Flight evaluation of power, control, thermal, and pulse code modulation signal subsystems during first orbit of Nimbus C satellit

Morgan, F., Jr.

English

NASA Technical Reports Server

66SD4375 26 May 1966 
NIMBUS C 
F I R S T  ORBIT FLIGHT EVALUATION REPORT 
F .  MORGAN, JR. ,  SUPERVISORY ENGINEER, 
FLIGHT ANALYSIS 
APPROVED BY: 
M. GUNZBURG, MANAGER 
GODDARD F I E L D  OPERATIONS 
https://ntrs.nasa.gov/search.jsp?R=19680003211 2020-03-12T11:47:28+00:00Z
1'  / NIMBUS C ORBlT ONE - FLIGHT EVALUATION REPORT 
1.0 SUMMARY 
Following an uninterrupted terminal countdown, Nimbus C was launched from VAFB into 
a near circular polar orbit at 00:55:34 PDT on 15 May 1966 aboard a thrust augmented 
Thor booster with an Agena B second stage. 
Following the planned powered flight profile, the spacecraft separated from the Agena B 
approximately 57 minutes after liftoff. Contact was maintained with the spacecraft through 
the launch and initial orbit insertion phases. Real time data was received at VAFB for 
two minutes after liftoff. Forty-eight minutes after liftoff, spacecraft telemetry was 
acquired at Johannesburg, South Africa. During this pass, spacecraft booster separation 
occurred and spacecraft initial stabilization was confirmed. In addition, the magnitude 
of the clock jump, resulting from paddle unfold squib firing, was determined. This inform- 
ation provided confidence in spacecraft performance and also permitted rapid clock reset 
during the subsequent Ulaska pass. Additional spacecraft confidence performance was 
observed at Winkfield, England, which monitored the beacon telemetry signal strength. 
Following separation from the Agena at the planned pitch attitude, the control subsystem 
maintained attitude of the spacecraft, within the fine limit cycle deadband, 3 . 8 ,  3 . 6 ,  and 
7 minutes respectively in pitch, roll, and yaw. 
Available data showed that the power subsystem and power supplies of the various other 
subsystems that were turned on functioned normally throughout this early flight phase, 
The values of pressurized components remained constant. The spacecraft's PC M beacon- 
signal was satisfactorily acquired by the Ulaska tracking station at approximately 09:23:34 
GMT with loss of signal occurring approximately 19 minutes later. 
2.0 - INTRODUCTION 
The Nimbus Program is the result of a continuing NASA need for a precisely controlled, 
near-space , earth-oriented, long-life spacecraft with modularized payload interchange- 
ability for use by the scientific community. Objectives of the Nimbus C flight include 
operational utilization of the spacecraft control, clock/command, thermal, power and 
sensory subsystems . 
This report presents a discussion of the data obtained during orbit one of the current 
Nimbus C flight. Conclusions concerning any discrepencies observed in the data have 
been reserved for inclusion in the First 30-Day Flight Evaluation Report. 
2 
3 .0  LAUNCH EVENT SUMMARY 
Table 3-1 presents a summary of the significant orbit one flight events. Figure 3-1 presents 
a diagram of the launch sequence. 
TABLE 3-1. FIRST ORBTI' FLIGHT EVENT SUMMARY 
Time of Occurrence (Nominal) 
GMT) 
Liftoff 
3ocket Motor Cutoff 
3ocket Motor Ejection 
Meco 
Veco 
I'hor/ Agena Separation 
4gena First Burn Ignition 
Vose Shroud Ejection 
\gena First Burncutoff 
Agena Second Burn Ignition 
Agena Second Burn Cutoff 
Agena Pitchup Maneuver Initiation 
Agena Pitchup Maneuver Termination 
Spacecraft/ Agena Separation 
Spacecraft Paddle Unfold Initiation 
Spacecraft Pitch and Roll Control Loops Enabled 
Spacecraft Yaw Control and Solar Array Drive Loops Enabled 
Alaska Nimbus Beacon Signal Acquisition 
Alaska Nimbus Beacon Signal Loss 
07:55:34 
07:56:17 
07:56:39 
07:58:03 
07:58:12 
07:5 8:19 
07:58:47 
0 7 5 8 ~ 5 7  
08: 02: 43 
08:48:14 
0 8:4 8:2 0 
08:5 1:11 
08:5 2:3 1 
08:52:42 
08:52:45 
08:5 2:5 5 
08:53:35 
09:23:34 
09:42:34 
3 
SUN 
1. LAUNCH FROM VAFB. 6. NOSE SHROUD SEPARATION. I I. TERMINATE PITCHUP. 
2 TAT MAIN ENGINE CUTOFF 7. AGENA FIRST CUTOFF. 12. SEPARATE SPACECRAFT. 
(MECO). 8. AGENA SECOND IGNITION. 13. INITIATE YAW/ROLL 
3. TAT VERNIER CUTOFF 9. AGENA SECOND CUTOFF. ' MANEUVER 
(VECO) IO. INITIATE 60e/MlNUTE 14. TERMINATE YAW/ROU 
4. TAT/AGENA SEPARATION. PITCHUP. MANEUVER, FIRST RETRO. 
5. AQENA FIRST BURN. IS. SECOND RETRO. 
Figure 3-1. Launch Sequence Diagram 1 4 
4 . 0  INITIAL ORBITAL PARAMETERS 
Table 4-1 presents the initial orbital parameters. 
TABLE 4-1. INlpTIAL ORBITAL PARAMETERS 
Parameter 
Eccentricity 
Inclination 
Argument of Perigee 
Nodal Regression 
Period 
Ped gee 
Apogee 
Solar Incidence 
units 
- 
Degrees 
Degrees 
Deg/ Day 
Minutes 
NM 
NM 
Degrees 
Value 
0.00558 
100.311 
337.710 
1.0041 
108.06 
590.5 
636.0 
9.74 
*Source: NASA Tracking and Data Systems 
5 
5.0 POWER SUBSYSTEM 
Normal power subsystem performance was experienced throughout the entire first orbit. 
The initial current provided by the solar arrays after paddle unfold was 14 amps, one amp 
greater than normal. Initial power loss in the slip rings was approximately 15 watts. 
Figure 5 -1 shows specific power subsystem parameter histories for the first orbit which 
included powered flight, separation, initial acquisition and stabilization. A s  shown in 
Figure 5-1, the cell face of the left array was exposed to sunlight as soon as the spacecraft 
exited the umbra. This was substantiated by the sharp rise in the left array temperature 
and the sudden appearance of the paddle board voltages at that time. The right paddle 
board temperature reflects the general spacecraft warmup as it became exposed to sunlight. 
The left array cell surface was exposed to sunlight rather than the right, because of the 
9.7 degree solar incidence angle. The right array surface was therefore shielded from 
sunlight. As the spacecraft approached more northern latitudes, the cell surface of the 
arrays began to be shielded from the sun. This is apparent from the sharp initial decrease 
in paddle board voltages. These voltages again began to rise during the pitch-up maneuver 
as the cell array surfaces were once more exposed to sunlight. At  the time of paddle 
unfold, and until the solar array drive was enabled, the paddle board voltages were again 
lost. This was because the back side of the array was facing the sun. A s  soon as the 
arrays were properly pointed toward the sun, the unregulated bus current rose to 14 amps, 
and the batteries began to charge. A l l  batkries charged to identical voltages. The lowest 
level of discharge was 28.8 volts, which was associated with battery eight. 
6 
8 
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SPACECRAFT 
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SOLAR ARRAY AUX AND AVCS COMP 
DRIVE ENABLED LOADS COMMANDED ON 
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LEGEND -PADDLE BOARD VOLTS .- -.- - ,-- .-- _ _  - -- - -- UNREGULATED BUS VOLTS q 
I 
A4 4 
ULASKA ACQUIRED 
BEACON SIGNAL 
20 
I O  
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-10 0 I O  20 30 40 50 60 70 80 90 100 
TIME FROM LIFT-OFF (MINUTES) 
Figure 5-1. Power Subsystem Parameter Histories for Orbit One 
7 
6 . 0  
Event 
Spacecraft/ Booster Mechanical Separation (T ) 
0 
Solar Array Unfold Initiation 
CONTROL SUBSYSTEM 
Elapsed Time 
(Set) 
0 
2.5 
The control subsystem performed in an excellent manner, in that it acquired the earth 
reference and stabilized the spacecraft within the fine limit cycle deadbands, within 3 . 6 ,  
3.4 and 6 . 1  minutes respectively in pitch, roll and yaw after each of these loops were 
enabled. Following the period of initial acquisition, no further solenoid firings were 
observed on orbit one. The solar array drive pointed the array surface properly toward 
the sun within 5.5 minutes after the drive loop was enabled. 
Telemetered control subsystem functions verified the acquisition sequence of events which 
appears in Table 6-1. The useable pneumatic impulse available at liftoff was 715 pound- 
seconds. 
TABLE 6-1. SPACECRAFT SEPARATION AND INR'IAL 
ACTIVATION SEQUENCE OF EVENTS 
Pitch and Roll Loops Pneumatics Enabled 
Solar Array Drive and Yaw Loops Enabled 
12.5 
52.5 
I I 1 
Figure 6-1 illustrates the period of Agena pitch-up , spacecraft/ Agena separation and 
initial acquisition. The last 15 degrees remaining in the Agena pitch-up schedule was 
observed in the coarse pitch e r ror  channel and is illustrated as a rapidly decreasing pitch 
error ,  approximately 1.1 minutes, prior to separation. The Agena yaw limit cycling is 
illustrated as cyclic variations in the spacecraft roll error  prior to separation. 
When the pitch and roll control loops were enabled after separation, the positive pitch 
solenoid fired continuously for six seconds turning off as  the pitch e r ror  came within the 
8 
pneumatic deadbands. As the pitch error became positive, the negative pitch solenoid 
fired continuously for three seconds. The pitch error  overshot by five degrees causing a 
negative pitch solenoid gate to occur. Following two subsequent overshoots of 1.0 and 0.6 
degrees, the pitch e r ror  was controlled within the fine limit cycle deadbands. These 
overshoot characteristics are illustrated in a phase plane plot in Figure 6-2. 
No roll solenoid firings were required to stabilize the spacecraft on this axis. Overshoots 
in this axis are illustrated in the phase plane of Figure 6-3. 
Following enabling of the yaw loop, yaw flywheel pneumatic gating occurred seven times. 
The second gate in the series was a timed re-enable gate; a negative gate occurring 30 
seconds after the first negative gate. Yaw fine limit cycle control was achieved 6. 1 minutes 
after it was enabled, with a zero average heading er ror  resulting. No further yaw solenoid 
firings occurred prior to the Lnaska station pass. 
The slew of the array is also shown in Figure 6-1. The array drove backwards from the 
180 degree position for approximately 5.5 minutes before the SAD sun sensors became 
nulled. 
At  Thor engine ignition, sufficient vibration apparently occurred to cause the pneumatic 
regulator diaphragm to chatter, permitting high pressure gas to leak into the low pressure 
manifold; this is shown in Figure 6-4 as the pressure going out-of-band high for the 
sensor monitoring this pressure. 
The low pressure control solenoids in the manifold are  forced open at 90 psi. Had the 
regulator not become reseated, a noticeable drop in "tank pressure" (gas locked up in the 
line down stream of the tank solenoid and up stream of the regulator) would have resulted. 
However, this did not occur, therefore, the regulator probably became reseated shortly 
after liftoff (a period of high vibration levels). The manifold pressure returned to normal 
at the time the pitch and roll loops were enabled. 
9 
30 
a 
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-10 
a 
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LL 
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0 
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-30 
200 
W 
s- 
-200 
2000 
-2000 
PITCH AND ROLL 
CONTROL LOOPS 
ROLL FLYWHEEL SPEED 
YAW FLYWHEEL SPEED 
ATIVE YAW SOLENOID GAT 
I l l r l r r l l l r l l l r r l l l r l  
-2  -I 0 I 2 3 4 5 6 7 8 
TIME FROM SEPARATION (MINUTES) 
Figure 6-1. Control Subsystem Parameter Histories of Spacecraft 
Initial Attitude Reference Acquisition 
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CD 6 J  m t t 13 CI) 
13 
Figure 6-5 presents the launch and initial control subsystem housing panels and thermal 
control shutters re spec tive ly . 
On the launch pad, the sensory ring was maintained at approximately 17OC by forced cool 
air. As  can be seen in the figure, the control housing panels were 5 to 10 C hotter 
than the ring. As a result, the number one and two thermal control shutters were 8 and 
6 degrees open respectively, which was normal for their associated panel temperatures. 
0 
During the period between MECO and Agena first burn cutoff, shutter No. 2 seemed to be 
driven further open, up to 53 degrees prior to first burn cutoff . Following engine cutoff, 
the shutter returned to the 28 degrees open position and began cycling about this level in 
response to ita associated panel temperature. 
Shutter No. 1 remained at the proper position until Agena engine second burn ignition, at 
which time it also went to the 28 degrees open position. This shutter subsequently began 
responding to its associated panel temperature as it began to warm up. The 28 degree 
open position finally assumed by the shutters was approximately 20 degrees more than 
normally would be expected when compared to the associated panel temperatures. 
Figure 6-5 illustrates the housing panels temperature responses to the spacecraft exiting 
the umbra, the pitch-up maneilver and crossing of the ecliptic plane. 
The control subsystem component temperatures were maintained within their specified 
ranges during the launch and orbit injection phases. The gyro fluid temperature was 
precisely controlled at 74OC while the yaw and roll flywheels were maintained at 30 and 
31OC respectively. Figure 6-6 shows the initial cool down of the IR scanner sink temper- 
atures and the solar array drive motor temperature rise during the period of initial sun 
acquisition. 
14 
FIRST BURN 
CUTOFF 
AGENA SECOND 
BURN IGNITION 
- 1  
1 PANEL N 0 . 5  8 SHUTTER NO. 2 I I I 1 I I I 
I 
I 
I --- 1 LEGEND 
PANEL NO. 5 - --- SHUTTER NO. 2 POSITION 
PANEL NO. 4 TEMPERATURE ---- PANEL NO. 6 TEMPERATURE 
20 0 20 40 60 80 100 120 
TIME FROM LIFT-OFF (MINUTES) 
Figure 6-5. Histories, Launch and Orbit Insertion 
15 
DATA 
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7.0 THERMAL SUBSYSTEM 
On the launch stand, ground cooling air was wed  to supercool the sensory ring and ita 
associated components at approximately 15 to 17OC. This was done to ensure that the 
thermal control shutters would remain closed throughout the launch and injection flight 
phase, thereby avoiding any damage due to mechanical interference during the various 
separation events. 
After  liftoff, a gradual cooling down was experienced until the spacecraft emerged from 
the umbra. The cooldown trend was then reversed and general heating was observed. The 
average ring temperature for the first orbit was 15.5'C and the H-frame center section 
structure averaged about 16OC. The extremes for the average sensory ring component 
temperatures for the first orbit were 12.5 to 22.3OC. 
17 
8.0 PCM AND CLOCK/COMMAND SUBSYSTEMS 
The PCM subsystem delivered 98.5% good data in the first interrogation playback. Trans- 
mitter power output was 0.42w, operating on beacon No. 2. The first real-time telemetry 
data frame was processed at an antenna elevation angle of 5.7 degrees. Preflight estimates 
based on a 0.3~ transmitter expected approximately 14 degrees antenna elevation before 
data would have become usable. The PCM recorder maintained a pressure of 13.7 psi at a 
temperature of 19'~. 
The clock subsystem provided proper reference frequencies and phasing for other space- 
craft subsystems during the first orbit. The command section was enabled properly at 
separation and received, loaded, and executed encoded commands properly during the 
Alaska station pass. Crystal oven temperature was maintained at 59.5'C. 
During the separation and unfold sequence, available data indicated that the following 
anomalies occurred: 
a. 
b. 
A clock jump of 512 days, 60 hours, 10 minutes, 17 seconds. 
A PCM multicoder jump of one word. This resulted in improper computer sync 
with PCM data until the start of the next main frame, resulting in loss of meaning- 
ful data during unfold. 
Regulated bus voltage is suspected to have dropped initially to about 21.8 volts 
then surged to 28.1 volts. Changes in the IRIS position monitor TMV levels 
during this period, considering its relationship to the regulated bus voltage, was 
used to  derive this variability in regulated voltage. 
c. 
18 
9.0 
L 
I - 
Suljsystem Pressure. (Psi) Temperature, OC 
AVCS 17.3 16.2 
HRIR 16.4 17.4 
MRIR 14.9 19.0 
SENSORY SUBSYSTEMS 
With the exception of the MRIR record motor and HRIR and MRIR radiometer motors, no 
sensory components were operating during launch. Recorder temperatures and pressures 
remained constant throughout. (See Table 9-1). The HRIR detector cell temperature 
showed slight cooling during the initial satellite night period, returning to its original 
temperature after satellite sunrise, then cooling rapidly again after separation to -55 OC 
at the time of A-stored playback at  Ulaska. The temperature eventually stabilized at 
-75OC . 
The MRIR record motor current indication dropped from 3.8 to 3.4 TMV at the time of 
unfold squib firing. This motor is normally in continuous operation whether o r  not MRIR 
data is being recorded. 
19 


Nimbus C first orbit flight evaluation report

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