microsatellites at very low altitude

15 Aug 20061 Proprietary of Rafael - Armament Development Authority Ltd

20th Annual Conference on Small Satellites, 2006

20th Annual Conference on Small SatellitesAugust 14-17, 2006

SSC06-II-3

Microsatellites at Very Low Altitude

Dr. Hezi Atir, Dr. David Mishne

Space Systems Directorate, Rafael Ltd.



The Triad of the Presentation

Venµs TechnologicalMission


Hall EffectThruster

Hall EffectThruster

Microsatellites at Very Low AltitudeMicrosatellites at Very Low Altitude



Microsatellites at Very Low AltitudeMicrosatellites at Very Low Altitude



Low Altitude DefinitionWhere drag is significant and has to be considered (in orbit maintenance and other mission tasks)Affected by the cross section area Individual per architecture and mission.



Altitude vs. Cross Section Area5 10 15 20 25

240

260

280

300

320

340

360

380

400

0 1 2 3 4

Cross section (m2)

Alti

tude

(km

)

Kg hydrazine/yearKg hydrazine/year



Low Altitude Imaging MissionsLow altitude imaging missions benefit the low cost of smaller and less complex payload compared to higher altitude missions with similar performance Additional subsystems (i.e. communication), as well as the whole satellite become less expensive. The same applies to the launch cost

However, several issues should be considered as well in order to justify a low altitude mission compared with a higher altitude one.

However, several issues should be considered as well in order to justify a low altitude mission compared with a higher altitude one.


20th Annual Conference on Small Satellites, 2006 Low Altitude Advantages for

Imaging MissionPayload size drastically reducedImpact on payload, satellite and launch cost

200 300 400 500 60015

20

25

30

35

40

45

50

55

Satellite Hight, km

Req

uire

d A

pertu

re, c

m

Ground Resolution Distance = 1 mWavelength = 0.7 micrometer

Payload Size Vs Height

GSDhD λ22.1

=

Typical Imaging Missions

Low Alt. Imaging Missions

Altitude (km)

D(cm)h= Altitude

GSD=Ground Sampling Distance

λ=Wavelength



Drag ConsiderationsLow altitude requires frequent orbit maintenance

300 350 400 450 5000

2

4

6

8

10

12

14

16

18

20

Altitude, km

Fuel

Req

uire

d, k

g

XenonIsp = 1350 sec

HydrazineIsp = 220 sec

Area = 0.8 m2

Mission = 3 YearsSolar max. conditions

220 240 260 280 300 320 340 360 380 4000

2

4

6

8

10

12

14

16

18

20

Altitude, km

Fuel

Req

uire

d, k

g

XenonIsp = 1350 sec

HydrazineIsp = 220 sec

Area = 0.2 m2

Mission = 3 YearsSolar max. conditions

Fuel Required to Compensate for Drag



Hydrazine Propulsion vs. Electric Propulsion

A microsatellite with high quality payload is usually intended for long mission durations.

0

1

2

3

4

5

6

250 300 350 400

Altitude, km

Year

s

mAtf ∆>∗∗− *)1(ξ

ξ = Hydrazine / Xenon fuel consumptionf = Xenon required / (year * area) (kg/y*m2)t = Mission duration (years)A = Cross section area (m2)∆m = ∆ (empty mass) of EPS (kg)

A=0.5m2

∆m=15kgElectric Propulsion Zone

Hydrazine Propulsion Zone



Thrust Level

Orbit maintenance should not interfere with the mission → propulsion duty cycle should be short Also, short periods of very high drag due to extreme solar activity are expectedTherefore, thrust should be much higher (5-10 times) than the drag



Attitude Dynamics - 1

Moment due to drag:

Constant + attitude dependent

θV

( ) ( )θρρ ,0 MMM drag +=

Density Pitch angle



Attitude Dynamics - 2

( ) ( ) ( )θθρρθ ggMMMI ++= ,0&&

Pitch dynamics:

Non-attitude dependent drag moment

Gravity gradient moment

Attitude-dependent drag moment

Linearized dynamics:

( ) ( )ρθρθ θ 0MMI =+&&



Attitude Dynamics - 3( ) ( )ρθρθ θ 0MMI =+&&

The density is periodic (due to day/night cycle), with basic frequency equal to the mean motion n.

A good design approach:

Provide adequate aerodynamic stability such that the natural frequency of the pitch dynamics, , is much higher (at least twice) than the mean motion n.

IMθ



Orbit ControlThe orbit control law has to compensate for the exerted drag. Thus the satellite should have some autonomy in the orbit control Orbit determination becomes more complex both because of the unknown drag and because of the frequent thrust activations. The GPS system can be used for this taskSatellite autonomy should be included also in the mission planning (because of the orbit determination process)



Power Considerations

Electric propulsion system requires extra powerFor thrust level of 15mN, with efficiency of 40%, the system requires about 300wRequired solar panels area is about 1.5m2, and the additional weight is 6kg.



Coverage ConsiderationsSmaller field of regard (reachable area)Need more satellites or more time to cover a given area of interest



Communication ConsiderationsRequire less power and smaller componentsShorter real time data communication range

range

200 250 300 350 400800

1000

1200

1400

1600

1800

2000

2200

2400

2600

Satellite Altitude, km

Dis

tanc

e fro

m Im

aged

Site

to G

roun

d S

tatio

n, k

m

Elevation = 10 deg

GRS Antenna Elevation = 0 deg

Elevation = 5 deg



Hall EffectThruster

Hall EffectThruster

A key ingredient for low altitude

orbit maintenance

A key ingredient for low altitude

orbit maintenance



The HET- 300W Thruster

A small Hall effect thruster is developed, suitable to microsatellitesWill be tested aboard the Venus satellite

150mm



HET General PropertiesThruster mass – 1.5 kg

Thruster dimensions 169X119X91mm



HET PerformanceHET Performance

Nominal thrust: 15-33 mN

Nominal specific impulse: 1300-1650 sec

Input power (operating): 300 – 600 watts



Thruster ModelThrust Vs. Anode Power Isp Vs. Anode Power

0

200

400

600

800

1000

1200

1400

1600

1800

200 300 400 500 600

Power(watts)

Spec

ific

Impu

lse(

sec)

0

5

10

15

20

25

30

35

200 300 400 500 600

Power(watts)

Thru

st(m

N)

cathode constant mass flow of 0.1 mg/sec



HET SystemTotal HET system dry weight (with 2 thrusters) – 20 kg

PV

PT

FilterPressure

Regulator

PPU/TSU

XenonTank

HETGimbalsXFC

Sate

llite

C2

PPU/TSU- Power Processing Unit/ Thruster Selecting XFC- Xenon Flow Controller HET – Hall Effect Thruster PT – Pressure Transducer PV – Pyro Valve - - - - Control and Electric Line Fuel Line



Orbit Maintenance

Example: Sun-synchronous orbit IHET operation zones

Two activations of two different thrusters provide symmetrical orbit maintenance Eclipse





A test bed for low altitude imaging

mission

A test bed for low altitude imaging

mission



Venµs ProgramVegetation and Environment New µSatellite

Dual missions: – Scientific Mission: A research demonstrator mission

for the GMES program (Global Monitoring for Environment and Security), dedicated to monitoring vegetation and water quality – using a Multi-Spectralcamera

Ben Gurion University

SOREQ



Venµs ProgramVegetation and Environment New µSatellite– Technological Mission: Validation of the Israeli Hall

Effect Thruster (IHET) and demonstration of its mission enhancement capabilities:

• Orbit maintenance, • LEO to LEO orbit transfer• Enabling imaging mission in a high drag

environment.

Ben Gurion University

SOREQ



The VENµS Mission PhasesPhase 1: Scientific mission (hyperspectral imaging at 720km) – 2 yearsPhase 2: Descent to low altitude (410km) using Hall thruster (3 months)Phase 3: Demonstration of combined scientific mission and orbit maintenance at low altitude



Technological Mission Rationale

Validation and verification of the IHET systemDemonstration of imaging mission enhancement capabilities: Orbit transfer and orbit maintenance with high drag compensation requirements

Venus technological mission is an emulation of imaging mission at very low altitude

Venus technological mission is an emulation of imaging mission at very low altitude



The Orbit

Low OrbitHigh Orbit

Circular, sun-synchronous

Circular, sun-synchronousOrbit Type

410 km720 kmAltitude

2 days (31 orbits)2 days (29 orbits)Revisit Time

13 km27.5 kmSwath Width

2.9 m5.3 mResolutionConsiderableNegligibleDrag



The VENµS Technological MissionImaging orbits in Yellow, IHET orbits in blue



Phase 3 Orbit Maintenance IHET activations are performed in the IHET allocated orbits, while the scientific mission continues in the imaging allocated orbits.

IHET activation every 4th orbit, to allow orbit maintenance during solar max conditions

Imaging orbits are fixed over Israel and Europe, as planned during Phase 1

IHET allocated orbits are constant and are repeated every satellite Earth repeating cycle (31 orbits): – 4, 8, 12, 16, 20, 24, 28, 31



VM3 simulation results - 1

0 50 100 150 200 250 300 350 400 450-0.2

-0.15

-0.1

-0.05

0

0.05

0.1

0.15

0.2

0.25Changes in mean semi major axis [km]

orbits since Epoch

∆ s

emi a

fro

m n

omin

al [k

m]

[15mN 4:4:30,31] 27days

Change in Mean Semi Major Axis [km]

in 28 Days

Same as referencesatellite with no drag



ConclusionsLow altitude imaging missions can be performed with microsatellitesAdvantages: resolution, costDisadvantages: coverage, communication rangeThe configuration should be properly designed: low cross section, proper dynamic stability, adequate solar panel areaElectrical propulsion provides long-term orbit maintenance, without compromising the missionTest bed – Venus (2009). The technological mission will demonstrate the feasibility of using HET for orbit control and of carrying out undisturbed imaging mission, similar to what should be required from a microsatellite at very low altitude

microsatellites at very low altitude

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