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Tethers In Space Handbook

Edited by

M.L. Cosmo and E.C. Lorenzini

Smithsonian Astrophysical Observatory

for

NASA Marshall Space Flight Center

Grant NAG8-1160 monitored by C.C. Rupp

M.L. Cosmo and E.C. Lorenzini, Principal Investigators

Third Edition

December 1997

The Smithsonian Astrophysical Observatory

is a member of the

Harvard-Smithsonian Center for Astrophysics

Front Cover: (left) Photo of TSS-1 taken from the Shuttle cargo bay, 1992;

(right) Photo of SEDS-2 in orbit taken from the ground, 1994.

1.2 The Small Expendable Deployer System (SEDS): SEDS-1 and SEDS-2

Missions

The SEDS project started as a Small Business Innovative Research contract awarded t o

Joe Carroll by NASA MSFC. SEDS hardware proved to  be able to succesfully deploy  a  20  km
tether in space. Both flights of SEDS-1 (March 29, 1993)  and SEDS-2 (March 9, 1994) flew
as secondary payloads on Delta II launches of GPS satellites.  After  the  third  stage separation
the end-mass was deployed from  the  second stage. SEDS-1 demonstrated the capability of
deorbiting a 25 kg payload from LEO.  SEDS-2, on  the other end, demonstrated the  use of a
closed loop control law to deploy a tethered payload along the local vertical.

SEDS‘ hardware, as shown in figure 1.9, consists of a deployer, brake/cutter and

electronics box. All the components that are in contact with the tether, except for the  brake
post,  are coated with teflon. The deployer consists of baseplate, core, tether  and canister.
The tether is wound around the  core. In addition there  are  three  Light  Emitting  Diodes
(LED). Two of the  LED‘s are used to  count the turns of deployed tether, while the third is
used to  check when the tether is almost completely unwound. The canister provides a
protective cover for the tether and restrains it during deployment. The tether material is
SPECTRA-1000.

Figure 1.9 SEDS and Endmass on the Delta Second Stage

The brake/cutter components are: brake post, stepper motor, tensiometer, temperature

sensor, pyro cutter, exit guide. The tether post is coated with hard anodize. The stepper
motor is used to wrap or unwrap the tether to vary the  deployment  tension  and the  resulting
deployment velocity. The brake mechanism is a friction  multiplier and the  multiplier
function is proportional to the friction surface area between the tether  and brake post.  SEDS
functional diagram is shown in figure 1.10.

Figure 1.10 SEDS Functional Diagram

The main differences between SEDS 1 and SEDS-2 are shown in table 1. SEDS-2 closed

loop was implemented by deploying the tether according to a pre-mission profile. The
deployment control logic acted on the brake mechanism by increasing or decreasing the
deployment velocity to follow the profile and bring the payload at the end of the tether
deployment to a smooth stop along the local vertical.

Table 1. Main differences between SEDS-1 and SEDS-2

Tether Cutter Pyrotechnics

Active

Inactive

Control Law

Open Loop

Closed Loop

Tether Solder Lumps

Study Tension Pulses

None

Tether Fabrication

Tether Application

Cortland/Hughes

Mission Initiation

Prior to Depletion Burn

After Depletion Burn

Brake Usage

Minor

Significant after 1 Km

Tether Stabilization

None

Yes

SEDS-1

SEDS-2

The end-mass payload (EMP) was developed by NASA LaRC in order to monitor the

dynamics of a tethered susbsatellite. EMP consisted of three primary science sensors: a three-
axis  accelerometer, a three axis tensiometer  and a three  axis magnetometer. The EMP
measured 40.6X30.5X20.3 cm and weighted about  26  kg.  The  end-mass  was completely
autonomous and carried its own battery,  electronics,  computer  and S-band telemetry system.
As schematic of EMP is shown in fig. 1.11. The three axis tensiometer was also developed a t
NASA LaRC.

SEDS-1 mission objectives were to demonstrate that SEDS hardware could be used t o

deploy a paylod at the end of a 20 km-long tether and study its reentry after the tether was
cut. The orbit chosen had an inclination of 34 degrees and a perigee altitude of 190 km and

Figure 1.11 Schematic of SEDS EMP

an apogee altitude of 720 km. The EMP transmitted over 7900 seconds of data

before burning into the  atmosphere  (1Hz  sampling rate for the magnetometer  and 8 Hz for
the tensiometers and accelerometers). As predicted, SEDS-1 reentry was off the coast  of
Mexico (see fig. 1.12a). NASA stationed personnel at Cabo San Lucas,  Puerto Vallarta and
Manzanillo to make photographic and video observations.  The Puerto Vallarta site was able
to obtain observational data as shown in figure 1.12b

Figure 1.12a SEDS-1 EMP reentry
trajectory

Figure 1.12b Observational Data of SEDS-
1 reentry

SEDS-2 mission objectives were to demonstrate the feasibility of deploying a payload

with a closed-loop control  law (i.e. a predetermined trajectory)  and bring it to  a small final
angle (<10 degrees) along the local vertical. A secondary objective was to study the long term
evolution of a tethered system. The orbit this time was chosen to  be circular with an altitude
of about 350 km. The SEDS-2 tether was allegedly cut by a micrometeroid or debris after five
days.  The EMP transmitted over 39,000 seconds of  data  before  the battery died (1 Hz
sampling rate for all the three primary science sensors).

SEDS-1 and SEDS-2 Flight Data

SEDS data base is available through anonymous ftp at the node optimu@gsfc.nasa.gov

(128.183.76.209) SEDS1 data are in the  subdirectory /pub/projects/tether/SEDSMission1 and
SEDS-2  data are in the directory /pub/projects/tether/SEDSMission2. Each directory is
organized in different subdirectories with deployer  data,  EMP  data,  radar, etc..  Each content
of a directory is described in a read.me file.

SEDS-1

The turn counter data are shown in Figure 1.13a, the tension at the deployer is shown in

figure 1.13b and the tether rate in 1.13c. In order to compute the tether length and its rate,
the turns had to be mapped and converted into deployed length. Note that the velocity at the
end of the deployment was about 7 m/s explaining the huge jump in tension and the
consequent rebounds.

Figure 1.13a. SEDS Deployer Turns

Counts

Figure 1.13b. SEDS Deployer Tension

Figure 1.13c. SEDS Deployer 10-sec Average Length Rate

The magnetometer and tension moduli at the EMP are shown in figures 1.13d and

1.13e, respectively. Note that the magnetometer was affected by a bias estimated  to  be 3065
nT, -3355 nT and -4188 nT on the x, y and z axes, respectively.   Procedures on  the  data
calibration and validation are given at the ftp site as well as are described in  several  papers
presented at the Washington Conference.

Figure 1.13d. EMP Magnetometer Modulus

Figure 1.13e. EMP Tension Modulus

SEDS-2

The tether deployment rate and the tension at the deployer are shown in figure 1.14a,

and 1.14b,  respectively.  The  deployment  law was so effective that the final tether rate was
about 2 cm/s. As computed by the modulus of the EMP tension, shown in fig 1.14c,  the  final
libration was about 4 degrees, and it was confirmed also by the radar tracking. Even in SEDS-2
the magnetometer signal was affected by a  bias anomaly that was estimated to be -1128 nT,
1312 nT, and 2644 nT on the x,y and z axes, respectively.

Figure 1.14a. Tether Rate

Figure 1.14b. Tension at Deployer

Figure 1.14c. EMP Tension Modulus

Contacts for the SEDS Project:

•

J.Harrison , H.Frayne Smith, K.Mowery, C.C. Rupp - NASA/MSFC

•

J.Carroll - Tether Applications

•

J. Glaese - Control Dynamics

•

M.L. Cosmo, E.C.Lorenzini, G.E. Gullahorn -SAO

•

T.Finley ,R.Rhew, J.Stadler - NASA/LaRC

•

W.Webster - NASA/GSFC