
Sensor Data
2 Hz
5 Hz
Novice
Expert
Video
Average
15.2
Elapsed time min
12.8
min
17.4
10.6 min 23.5 min 4.6 min
Average
Number
collisions
12
16
15
19
15
Video +
Range
0
of
Table 3 Results from operator control study show range sensing eliminates collisions under operator control
We evaluated various concepts for incorporating range sensing into the vehicle. The next section will describe our
effort and detail the method we selected to assist the operator in navigation.
Collision Avoidance
We developed and tested in simulation a collision avoidance concept which did not provide a range estimate but did
allow for an adjustable “proximity sensor” utilizing an optical flow technique similar to that used by insects to avoid
collision. To evaluate this algorithm we tested a version of the software, running on a processor identical to that
selected for the vehicle, which was fed imagery that was collected from a simulation run where an operator flew the
vehicle in an urban environment and eventually collided with a building. The algorithm performed very well and
showed distinct spikes in proximity detection when the operator flew too close or collided with a building (Figure
19). Therefore integrating a collision avoidance system with a range estimator would definitely help the operator in
navigating the Samarai vehicle through tight spots and low light situations.
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Figure 19 -- Our collision detection algorithm uses optical flow and can alert the operator when proximity
reaches a certain threshold
At the conclusion of the NAV phase 1 program Lockheed Martin presented to DARPA a complete system
configuration based on these technologies, including identified electronic components and electric power, along with
predictions of aerodynamic and mission performance.
IV. SUBSEQUENT SAMARAI DEVELOPMENT
Subsequent to the DARPA NAV program, Lockheed Martin continued investment in development of the Samarai
system and underlying technologies. This effort has resulted in the family of maple seed-inspired air vehicles,
ranging in radius from 17cm to 72cm, as depicted in Figure 20. While the vision of a working 10g Samarai NAV has
not yet been achieved, some important milestones have been achieved in flight capabilities with these larger Samarai
prototypes.
R=17 cm, 50g
na
tur
R=30 cm, 200g
R=72 cm, 600g
inspired by nature
e
Maple Seed (Samara)
Nano Air Vehicle (NAV)
Micro Air Vehicle (MAV)
Demonstrator Air Vehicle (DAV)
Acer diabolicum Blume
Figure 20 -- The Samarai family of Maple seed-inspired UAVs have demonstrated important capabilities such
as vertical take-off and landing, full authority flight controls & onboard sensing, guidance and navigation
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A 2009 effort led to the demonstration of fully autonomous indoor and outdoor controlled flight with a 76cm, 600g
prototype (Figure 21) known as the demonstrator air vehicle (DAV), which is similar to that flown during the NAV
program [6]. Additional details about the DAV can be found in [7].
Figure 21 -- Samarai Demonstrator Air Vehicle, flown in 2009, uses an electric propeller opposite the wing
and has a stabilizer rod. The vehicle takes off vertically from a launch pin in the ground.
The DAV airframe consists of a single balsa wing rigidly attached to a hub that also has a balancing mass attached
to a boom of suitable length and inclination for stability. An electric motor and propeller mounted in a tractor
configuration on a separate boom orthogonal to the mass boom provides its propulsion. Stability of this design is
enhanced by the use of a vertical fin at the wing tip. The sole control surface on this vehicle is an aerodynamic flap
driven by a servo torque-rod assembly. The vehicle features full on-board autonomous operation with the ability to
automatically execute closed-loop waypoint navigation in both indoor and outdoor environments. Avionics that
enable these capabilities are housed in a pod at the vehicle hub.
Subsequent effort in 2010 led to the demonstration of the Samarai micro air vehicle (MAV) - a more capable 30cm,
200g prototype with an integrated camera unit. Similar to the DAV, the Samarai MAV features a motor driving a
small propeller mounted on the tip of the blade to provide propulsion. A portion of this blade is cut out to act as a
trailing-edge aerodynamic flap driven by a servo torque-rod assembly mounted at the root end of the flap for use as
the control surface. The other end of the blade is attached to an avionics/payload pod through adjustable spars that
enable its angle of inclination relative to the pod to be adjusted as needed. A pair of music wire bent into a skid
extending about 2-3 cm in front of the wing leading edge act as landing gears. These features are shown in Figure
22. In certain configurations, the MAV has demonstrated flight endurance of over 17 minutes, and of switching
seamlessly in flight between three flight modes, ranging from RC controlled to fully autonomous, all with downlink
of stabilized video from onboard the aircraft.
Figure 22 -- The Samarai Micro Air Vehicle has a tip-mounted propeller and requires no stabilizer bar. It
can be launched by hand or take off and land directly on any flat surface.
Continued Samarai development in 2011 has resulted in the demonstration of a 17cm NAV prototype weighing less
than 50g (Figure 23). This prototype is more similar to the DAV in that the propulsion system is not directly on the
wing but on a separate boom with adjustable effective lever arm to the avionics pod. This vehicle has shown
excellent flight characteristics. It features a highly miniaturized avionics suite enabling it to be automatically
controlled in all degrees of freedom.
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Figure 23 -- The Experimental Samarai Nano Air Vehicle design has a propeller mounted on an adjustable
boom. It has shown excellent flight handling abilities, including very stable hover in tight indoor spaces.
The work done on the various Samara prototypes under Lockheed Martin funding holds promise to yield both
militarily useful capabilities as a portable UAS for short-range surveillance missions and as an enabler for further
technology developments that will pave the way for the realization of the original Samarai NAV vision.
Acknowledgments
Much of the work described in this paper has been sponsored by the Defense Advanced Research Projects
Agency (DARPA) under contract number W31P4Q-06-C-0324. The authors would like to acknowledge the vision
and leadership of Darryll Pines, who created the Nano Air Vehicle program at DARPA, and of Todd Hylton, who
effectively carried on this vision. We would like to acknowledge the vision and leadership of Russ Frew of
Lockheed Martin, under whose sponsorship the post-NAV Samarai research and development work has been
conducted.
The views expressed are those of the authors and do not reflect the official policy or position of the Department
of Defense or the U.S. Government.
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