Holleman, E. C.

English

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261
UTILIZATION OF THE PILOT DURING BOOST PHASE
OF THE STEP I MISSION
By Euclid C. Holleman
Flight Research Center
SUmmARY
Some of the capabilities of the human pilot for controlling the
Step I Dyna-Soar launch have been assessed by requiring the pilot to
control the simulated launch. The piloting task was well within the
capability of the human pilot. With only rudimentary presentation, the
pilot could control the launch to within acceptable limits of the desired
velocity and altitude. As the primary controller of the launch, it is
believed that the pilot can add materially to the reliability and flexi-
bility of the launch.
INTROIYJCTION
The role of the pilot in the launch of a multistaged vehicle with
orbital capability has been examined extensively during the past year
(for example, ref. I). Generally, these studies used launch simulations
in which the pilot, presentation, controller, and analog computer formed
a closed-loop system. In one study the effects of the launch-acceleration
environment on the performance of the pilot was investigated, and the
human centrifuge was used to close the launch-acceleration loops in
normal and longitudinal acceleration. The results of these studies were
generally encouraging and showed that the use of the pilot as the primary
controller of the launch of multistaged vehicles holds promise.
It is the purpose of this paper to indicate some of the capabilities
of the pilot for controlling the Step I Dyna-Soar launch based on a
fixed-base simulation program and on the results of previous investiga-
tions at the Flight Research Center.
https://ntrs.nasa.gov/search.jsp?R=19720063149 2020-03-17T06:27:30+00:00Z
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ax
g
h
2_h
Iy
q
R
S
V
AV
7
A7
5h
5n
_n
SYMBOLS
longitudinal acceleration_ g units
acceleration due to gravity, ft/sec 2
altitude, ft
altitude error, ft
vehicle moment of inertia in pitch, slug-ft 2
dynamic pressure, lb/sq ft
range, nautical miles
reference area, sq ft
relative velocity, ft/sec
velocity error, ft/sec
angle of attack, deg
flight-path angle, deg
flight-path error, deg
stabilizer position (X-15), deg
nozzle position (Dyna-Soar), deg
damping ratio
vehicle undamped natural frequency, radians/sec
LAUNCH SIMULATIONS
In figure i is shown the nominal Step I Dyna-Soar mission. The
two-stage launch to a velocity of 19,000 ft/sec with a range capability
for the lifting glider of 3,000 to 4,000 miles is shown. This study
investigated primarily the boost phase of the mission but did consider
briefly the effects that piloting errors at burnout would have on the
range capability of the glider.
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4
Control, presentation, and guidance similar to that which proved
satisfactory during previous launch programs was used to enable the
pilot to control the desired boost trajectory. For pilot's control, the
Flight Research Center's three-axis controller was used by virtue of its
generally satisfactory performance during previous fixed-base and centri-
fuge investigations. The previous paper by Brent Y. Creer, Harald A.
Smedal, and Rodney C. Wingrove showed that more conventional controllers,
for instance, a two-axis controller with toe pedals, would have been
satisfactory at the level of acceleration (about 5g) expected for this
vehicle. It was determined during the Flight Research Center's centri-
fuge boost program that longitudinal staging accelerations up to 9g had
very little effect on the ability of the pilot to perform the boost
control task. In fact, at this level of acceleration the pilots esti-
mated that only 50 to 40 percent of their physical effort was required for
the control task. There was some loss in peripheral vision due to the
norz_l component of the acceleration enviromnent, but actual data show
no deterioration in performance at this acceleration level. Since the
Dyna-Soar launch is not expected to require an acceleration higher than
6g, little effect of the acceleration environment on the pilot's per-
formance would be expected. However, a good support system, such as
the molded seat used during the centrifuge program (ref. 1), is vital
for the pilot's comfort and for fixing the pilot-controller position
during acceleration.
For the present study, no new presentation concepts were developed.
Rather, known required quantities were presented to the pilot on con-
ventional instruments as is shown in figure 2. Primary control quan-
tities were angle of attack, angle of sideslip, angle of bank, altitude,
and velocity. No vernier rockets were used for control of final velocity,
but a sensitive presentation of the final thousand feet per second proved
useful for indicating when to cut off thrust. Other useful quantities
were pitch attitude, pitch and yaw program errors, and remaining burning
time. A stage warning light was useful, especially for controlling
vehicles with unstable aerodynamics. A card of the desired attitude-
altitude provided alternate guidance.
RESULTS ANDDISCUSSION
Figure 3 shows a typical piloted launch from the fixed-base Step I
simulation. The performance quantities are shown in the upper half of
the figure, and the control quantities are shown in the lower half. The
control task was initiated 20 seconds after ground launch with the
vehicle at an initial angle of 87 °. In order to accelerate the 9,000-
pound glider to the desired end conditions of about h = 250,000 feet
and V = 19,000 ft/sec, two stages of about 5g each were required.
264
Typical mass and aerodynamic characteristics for the finned vehicle
were assumed. For this launch the vehicle longitudinal stability was
statically stable for the first stage and unstable for the second stage;
however, several levels of stability - both stable andunstable - were
investigated. Representative characteristics for the lateral and direc-
tional modeswere assumed, but primary emphasis was placed on the longi-
tudinal modesof motion. Titan (Lot J) missile weight and inertia
characteristics were used (table I), as were the Titan nozzle-deflection
and rate limitations.
For primary guidance, flight-path error was presented to the pilot.
This error was controlled by controlling angle of attack through nozzle
angle. Shownalso in figure 3 is the vehicle first-stage structural
limit of c_ = 3,750 considered during the study. 0nly small values
of a were required to correct flight-path error during the first stage,
but considerably higher values were required during the second stage
where aerodynamic lift was small. Of interest also was the absence of
'disturbances during staging where a limit of _q = 350 was used.
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Reference i has also shown that a control problem could exist at
staging for vehicles of this type. Figure 4 orients the assumed Dyna-
Soar vehicle aerodynamic characteristics in pitch relative to previous
investigations. The crosshatched region shows the scope of previous
investigations of static stability and damping. Included is the piloted
controllability limit for zero-time thrust delay between stages.
Indicated are points investigated in considerable detail under the
acceleration environment during the Flight Research Center's centrifuge
program and the two levels of damping at which the piloting controllability
limits were verified during closed-loop centrifuge operation. Shown also
in figure 4 are the first- and second-stage Dyna-Soar longitudinal aero-
dynamic characteristics representing the basic unaugmented configuration.
The Dyna-Soar vehicle appears to be easily controllable, but lightly
damped. With reference to figure 33 which illustrates the control task
with the basic configuration3 it can be seen that the control motions
are characterized by small precisely timed inputs. The pilots commented
that even stable static stability is not appreciated without damping.
For staging, the "fire-in-the-hole" technique (or firing of the
second stage before separation of the first stage) proposed for the
Dyna-Soar vehicle proved very beneficial during thrust delays, but
second-stage unstable aerodynamics can result in a control problem if
staging occurs at an angle of attack.
Figure 5 shows the results of an investigation of the control of
the second stage of the Dyna-Soar vehicle. Shown is the ratio of angle-
of-attack excursions to the staging angle of attack for various levels
of second-stage instability. These data indicated that for the basic
level of instability (_n 2 = -2.5 radians2/sec2), an excursion in _ of
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265
approximately 2.5 ° can be expected for each degree of staging angle Of
attack. Staging up to about 1.5 ° could be tolerated to restrict the
excursions to the assumed _q limit. However, it was relatively easy
to control staging angle of attack to low values.
In order to determine the effect of vehicle aerodynamic characteris-
tics on the pilot's performance, launches were made at several levels of
vehicle stability and damping. The performance of one pilot is summarized
in figure 6, which shows a typical launch as a function of velocity and
altitude. Also indicated in this figure is the spread in altitude and
velocity at first staging and also at the final cutoff velocity of
19,000 ft/sec. Shown in the left inset is a typical set of second-stage
end velocities and altitudes for the basic vehicle and two other levels
of vehicle stability and damping. No variation in performance with
stability or damping was indicated. However, it was indicated that the
pilot can control the final velocity and altitude for this mission with
the simple presentation used to within about 20 to 30 ft/sec in velocity
and 3,000 to 4,000 feet in altitude.
The ability of the pilot to adjust to more demanding control tasks
during the launch was investigated by unexpectedly failing augmentation
loops and guidance during the launch. The results of these simulated
emergencies are shownin the other insets as final incremental altitude
and velocity about the desired quantity. It can be seen in figure 6
that the pilot has the capability of performing this launch control task
even with limited presentation.
A heading change has been proposed during the Dyna-Soar launch to
avoid dropping the first-stage booster in a restricted area. To determine
the effect that this more complex piloting task might have on the pilot's
performance, heading changes of lO ° and 20 ° were made during the second
stage.
A comparison of the pilot's performance with and without the heading-
change task is shown in figure 7. Also shown is the variation in altitude
and velocity for the two tasks. It is apparent that the addition of
heading-change task had little effect on the ability of the pilot to con-
trol the vehicle burnout altitude and velocity. Figure 8 shows the effect
of piloting errors in velocity and heading at burnout on the range capa-
bility of the lifting glider. The crosshatched region shows the range
resulting from errors in velocity of 50 ft/sec and in heading of 2°.
It can be seenthat the expected piloting errors are insignificant com-
pared to the maneuvering envelope of the vehicle for the 19,000 ft/sec
mission.
Since the North American X-15 is a rocket-powered vehicle and is
designed to be piloted to 250,000 feet, a brief comparison will be
drawn between the piloting requirements for the X-15 design altitude
266
mission and the piloted Dyna-Soar launch. Typical launches are compared
in figure 9- Shown in this figure are the longitudinal acceleration,
velocity, altitude, angle of attack, and pilot's control position.
The launch accelerations during boost are quite similar and the
piloting tasks are similar once the X-19 is rotated to the proper
attitude angle of 31 °. The X-19 launch requires constant pitch attitude
to burnout, whereas the Dyna-Soar ideally requires constant angle of
attack (zero).
Piloting the X-l_ during the launch would serve to delineate the
piloting problems of the Dyna-Soar vehicle. Based on simulator investi-
gations of the control task and of the effects of acceleration environ-
ments, both control tasks appear to be well within the capability of the
human pilot.
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CONCLUDING REMARKS
In summary, it appears that the human pilot is capable of control-
ling the launch of the unaugmented Dyna-Soar vehicle. The launch accel-
eration environment anticipated will have a negligible effect on the
performance of the pilot. With augmented damping, some negative stabil-
ity could be controlled by the pilot. With only rudimentary presentation,
the pilot can control the vehicle to within acceptable limits of the
desired velocity and altitude. The inclusion of the turn task had
little effect on the pilot's control of final altitude and velocity.
As the primary controller of the launch, it is believed that the pilot
can add materially to the reliability and flexibility of the launch
maneuver.
REFERENCE
l,
Holleman, Euclid C., Armstrong, Neil A., and Andrews, William H.:
Utilization of the Pilot in the Launch and Injection of a Multi-
stage Orbital Vehicle. Paper No. 60-16, Inst. Aero. Sci.j
Jan. 29-27, 1960.
267
D
TABLE I.- TITAN MISSILE CHARACTERISTICS
Stage I (at launch):
Weight, lb .......................
Thrust at sea level, lb ................
Iy, slug-ft 2 ......................
Control arm, ft ....................
Burning time, sec ................
Stage II:
Weight, ib ........................
Thrust, ib .......................
Iy, slug-ft2
Control arm, ft ....................
Burning time, sec ...................
Glider :
We ight ib
Wing area, sq ft .....................
232,400
300,000
3,310,000
42
138.5
54,500
80,000
221,000
21
157-5
9,000
33O
268
TYPICAL DYNA SOAR-LAUNCHES
Figure 1
PILOT'S PANEL
----J@,Q£G°e
Figure 2
269
TYPICAL PILOTED LAUNCH
2
0 = t = I
ax, g
V, FPS
8O
y, DEG 4C
I I
3x104
i
TIME, SEC
Flgu_e 3
3xlO 5
h, ft
4OO
R
_-OO nout ca
miles
VEHICLE CONTROLLABI LITY
20
STABLE
UNSTABLE UNCONTRO
I
-200 2
DAMPING,
Figure 4
6
270
Q"E XCURSION
Q"AT STAGING
STAGING CONTROL PROBLEM
0
BASIC
I I
-2.,5 UNSTABLE -5.0
SECOND-STAGE STATIC STABILITY, uJn2
Figure 9
2OO
ALTITUDE
FT
PILOT'S CONTROL OF FINAL VELOCITY AND ALTITUDE
500x 10 ;5
IOC
0 4 8 12
V, FPS
+Ah,
FT
I I
16
I I
BASIC
4,000 J-
3,DO0_,
2,000 :
I,OOO
m
0 40
I
20 x 103
DAMPERS GUIDANCE
FAILED FAILED
_e J
4 i
jo
i " t
0_40 0 40
AV, FT/SEC
Figure 6
37Y 271
EFFECT OF HEADING-CHANGE TASK
Ah, FT
NO HEADING CHANGE
8xlO 3
4 0
O0 OO
0 0
ODOD 0
0 o---
0
0 0
o
0 0
I
-80 20 40
-4
AV, F T/SEC
HEADING CHANGE
0 0
C_
0
o 0
)0 0
0
000
- 0 (30
O0
I
0 20 40
AV, FT/SEC
Figure 7
PILOTING ACCURACY
1,000
I I
2,000 3,000 _i
Figure 8
2"(2
COMPARISON OF X-15 AND DYNA-SOAR LAUNCHES
6
ax, g :3
0
2xlO 4
V, FPS I[
0
3×10 5
m DS
..... X-15
_-;---:---:....7-, , ,
h' T l
-I01 i _ i i i i i i i r i r I
DEG vI
-_o
, "_:-,_" ,'-7--7--__, , , ,-- ........... 1
4-0 60 80 I00 120 14-0 160 180 200 220 240 260 280
TIME, SEC
Figure 9


Utilization of the pilot during boost phase of the step 1 mission

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