NASA CONTRACTOR REPORT 177394 

, 

(PASA-€E-177394) SSULP CP lEOr SYiTEflS N87- 193C 8 

CCPICEE'IS FCfi A 'll'1AN ATMCSEhEEIC EBCBE 

Einal Technical &€Fort (HG~~EI Aircrdft Cc.) 

196 E CSCL 22a Unclas 

G3/d8 43770 \ 

Study of New Systems Concepts for a Titan Atmospheric Probe 

CONTRACT NASZ- 12049 

April 1986 

Doug Bernard 

Todd Citron 

Bob Drean 

Scott Lewis 

Martin Lo 

John McCarthy 

Robert Soderblom 

Dave Steffy 

Tina Vargas 

Marty Wolf f 
*' 

NASA CONTRACTOR REPORT 177394 

Study of New Systems Concepts for A Titan Atmospheric Probe 

Doug Bernard 

Todd Citron 

Bob Drean 

Scott Lewis 

Martin Lo 

John McCathy 

Robert Soderblom 

Dave Steffy 

Tina Vargas 

Marty Wolf f 

Hughes Aircraft Company 

Space and Communications Group 

El Segundo, California 

Prepared for 

Ames Research Center 

Under Contract NAS2-12049 

April 1986 

National Aeronautics and 

Space Administration 

Ames Research Center 

Moffett Field, California 94035 
t b 

Table of Contents 

Section 

1. Summary and Introduction 

2. Mission Analysis 

2.1 Carrier Deflection/Coast 

2.2 Entry 

2.3 Descent 

2.4 Relay Communications 

2.5 Surf ace Survival 

3. Deceleration Module 

3.1 Decelerator Configuration 

3.2 Thermodynamics 

3.3 Aerodynamics 

3.4 Parachute 

4. Descent Module 

4.1 Configuration 

4.2 Structure 

4.3 Thermal Control 

4.4 Power Subysystems 

4.5 Communicatios Subsystem 

4.6 Command Data Subsystem 

5. Carrier Mounted Hardware 

5.1 Separation 

5.2 Receive Antenna 

5.3 Receiver 

6. Conclusion 

6.1 Technology Readiness 

APPENDIX A Symbols and Abbreviations 

References 

Abstract 

Page 

1 

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14 

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54 

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106 

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133 

137 

152 

152 

171 

173 

189 

190 

A- 1 

A- 3 

A-4 
1. SUMMARY AND INTRODUCTION 

This final report documents results of a study of new systems concepts €or 

a Titan Probe performed for NASA Ames Research Center under contract NAS .2-12049. 

This effort extends the work previously performed (Hughes, 1978) where a Titan 

Probe was conceived based on minimum deviation from Galileo Probe hardware. 

Though not bounded by existing hardware, this conceptual study was constrained 

to readily available and low risk hardware projected for the late 1980's time 

frame. 

as a point of departure for the present study. 

The Titan atmospheric probe concept described in Swenson, 1984 was used 

This study also reflects the latest thoughts of planetary scientists relative 

to investigation goals which are somewhat different from the previous studies. 

Foremost of these is the desire to sample the organic haze layer which surrounds 

Titan at high altitude, up to 200 km above the surface. 

this altitude requires a low ballistic coefficient. 

light weight Probe design developed in this study provides the desired atmospheric 

s amp li ng . 

Subsonic gas sampling at 

The blunt, large base diameter, 

A second new requirement is lengthened descent time of 2 to 4 hours, to 

provide adequate gas sampling. This defines principally, the parachute design, 

the descent thermal control design, and the power source capacity. 

A third new requirement is the ability to image the surface before impact. 

This leads to a data rate one or two orders of magnitude beyond previous atmospheric 

probes. 

cations relay enables the necessary data transmission. 

However, the relatively close Carrier flyby altitude during the communiA fourth new requirement is surface operation. Because of the major effect 

on probe cost of a hard requirement for surface impact survival and operation under 

a wide range of surface conditions, the current study only describes the necessary 

changes; no attempt was made to alter the baseline design to enhance chances of 

surface survival. 

Table 1.-1 lists the key tradeoffs performed during this study and described 

in detail in this final report. 

Figure 1.-1 shows the baseline Probe configuration derived from the study 

tradeoffs. The Probe basically consists of two major parts, plus separation 

and rsdio relay hardware left with the Mariner Mk I1 Carrier spacecraft after 

Probe release. 

taining all instrumentation and support subsystems, attaches directly to a solid 

beryllium heat-sink nosecap which provides thermal protection during atmospheric 

entry. The Probe attaches to the Carrier through three attachment points within 

the beryllium nosecap. 

vide the necessary low entry ballistic coefficient (high drag area). 

uses an ablator entry thermal protection system. 

The central probe element, the 1 m diameter Descent Module conThe outer decelerator skirt surrounds the nosecap to proThe skirt 

Figure 1.-2 illustrates the mounting of the baseline Probe on the Carrier. 

Minimum mass and cost result from the nose inward mounting. The fixed decelerator 

skirt has less risk than a deployable design and fits within the Shuttle envelope. 

1 
Figure 1.-3 shows the sequence of events during Probe entry and descent. 

The Probe first encounters the sensible atmosphere at an altitude of about 

700 km. 

of about 190 km where the velocity drops below sonic speed. At this point, 

the Descent Module separates from the decelerator skirt by differential drag. 

As soon as the skirt is clear, a lanyard from the skirt deploys a parachute 

from the rear of the Descent Module. 

instruments deploy, the communications link is established with the Carrier 

spacecraft, and the Probe begins the nominal 3 hour descent to the surface. 

The Probe undergoes supersonic aerodynamic deceleration to an altitude 

After this, the optical heads of certain 

The individual sections of this report give mass breakdowns of each 

Probe subsystem or major element. Table 1.-2 summarizes the Probe mass. As 

shown, the total of 154.8 KG separated mass is over one third decelerator. 

Carrier mounted hardware comprise 34.7 Kg with a total mass including 

contingency of 208.5 Kg. 

according to its various subsystems. 

Table 1.-3 further breaks down the descent module mass 

The next section of this final report describes the mission analysis 

performed to determine compliance with the high altitude sampling and descent 

time requirements. 

geometry to enable high rate surface imaging and defines some of the 

considerations of surf ace survival. 

It also describes the selection of Probe/Carrier relay 

Section 3 describes the derivation of a baseline decelerator design. 

Entry thermodynamics analysis determines the decelerator materials and thickness 

and aerodynamic stability constrains the shape. 

the parachute is included in the decelerator section. 

For purposes of this study, 

Section 4 derives the baseline Descent Module design. Tradeoffs result in 

an integrated descent module/decelerator nose with aft separation of the skirt 

by differential drag. Some subsystems retain Galileo technology. The most 

significant advancement from previous probe technology is the Command/Data subsystem reflecting the much higher data rate and increased emphasis on redundancy. 

Section 5 describes the elements of the Probe System left on the Carrier 

after separation. These include the separaration hardware, a simple three 

spring simultaneous spin/eject mechanism, and the radio relay hardware, a 

BPSK receiver and antenna. 

Finally, section 6 summarizes the study results. 

The following individuals are responsible for the preparation of this 

final report. We also acknowledge the study contributions of our subsystem 

specialists, of our subcontractcr, general Electric Re-entry Systems 

Organization, under the management of Dwight Florence, and of the NASA contract 

monitor for this study, Byron L. Swenson of NASA Ames Research Center, Moffett 

Field, CA. Finally, we acknowledge the support of our secretary, Donna Williams, 

who typed and assembled this report. 

Doug Bernard 

Todd Citron 

Bob Drean 

Scott Lewis 

Martin Lo 

John McCarthy 

Robert Soderblom 

Dave Steffy 

Tina Vargas 

Marty Wolf f 

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2. MISSION ANALYSIS 

Table 2.-1 lists the major assumptions made in performing the mission 

tradeoffs described in the following sections. The approach velocity of 

6.5 km/sec assumes the Carrier is in orbit about Saturn prior to probe release. 

There is considerable flexibility in Probe targeting and Carrier flyby 

geometry. Figures 2.-1 and -2 show the overallProbe/Carrier geometry for 

a set of representative conditions. Tradeoffs have demonstrated small 

sensitivity to these characteristics. Two model atmospheres were used for 

this analysis, a model by Hunten, 1982 and a model by Lindal, 1983. The 

Hunten model atmosphere was imposed as an upper density bound. Figure 2.-3 

shows the slight offset for the Lindal model. 

It should be noted that for this preliminary analysis, the effects of 

the small rotation of Titan (approximately 1 deg/hr) were neglected for 

simplicity. This assumption was made because the precise azimuth of the 

entry trajectory relative to the pole of Titan is a function of the mission 

opportunity, the orbit inclination about Saturn, and other Saturn orbit 

mission characteristics which have not been chosen as yet. The final entry 

trajectory will have to be appropriately biased to account for this rotation, 

but for the purposes of this preliminary analysis of Titan probe characteristics, this assumption is sufficiently accurate. Finally, note that no 

atmospheric winds were considered in the bulk of the present analysis. This 

assumption was made because at the start of the study, a wind profile for the 

circulation of the Titan atmosphere had not been defined. Subsequently, a 

rather strong zonal wind pattern was defined based upon the modelling of 

Flasar, 1981. The rather substantial effects of that wind profile on the 

communications link are discussed in Section 2.4. 

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2.1 Carrier Deflection/Coast 

The Probe design is relatively insensitive to the time of separation 

compared to entry. 

its stability is not a critical factor since the Probe sequence resets based 

on direct measurement of entry by acceleration switches. 

does have an effect on the Carrier since the Carrier retargeting maneuver 

(to position the Carrier for flyby and communications relay) is larger for 

later separation, (see figure 2.1-1). 

however, as there must be time for orbit determination after the final Carrier 

orbit correction maneuver at about E-20 days. The figure shows a modest 

penalty for delaying the Carrier an additional 20 to 40 minutes over the 

nominal 3 hour delay. 

The coast timer imposes a very small battery drain and 

The separation time 

Separation can not be too early, 

Carrier phasing is defined in section 2.4 

Probe entry angle is not critically sensitive to Carrier orbit determination. 

Figure 2.1-2 shows a target plane error of 100 KM results in less than 2O path 

angle error for steep entry. 

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2.2 Entry 

Although steep entry path angles impose a greater deceleration load 

than shallow entry path angles, figure 2.2-1 indicates that the worst 

case deceleration load at a -90° path angle of 20 gE is still very benign 

and comparable to the launch load environment. Figures 2.2-2 and -3 show the 

time variation of several key trajectory parameters during entry for 

entry angles of -4OO and -90°, respectively. 

The entry path angle also has an influence upon the altitude where 

subsonic speeds are reached. Figure 2.2-4 shows that the baseline probe 

design with a ballistic coefficient of 14.49 Kg/m2 a Mach number of unity 

at 190 KM. It is not easy to increase the altitude. As shown, a 10% 

larger probe diameter (as could be realized through a complex deployment) 

only adds an additional 10 KM to the sonic altitude. Shallow path angles 

offer somewhat greater performance. At a practical communication limit 

(i.e. 70° probe aspect angle) the resultant -35O entry path angle adds 35 

KM to the sonic altitude. 

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2.3 Descent 

The diameter of the parachute is the principal parameter to adjust the 

descent time from parachute deployment to impact. Figure 2.3-1 shows that a 

3.13 M diameter parachute gives the desired nominal 3 hour descent. Figure 

2.3-2 shows the corresponding descent velocity profiles. Small parachutes 

(e.g. 2m diameter), result in an initial speed-up since they have less drag 

than the decelerator. The descent time can not be exactly controlled due to 

various uncertainties. The total uncertainty is important to the relay 

communication analysis as described in the next section. Table 2.3-1 shows 

how errors combine to a - +8.1% RSS or a sum of - +13.3%. 

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2.4 Relay Communications 

Probe communication to the Carrier has two basic phases. The initial 

phase is characterized by a low rate of 100 bps to support basic atmospheric 

structure and composition measurements. This phase does not critically size 

the link. The second phase requires a much higher rate of 10,000 bps to 

satisfy surface imaging. The mission parameters affecting the link include 

the Probe entry trajectory and the Carrier flyby trajectory (i.e. periapsis 

altitude and phasing) which in turn determine Probe aspect angle, Carrier 

aspect angle, and doppler frequency shift and rate. Hardware variables 

including transmitter power, probe antenna size, receive antenna size and 

pointing mechanization, and modulation type combine with the mission parameters to determine link margin, Eb/No as shown schematically in figure 

2 -4-1. 

Figure 2.4-3 defines the communications aspect angles. Since the exact 

mission date is not yet defined, the Carrier attitude (Earth oriented) is not 

known. Therefore, Carrier aspect angles refer to the direction of approach 

vel0 ci t y . 

Figure 2.4-4 further defines the Carrier communications angles. The 

key performance factor depends on the offset of the receive antenna boresight 

(look angle) and Probe line of sight (Carrier aspect angle). This offset 

(communications angle) is nominally zero (with a 2O pointing error) if the 

receive antenna is mechanized to track the probe. 

For purposes of this analysis, the mission parameters given in Table 2.-1 

will be assumed together with the following constraints 

2.5W transmitter power 

S-band frequency (2295 MHz) 

10,000 bps 

Figure 2.4-5 depicts the time variation of Probe aspect angle during descent 

for several entry angles. As shown this angle begins at a nominal value which is 

offset most -from local vertical, for shallow path angles. As the Carrier flies 

overhead, this angle quickly goes through zero and then exceeds a value of 90° as 

the Carrier passes over the local horizon. By arranging for Probe impact some 30 

ninutes or more before Carrier periapsis, the Probe antenna beamwidth depends 

only on the entry path angle. 

dently of any communications performance considerations, two representative entry 

conditions and associated Probe antenna beamwidths have been studied as indicated 

in Table 2.6-1. 

Since entry path angle may be selected indepenThe first constraint imposed on the communications relay is that all lines 

of sight fall within antenna 4 dB cutoff points. Table 2.4-2 summarizes the key 

characteristics of all considered Probe and Carrier antennas. Figure 2.4-6 shows 

the effect of the cutoff €or the two representative missions and as a function of 

Carrier periapsis attitude. The lower limit periapsis altitude 1000 km is 

selected since it allows the latest communications before cutoff. 

A second constraint involves doppler shift or rate, the exact limit depending on the modulation technique (see section 5.3). For a BPSK receiver, the 

26 
doppler rate must be less than 2 Hz/sec (2nd order tracking loop) and 

110 Hz/sec (3rd order loop). As shown in figure 2.4-7, this also requires 

targeting such that the Probe impact is at least 30 minutes before Carrier 

periapsis. An FSK receiver is sensitive to doppler shift. Figure 2.4-8 

shows that a typical 5 kHz filter also is consistent with about 30 minutes 

of phasing. 

A third constraint comes from the receive antenna geometry if the 

antenna is not free to track the predicted Probe location. If the receive 

antenna is positioned for no nominal offset at the start of Probe descent, 

the 4 dB cutoff requires a minimum phasing limit of 30 to 45 minutes as 

shown in figure 2.4-9. 

Table 2.4-3 summarizes the phasing constraints resulting from the above 

considerations. As shown, the baseline design (i.e. steep probe entry, BPSK 

2nd order loop receiver, and a pointed receive antenna) requires a minimum 

phasing of 40 minutes. 

The final measure of performance is the link margin which varies with 

range and communications angles. Figure 2.4-10 shows for the shallow entry 

case (i.e. case 1) a short time where 10,000 bps can be supported. Figure 

2.4-11 shows the much longer high rate communications time is available for 

the steep entry case. 

This characteristic time provides a bench mark to compare the various 

factors in the communications tradeoff. Table 2.4-4 shows the maximum time 

where 10,000 bps can be supported (considering all the constraints and link 

margin) as a function of Probe entry angle and antenna diameter (Case 1 

and 2), pointed or fixed Carrier antenna, Carrier antenna size, and Probe 

communications modulation technique. As noted the minimum desired time is 

70 minutes to account for timing uncertainty (see section 2.3) and a minimum 

10 KM descent at 10,000 bps. The 145 minutes provided by the baseline 

design has ample margin. 

in the wind environment (see next section). 

This margin is required to account for uncertainties 

Wind Effects 

After completion of the preceding mission analysis a wind model has been 

suggested which significantly alters the quantitative results, but does not 

significantly alter any conclusions. To ensure the fundamental validity.of the 

analyses, a brief examination was made of the impact of the wind model. Figure 

2.4-12 shows the wind velocities experienced by the Probe as a function of time 

for the 3 hour descent. For this preliminary study, an analysis has been made 

of the effect of the Probe horizontal motion due to this wind on the Probe 

position and the increased Probe communication angle due to the rotation of 

Titan about its axis. The modeling did not include any attitude offset due to 

wind shear. Figure 2.4-13 shows the Probe longitude rotatiton due to the wind. 

The contribution of Titan rotation is also shown. The wind model results in a 

296 KM shift of probe impact site compared to no wind. This motion results in 

change in the communication angle profile as shown in figures 2.4-14 through -16. 

The Probe aspect angle exceeds the +11.40° cutoff of the previous baseline 

thereby requiring a broader Probe aitenna beamwidth. 

matches the new geometry. Table 2.4-5 lists the appropriate Prob; antenna 

characteristics. 

As shown a +13.7O cutoff 

27 
Figure 2.4-17 shows the effect the hroader Probe antenna has if there 

is no wind. The time where the link margin exceeds the 4.8 dB BPSK threshold 

decreases from 145 minutes to 125 minutes. Figure 2.4-18 shows the performance with wind. Whereas the previous Probe antenna would have a significant 

data outage where the communications angle exceeds the 4 dB cutoff, the new 

antenna would support the 100 bps and 10,000 bps phases with the later high 

rate phase allowed for 87 minutes. 

In the first analysis, Titan rotation is ignored. Now, coupled with 

the wind, the rotation mst be considered. Table 2.4-6 lists the various 

performance factors. With rotation, the high rate time becomes 75 minutes, 

just meeting the 70 minute target time (fig. 2.4-19). 

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53 
2.5 Surf ace Survival 

Approach 

For the purposes of this study, the following approach was adapted to 

determine requirements for surface survival and operation for a period of 

10-20 minutes. First, the baseline design was developed without imposing 

surface survival requirements. Second, an assessment of the survival 

potential of the baseline design was made and problem areas were identified. 

Finally, design modifications were conceived to enhance the survival potential 

of the Probe. Since the state of the surface is unknown, both liquid and 

solid surfaces have been considered. 

Specific concerns fall into three classes. The first class includes 

questions about the Probe's abiliLy ir, survive the shock of impact. The second 

class addresses the problems of operating on the surface, including thermal 

control, power availability, and probe buoyancy. The final class includes the 

various problems associated with maintaining communications with the Carrier. 

Surface Conditions 

At present, there are conflicting theories of the nature of Titan's 

surface. If solid, the surface is expected to consist of a mix of ices and 

clathrates. If liquid, present theories predict a global ocean of liquid 

ethane mixed with some liquid methane and liquid nitrogen (Lunine 1983). 

The estimated density of the ethane ocean is 0.61 g ~m'~, and the estimated 

depth is 1 km. 

Based on Voyager measurements, a surface temperature of approximately 

According ta the Conrath model (Conrath 1985), 

95K, surface pressure of 1.5 to 1.6 bars, and composition primarily of diatomic 

nitrogen (N2) is expected. 

there is no wind at the surface although wind profiles (described in section 

2.4) indicate the possibility of non zero surface wind. 

Science requirements for surface operation are not specified at present. 

There is some interest in obtaining samples of the surface for chemical 

analysis. 

the probe to provide an easily interpreted deceleration signal at impact. . 

In addition, it may be worthwhile to deploy a spike forward of 

For the purposes of this study, no special science requirements have 

been assumed. However, some of the difficulties of using the existing 

science complement on the surface have been identified. 

Surviving Impact 

As shown in Figure 2.5-1, the probe's impact velocfty (without winds) 

varies from 5.7 m sec'l (for a two hour descent) to 2.9 m sec'l (for a four 

hour descent). To give an intuitive feel, Figure 2.5-2 expresses these impact 

velocities in terms of the height from which the probe would have to be 

dropped on Earth to reach the same speed. 

height is 1.7m. 

The worst case equivalent drop 

54 
The baseline design has a special beryllium nosecap about 1 cm thick. 

Dropped from a height of 1.7 m, this nose is unlikely to crush. However, the 

impact load is likely to exceed the Probe structure capabilities since it is 

sized for relatively low loads (20 g's or less at launch and upon entry). 

The details of the impact, either in liquid or on a solid, are difficult to 

analyze. For example, the low velocities involved are outside the range of 

validity for empirical equations for solid surface penetrators. 

the amount of energy dissipated in splashing down in a liquid is comparable to 

the total irnpact energy of the probe, so one must develop a method of analyzing 

the complex behavior of the splash. For these reasons and the resources available for this study, no attempt was made to predict transient landing loads. 

Furthermore, 

A purely kinematic approach to the problem yields interesting results. 

If the deceleration is limited to the 20 g entry load 

Figure 2.5-3 shows the relaticr, b?twnx impact velocity, stopping distance, and 

average deceleration. 

value and the descent time is two hours, then the probe must somehow be stopped 

over a distance of no less than 8 cm. In principle, the stopping distance can 

be controlled by the use of crushable materials or perhaps spikes. To accommodate a two hour descent, at least 8 cm of crushable material is needed. For 

longer descents, the crushable layer can be thinner. 

Winds at the surface have two effects upon these results. First, the probe 

has a transverse velocity at impact. Although the Conrath model of Reference 3 

predicts zero wind at the surface, the wind at 5 km is 12 m sec-1, which is 

twice as large as the worst case vertical impact velocity. 

vector would be substantially rotated from the vertical, implying that the 

loads might be shifted from the relatively strong nose to the weaker probe side 

walls. Second, wind shear may induce pendulum-like motions, contributing 

another component of transverse velocity at impact. Further study of surface 

survival requires an upper limit on dnd velocity at the surface. 

The net velocity 

Operating on the Surf ace 

The first question to ask about survival on an ocean is whether or not 

the probe will float. As long as the probe does not leak and the probe's mean 

density is less than that of the liquid, it will float. The first condition is 

violated in the baseline design, which is vented to the atmosphere. Assuming 

that the probe is made leak-tight, then the probe will float, because the. 

cm3) is appreciably less than the density of the ethane ocean and of the other 

candidate surface liquids listed in Table 2.5-1. In fact, as shown in Figure 

2.5-4, the probe will float with roughly half of its volume out of the liquid. 

probe's mean density of 0.35 g ~m'~ (probe mass of 105 kg, volume of 302 x 10 3 

Table 2.5-1 

Densities of Candidate Liquids on Titan 

Liquid Density (g ~m-~) 

Ethane 

Me thane 

Nitrogen 

55 

0.61 

0.42 

0.81 
Another concern is probe stability on the liquid surface. To prevent 

the probe from floating upside down, the center of gravity must be kept below 

the center of buoyancy. Although detailed mass properties are unavailable at 

present, the baseline configuration appears to satisfy this requirement. 

In the baseline design, the outer probe surfaces are within a few 

degrees of the atmospheric temperature during the descent. The insulation 

contained within the probe itself controls the environment of the units and 

maintains a large temperature dif€erence between the units and the outside 

world. The atmosphere does leak into the vented units, but with limited 

convection it quickly warms up. The small heat capacity of the vented atmosphere has a negligible cooling effect on the units. 

When the probe lands, this situation is basically unchanged. The 

Probe's exterior temperature does not change significantly, and the insulation 

continues to maintain the proper environment. One can then extrapolate the 

temperature predictions for the baseline design. At the end of the mission, 

the time rate of change of temperature is approximately -2 degrees C per hour, 

so an extra 20 minutes of surface operation will only lower the bulk Probe 

temperature by 0.7 degrees C. 

The above argument presupposes that the Probe stays "dry", that is, that 

no significant amount of liquid gets into the interior of the probe. For a 

vented design, it is likely that this condition will be violated, leading to 

several undesirable effects. First, the insulation will cool, bringing the 

cold surface closer to the units. The resulting heat loss rate is much higher 

than quoted above. Furthermore, the liquid will vaporize, further cooling the 

units through heat of vaporization effects. In addition, the vaporized liquid 

will generate pressure with effects dii:+cult to predict. 

a leak-tight probe is very desirable for thermal control. 

For all these reasons, 

In the baseline design, four lithium battery modules supply power. For 

a three hour descent, the depth of discharge at impact is 67% (based upon a 

conservative derating of battery capacity). 

is estimated at 98%, leaving a margin of 31%. As shown in Figure 2.5-5, this 

margin is equivalent to 106 minutes of operation after landing. 

The limit on depth of discharge 

The baseline probe design provides inlets for seveal of the science 

instruments. Some of those inlets are in danger of being plugged or broken 

during landing. Should surface survival become a hard requirement, instrument 

accommodation will have to be reviewed carefully to prevent these problems. 

Maintaining Communications with Carrier 

To maximize communications time at the higher data rate, the optimnm 

Probe landing time is as late as possible in the mission sequence. 

the nominal landing time must be earlier than this optimum to ensure that the 

probe reaches the surface before the link deteriorates if there are errors from 

the target plan. A preliminary error analysis indicates that the nominal 

landing time will be 15 minutes earlier than optimum. Hence, if the descent is 

nominal, then full communications from the surface will be available for 15 

minutes. However, if communications from the surFace are required for worst 

case slow descents, then the nominal probe landing must be pushed back even 

However, 

56 
earlier in the sequence, reducing the duration of high data rate communications 

in the atmosphere. 

Excessive probe tilting will degrade the communications link. A wide beamwidth probe antenna mitigates but does not eliminate the problem. The presence 

of surface winds will favor jettisoning the parachute and raises concerns about 

the effects of ocean waves on the Probe's attitude. If the Probe lands on a solid 

surface, stabilizing legs may be required to maintain suitable antenna geometry. 

The survival of the Probe's antenna is another concern. Since the dielectric 

properties of the surface are unknown, there is a risk of RF absorption or electrical shorting if liquid or snow gets into the antenna. Some sort of RF transparent housing may be required. 

Conclusions 

Since the probe structure is being sized for entry and launch loads, which 

are 20 g's or less, there is considerable concern about the probe's ability to 

survive impact. Crushable materials may offer a way of limiting the landing 

loads. Approximately 8 cm of crushable material will reduce the load to 20 g's 

for a worst case impact velocity. Substantial uncertainties exist because of 

the uncertainties in surface wind speeds. 

Thermal control and buoyancy considerations indicate that the Probe 

design should be modified to prevent liquid leakage if surface survival is 

required. Providing this is done, the probe will float, and the probe temperature will drop only 0.7 degrees over a 20 minute period. Substantial battery 

power margin is available. Science instrument accommodation will need further 

review if surface operation is required. 

If the descent is nominal, 15 ??inUtes of communications time will be 

available from the surface. If communications is required for worst case 

conditions, then some high data rate communications time in the atmosphere 

must be given up. Concerns about probe tilting and antenna survival will 

require further study if surface survival is required. 

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, 

3. DECELERATION MIDULE 

The deceleration module consists of the decelerator skirt, nosecap, and 

parachute. The skirt slows the probe during entry; both it and the nosecap 

protect the descent module from entry heating. When the skirt separates at 

subsonic speed, it extracts the parachute, which stows on the aft side of the 

descent module, to provide the correct descent drag. 

63 
3.1 Decelerator Configuration 

The decelerator configuration strongly affects the entry ballistic coefficient and science deployment altitude. 

large area and the benign heating allows flatter designs with higher drag coefficients. 

results in a sonic altitude of 190 km; halving the coefficent increases the 

altitude by about 25 km. 

blunted cones with diameters of 3.3 m or less to accommodate the launch envelope 

of the Shuttle. 

Mild entry loads permit designs with 

As shown in Section 2.2, a ballistic coefficient of about 14.5 kg/m2 

The design effort has focussed on 60' sphericallyDecelerator Concept 

Both fixed and deployable decelerator skirts allow fitting reasonable drag 

areas in the available envelope on the Carrier. 

larger drag area, desirable because of the corresponding increased sonic altitude. 

However, deployable designs have higher risk and cost. 

more promising umbrella and hinged panel deployment concepts with a fixed solid 

material design. 

Deployable designs allow a 

The study compared the 

The umbrella design (Figure 3.1-1) consists of cantilevered struts supporting a flexible skirt material. 

meeting at a common point and the material folds along pleats much like a conventional umbrella. Although flexible materials exist that can protect against the 

heat loads of the Titan entry, all have significant insulation, flow-through, and 

structural problems. Common failures include internal material abrasion and 

burn-through caused by porosity. The deployment of the material to a smooth 

surface following 9 years of in-space storage also poses a design risk. 

the performance of the flexible materiol at low (100 K) temperatures is not known. 

For stowage the struts fold toward the center, 

Finally, 

The hinged panel concept, proposed for the VOIR aerobrake, avoids the 

matertal problems of the umbrella design (Figure 3.1-2). 

multiple triangular panels that approximate the desired conical shape when deployed. Hinges along the panel edges allow the decelerator to fold to a starshaped pyramid configuration for stowage. Because each of the panels remains 

flat, folding does not affect the material selection. Despite the number of 

hinges, floating pins and standard mechanical design practices ensure reliable 

deployment. However, each of the hinge lines must have a seal to prevent 

flow-through, a major design complication. 

The skirt consists of 

The decelerator deployment sequence €or either option begins by releasing 

the stowed members, which allows centrifugal force to deploy the skirt. Dampers 

slow the deployment as the struts reach their final positions. If the spin-totransverse inertia ratio of a deployable decelerator design is less than 1.0, the 

decelerator deployment and the probe itself must meet additional symmetry requirements to assure successful deployment. These requirements affect geometry, probe 

mass properties, release times, and the deployment dampers. 

Although constrained to a slightly smaller diameter, a fixed decelerator 

(Figures 3.1-3 and 3.1-4) eliminates the cost, risk, and complexity of the 

deployable designs. A honeycomb structure covered with thermal protection 

material forms the skirt portion, which separates from the nose when the probe 

slows to subsonic speed. 64 
.. 

While Titan Probe can use either a fixed or deployable decelerator skirt, 

this study focussed on the fixed design because of its many engineering advantages. The scientific advantages of the slightly higher deployment altitude of 

a larger diameter deployable design comes with increased design complexity, 

cost, and mass. 

Envelope 

Although deployable decelerators of about 3.5 m diameter are possible, 

the Mariner Mark I1 Carrier envelope constrains the fixed decelerator to about 

a 3 m diameter, as shown in Figures 3.1-5 to 3.1-8. The probe mounts nose-in 

on the opposite side of the Carrier from the RTGs, keeping the c.g. close to 

the Carrier for the stiffness-driven attachment. The moderate level of entry 

aerodynamic heating makes penetrations of the nose by the attachments practical. 

Allowing 1.5 fn. clearance to the JPL-defined fuel tanks, the 175 in. 

diameter Mariner static envelopc cilnp:zains the probe diameter to 2.97 m. 

(Figure 3.1-6). Additional tank clearance reduces the decelerator diameter, 

at least in local areas. The location of the Carrier thruster outriggers 

provides more than 45O clearance to the edge of the skirt. 

Clearance to the high gain antenna determines the Probe's longitudinal 

station on the Carrier (Figure 3.1-7). Increasing the clearance Forces the 

Probe aft, complicating the two attachments to the equipment bay corners. 

The third Probe support attaches to the propulsion module bulkhead. Section 

5.1 describes the interface in more detail. 

Decelerator Design and Integration 

Figures 3.1-9 and 3.1-10 show tbe fixed decelerator design and attachment to the descent module. A 6 mm thick aluminum honeycomb frustum with 0.18 

and 0.15 mm face sheets rigidly supports the thermal protection material. A 

ring at the outer edge of the skirt sriffens the structure. The forward edge 

attaches to an inner box-section torsion ring that surrounds the outboard 

edge of the descent module shelf and lies slightly aft of it. Three brackets 

attach this inner ring to the shelf. Locating these attach points at the 

same angular positions as the Carrier attachments transfers the decelerator 

launch loads directly to the Carrier. 

To jettison the decelerator skirt, fully-redundant cutters sever the 

bolts holding the three brackets to the shelf. Differential drag pulls the 

skirt aftwards. The mild entry loads allow the use of existing, qualified, 

dual bolt cutters. Separation guides prevent skirt rotation and ensure that 

the ring and brackets clear the descent module. 

Table 3.1-1 details the decelerator mass. This mass represents a conservative design, using aluminum structure and maintaining comfortable thermal 

margins at the bond that attaches the thermal material to the'decelerator 

skirt. Further analysis can assess the feasibility of more agressive designs 

that use higher-temperature substrates and use the transient thermal capabilities of the adhesive bond. 

Table 3.1-2 lists the decelerator structural design considerations. 

Table 3.1-3 lists the dimensions of each decelerator element. 

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3.2 Thermodynamics 

Model Verification 

The Titan Probe entry regime falls within the range of previously-developed 

models for spherically-blunted cone bodies used on Pioneer Venus, Galileo Probe, 

Viking, and classified missions. However, much of the existing data base 

applies to entries in air or other atmospheres that differ from Titan's. In 

addition, the Titan entry extends over a longer than typical duration. 

Approximate analytic models exist for simulating the aerothermodynamics 

of the Titan entry, but they assume the parameters of an air atmosphere. 

However, NASA/Langley developed more exact entry models for outer planet probe 

studies which assumed 90% nitrogenll0X methane atmospheres, Tiwari, 81. 

Comparison of the air model with the Langley results verifies its applicability for the Titan analysis. 

similar heating rates, with the air model being conservative toward the end of 

the 6 kmlsec entry. 

by less. 

existing approximate solutions developed for entries i n a i r atmosphere. 

As shown in Figure 3.2-1, the models predict 

Entries using lower ballistic coefficients should differ 

Therefore, the Titan Probe aerothermodynamics analysis can safely use 

Entry Heating 

Figure 3.2-2 shows a heating rate pulse for a typical Titan entry. While 

the maximum heating rate is low, about 30 BTU/ft2-sec, the relatively long (1C)O 

second) duration causes a significant integrated heat load. 

Figure 3.2-3 shows the dependence CF the heat load on ballistic coefficient 

and entry path angle. Although steeper trajectories increase the peak heating 

rate, they alleviate the integrated heat load, simplifying the heat shield and 

internal insulation. Since the Titan entries present only mild heating rates 

which do not drive the design, the thermodynamics favor a steep entry. 

Material Selection 

Heating rates on the outer portions of the decelerator decrease with increasing 

distance from the nose, reaching about 60% of the stagnation point values at the 

edge of the skirt (Figure 3.2-4). 

configuration in air and carbon dioxide verified the analytic model used to predict 

these heating rates. 

Tests performed for the Viking aeroshield 

The mild heating rates allow either an ablative or heat sink nose cap design 

and do not require ablation of the skirt material although an ablative skirt 

has been chosen due to its minimum mass. Figure 3.2-5 shows the amount of solid 

beryllium needed at the stagnation point for 500° and 1000° backface temperatures 

as a function of the integrated heat load. Typical entries require 1.5 to 2 cm 

thickness at the stagnation point. 

Although the thick beryllium heatsink nosecap can provide adequate stagnation point heating protection, a thin beryllium cap backed by carbon foam weighs 

less and provides comparable thermal protection in the same volume. Figure 3.2-6 

78 
shows the performance of the carbon foam insulation with a 3.7 mm beryllium shell. 

With a shell this thin, the stagnation point temperature responds rapidly to 

the heating pulse (Figure 3.2-71, but the 4.9 cm of carbon foam keeps the 

internal temperature below 450 %. 

The decelerator skirt experiences much milder heating loads than the nose. 

Figure 3.2-8 shows the demands on heat protection materials for several different 

entry missions; the Titan entry places minimal demands on the low density 

abalator materials. Use of ESM 1004-X RTV-impregnated quartz cloth, a typical 

low-density ablator, can adequatley protect the skirt. Figures 3.2-9 and 3.2- 

10 show the ESM thickness required to maintain the aluminum substrate temperature 

below 450 %, accounting for the heat sinking effects of the substrate. 

3.2-11 shows the response of the substrate to the entry heat pulse. The ESM 

material, used for backface protection on many Earth reentry vehicles, avoids 

the mass and thermal expansion problems of metals. However, the RTV develops 

glass-like properties at low temperatures that require further study. 

Figure 

In addition to the protection on the forward side of the skirt, the back 

requires a closeout made of uncompressed Litaflex. This closeout lines the 

inside of the skirt and covers the back of the descent module. Figure 3.2-12 

shows the required thickness. 

Conclusions 

The existing entry aerothermodynamics models apply to the Titan Probe 

entries and predict mild heating rates. 

long heating pulse. Beryllium proves practical for local heat sinks or the 

entire nosecap; ESM-1004-X material provides adequate low-mass protection for 

the skirt. 

Steeper entries minimize a relatively 
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3.3 Aerodynamics 

Entry Analysis 

Using three-degree-of-freedom entry simulation, the dynamic pressure, 

heating rate, deceleration loads, and Mach 1 altitude for path angles from -40 

to -90 degrees and ballistic coefficients from 8 to 60 kg/m2 have been determined. 

Figure 3.3-1 shows the effect of path angle and ballistic coefficient on the 

deployment altitude and deployment dynamic pressure. 

from straight-in to -60° has a minor effect on sonic altitude. 

shallow entries further raise the Mach 1 altitude, they present communications, targeting, and heating concerns that mitigate this advantage. 

Reducing the entry angle 

Although more 

Figure 3.3-2 shows the variation in velocity, altitude, dynamic pressure, 

and axial load during a typical eutry. Table 3.3-1 summarizes the key parameters 

for entries with flight path angles of -45, -60, and -90 degrees: Section 2 

describes the entry analysis in more detail. 

Entry Dynamics 

Separation analysis (section 5.1) and six degree-of-freedom simulations 

verify the proper attitude of the Probe throughout coast and entry. 

stability imparted at separation preserves throughout coast the initial attitude 

set by the carrier. 

due to separation tipoff, but small enough to 

Spin 

The spin rate must be large enough to provide low coning 

1) allow practical generation by a simple spin-up mechanism 

2) allow angle of attack convergence by aerodynamic forces before peak 

heating, and 

3) give an acceptably small angle of attack at the onset of transonic 

instability. 

Section 5.1 describes the separation attitude perturbations. 

Aerodynamic forces during entry cause the Probe's initial angle of attack 

to converge toward the velocity vector at the time of peak heating. To facilitate the process, the separation attitude of the Carrier places the Probe in 

this nominal Orientation. However, errors cause some misalignment at entry. 

Figure 3.3-3 shows the attitude convergence for arbitrary initial misalignment 

of 6 and 12 degrees, several times larger than predicted values. Although a 

slower spin rate aids the convergence, in all cases the angle of attack dro.ps 

to less than 4' before peak heating occurs. Angles this small do not affect 

decelerator performance during the heating; little or no ablation occurs during 

the Titan entry. 

Aerodynamic forces also cause potential roll and transonic instabilities 

(Figure 3.3-4). The slow spin rate minimizes the destabilizing roll effects; 

a 10 rpm entry shows mild divergence in angle of attack at a sJlightly higher 

altitude than the entry with a 5 rpm rate. However, the higher spin rate 

minimizes the effect of transonic instability. The simulations show little 

angle of attack divergence in all the cases studied, so spin rates of 5 to 10 

rpm are acceptable for these entries. The final choice of spin rate depends on 

the separation mechanism design and induced errors. 

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3.4 Staging and Parachute 

Staging Sequence 

The Probe's angular momentum during coast retains the entry attitude 

set by the Carrier at separation, aligning the spin axis with the velocity 

vector at the expected time of peak heating. Small attitude deviations 

do not affect the heat shield performance; low gyroscopic stiffness (5 to 

10 rpm) ensures that aerodynamic torques cause the attitude to converge 

before peak heating. 

For the baseline design, an initial jettison of the gas chromatograph 

inlet plug enables sampling in the transonic region, (at about 191 km 

altitude) immediately following peak heating. A sequence of staging 

events then exposes the remaining science instruments to Titan's atmosphere 

(Figure 3.4-1 and 3.4-2). Thest events occur after the Probe slows to 

subsonic speed (at about 182 km) to avoid the complications of supersonic 

deployments. For the optional design, a ballasted nosecap drops away 

from the rest of the Probe to expose the gas chromatograph inlet. 

For either design, cutters sever the bolts attaching the decelerator 

skirt to the descent module equipment shelf. The lower ballistic coefficient causes the skirt to fall behind the descent module. Guides ensure 

separation clearance. 

As the skirt falls behind, it pulls a tether which extracts the bagged 

parachute from its housing on the descent module, deploying the parachute 

bridal and risers and stripping the bag from the canopy. Because of the 

low dynamic pressure, canopy inflatiu: takes longer than on the Galileo 

Probe or Pioneer Venus Large Probe, and analysis and simulation studies are 

required to assure limited descent module attitude excursions before full 

inflation. Once the skirt clears the descent module, instrument sensors 

deploy for forward view. The optional design allows a larger skirt ring 

diameter so most instruments can look forward without deployment. Yowever, 

the larger ring and attachment brackets required to clear the fixed instrument inlets on the side of the Probe increase the skirt structure mass. 

Separation Tradeoff 

Although the forward release of the descent module results in the simplest 

hardware, a rearward extraction offers advantages of heritage and heat shield 

con t i nu i t y . 

As described above, the forward deployment of the descent module uses 

the inherent difference in ballistic coefficients, avoiding the need for an 

extraction parachute. 

parachute, although a mortar can deploy it. Forward release requires a seam 

in the heat shield where the skirt joins the nosecap. 

The separating skirt pulls out the single descent 

t 

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A rearward extraction by parachute takes advantage of the Pioneer Venus 

and Galileo separation experience. It allows a one-piece decelerator, and the 

high parachute drag required for extraction slows the initial descent. However, 

rearward extraction requires a mortar and pilot parachute, and the much larger 

separation parachute (approximately 10 m diameter) which adds significant mass. 

98 
. 

and deployment of a separate smaller descent chute. 

requirements presents formidable stowage problems. 

The large parachute volume 

A maneuvering extraction parachute moderately reduces the area required 

by creating horizontal as well as vertical clearance (Figure 3.4-3). Figure 

3.4-4 shows two options for such a parachute. However, these designs require 

large area to separate from the high-drag decelerator (Figure 3.4-5). 

Descent Parachute 

Descent times of 2 to 4 hours need a much smaller parachute, with a base 

area about ten times smaller than necessary for rear extraction (Figure 3.4-6). 

The required descent parachute area is similar to Galileo Probe's with the 

specific size tailored to the desired descent time. 

Ribbon, ring-sail, and :".s!-- cf-r-band designs have been used in similar 

applications. However, the ribbon design offers better stability and development heritage from Galileo and Pioneer Venus. 

The various entry designs studied have dynamic pressures at Mach 1 from 

0.7 to 2.5 psf. This is consistent with the 1 psf minimum desired to ensure 

deployment: parachutes have deployed at pressures as low as 0.5 psf. Evolving 

mission and Probe designs must avoid a low parachute deployment pressure. 

No parachutes have operated in an atmosphere as cold as Titan's. The 

Probe warms the parachute until deployment, but strength, bending, and flutter 

tests must measure the material performance at descent temperatures. 

In addition, all parachute designs have a drag uncertainty that may 

require tests for a sufficiently accurate estimate of the descent time. 

Spin Rate Control 

Below 10 km altitude, the Probe descent imager desires a constant spin 

rate of abolit 10°/sec to scan the surface. Although either a spin parachute 

or vanes can provide this rotation, both lack predictability and suffer from 

"helix angle" effects at higher altitude. 

The inherent spin of all parachutes and the friction of the connecting 

swivel dominate the spin effects. Large vanes can overcome the swivel's' 

friction but require actuators and a control system to compensate for the 

indeterminate parachute spin effects. 

Spin parachutes deliberately induce an aerodynamic torque, usually with 

vents in the canopy. NASA has studied high-altitude spin parachutes of the 

disk-gap-band design. Although these designs can produce spin rates below 

20 rpm during terminal descent, the spin rate, proportional to velocity, 

exceeds 100 rpm at higher altitudes. 

For both the standard and spin-inducing parachute designs, spin rates 

vary widely, and testing is required to bound the performance. Further 

definition of imager and parachute design can help to define an acceptable 

scanning approach. 

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4. DESCENT MIDULE 

4.1 Configuration 

Envelope, mass, and access considerations determine the descent module 

configuration. A narrow cylindrical envelope requires a small descent module 

diameter and requires a folding decelerator (see section 3.1). Equipment must 

be packaged on two or more shelves resulting in an elongated configuration. A 

sharply curved decelerator nose, steep cone angle, or large equipment shelf 

forces the descent module aft and increases attachment structure mass; flatter 

nose configurations permit greater shelf area. 

Figure 4.1-1 shows descent module layouts for 50 and 80 cm diameter cylindrical envelopes. 

than the baseline single shelf design, e.g. three battery modules instead of 

four, and assume tighter unit pdck;r&. The 50 cm design shows two shelves, but 

would probably require a third shelf to properly accommodate harness and blankets. 

Support o€ the aft shelves consumes much of the available space, and the second 

shelf and associated aftward mass introduces cg location and inertia ratio 

problems. In addition, all instruments must deploy to view along the sides, 

routing of the gas inlets is difficult, and access to units requires extensive 

disassembly. 

These preliminary configurations show somewhat less equipment 

The 80 cm design resembles the Galileo Probe configuration, with the instruments and most of the subsystem units on the main shelf and an aft shelf supporting the communications equipment and a few selected units. Although still cramped, 

adjustments in the shelf sizes and unit layout permit a two-shelf design of this 

diameter. Access, though better than the smaller-diameter design, remains difficult, and some instruments still requtro deployment. 

Using the full envelope available on the Mariner Mark I1 permits a flat 60° 

decelerator cone with a large spherical nose (Section 3.1). With this geometry, 

a larger diameter shelf fits near the nose, where it provides favorable mass 

properties and simplifies the structural design. Figures 4.1-2 show a descent 

module design based on a 100 cm diameter shelf; the configuration allows adjusting 

this diameter if required. All the units fit on a single shelf, minimizing the 

structure mass, avoiding shelf supports, and providing complete access during 

integration. 

Given a strong preference for a single shelf design, two additional options 

were considered relating to the integration of the decelerator nosecap. The next 

two sections describe an integral and separable nosecap respectively. 

Descent Module Arrangement with Integral Nosecap 

Figure 4.1-3 shows the design of the descent module with an integral nosecap. 

The descent module layout places the heavier batteries and command and data 

processor units on the forward side of the shelf, keeping the cg forward and 

close to the Carrier attachments as shown in Figures 4.1-4, 4.1-5 and 4.1-6. The 

other subsystem units and the science instruments mount on the aft side of the 

shelf and the NMS rests in a cutout similar to the mounting of the Galileo Probe 

unit. 

room beneath it for the communications equipment, minimizing the cable runs. 

Slightly elevating the antenna gives it a clear field of view and provides 

106 
i 

Three brackets on the forward side of the shelf support the fixed 

beryllium nosecap and the three attachments to the Carrier which penetrate 

it. The attachment bolt catchers run almost tangential to the inside of ' 

the nosecap well away from the batteries and CDP. The three decelerator 

skirt brackets fasten to the aft side of the shelf above the nosecap supports. 

The skirt jettison bolt cutter assemblies hold the brackets outside the 

thermal blankets that envelop the rest of the module. 

Because the decelerator skirt requires radial clearance around the shelf 

for its aftward separation, the nosecap extends slightly outboard of the 

shelf to meet it. Therefore, this forces deployment of the AS1 inlets, 

nephelometer reflector, and NFR/irnager optics. Deploying the sensors places 

them outboard of the nosecap edge to provide a clear field of view and exposure 

to clean airflow. Both the GC and NMS inlets pass through the nosecap near 

the stagnation point. The inlet design prevents interaction with the flow 

effects. Although two inlets are :h.;s/n, they might be integrated into one 

inlet in the final configuration. 

Blankets, held by aluminum supports, cover the entire descent module and 

line the inside of the nosecap. 

compartment to avoid disturbing the rest of the insulation when it deploys. 

The parachute stows in a separate blanketed 

Descent Module Configuration with Jettisoned Nosecap 

Figure 4.1-7 shows integration of the descent module with a jettisonable 

nosecap. Jettisoning the nosecap allows it to use lower mass organic ablation 

materials for entry thermal protection that would otherwise contaminate the 

gas samples. Jettisoning the noseca: eliminates or at least simplifies the 

instrument deployments. The shelf arrangkaent is unchanged, but bolt cutters 

hold the nosecap at the end of the supports. About 24 kg of ballast mst be 

added to the nosecap to separate it ftm the descent module and skirt when 

the cutters fire. Removal of the nosecap exposes the instrument gas inlets. 

Since the jettisonable nosecap extends further beyond the shelf, a 

larger hole in the decelerator skirt is required in order to clear fixed 

instrument sensors. This simplifies instrument design at the expense of an 

extra separation and the nosecap ballast mass. To 

nosecap, the long nephelometer sensor deploys in a 

the Galileo Probe. 

- Mass 

Table 4.1-1 gives the descent module mass for 

single shelf baseline configuration. 

Conclusions 

avoid an even bigger 

similar manner to that on 

the integral nosecap, 

A simple configuration using a single 100 cm diameter shelf supports 

all science and subsystem units while maintaining favorable mass properties 

and complete accessability. Shelf diameter can grow to provide additional 

area as required. An integral nosecap reduces overall Probe mass and risk 

at the expense of additional instrument sensor deployments. 

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4.2 Structure 

The single-shelf configuration, mild entry deceleration, and moderate 

launch loads simplify the Probe structural design. 

structural requirements. 

load as the worst case for the trajectories of interest and JPL has provided 

estimated launch loads for the Probe location on the Carrier. 

Table 4.2-1 lists the 

The entry analysis in section 2.2 defines 20 g axial 

Stiffness dominates the structural requirements. The overall Probe 

natural frequency of 30 Hz allows a 10 Hz margin over the JPL-defined minimum. 

Coupling of the structural modes then requires that the main shelf frequency 

exceed 50 Hz. 

Equipment Shelf 

The single equipment shelf supporting all subsystem and science units 

dominates the structure. Mdels of the shelf using NASTRAN show high load 

margins resulting from the stiff design. A 2.0 inch thick core of 5056 aluminum 

honeycomb with 50 mil 2024-T81 facesheets maintains the 50 Hz frequency. 

Locating the three shelf supports at two-thirds of the radius enhances 

the shelf stiffness (compared to edge support) while placing the Carrier attachments correctly for separation. Loads from the decelerator pass directly to 

the Carrier attachments and do not affect the shelf. However, the cutout for 

the NMS instrument (assumed to resemble the Galileo Probe NMS) lowers the shelf 

stiffness. A surface-mounted NMS or a local shelf stiffner would offer some 

saving in shelf mass and thickness. 

Mass 

Table 4.2-2 lists the mass of each structural element. As the prime loadcarrying member, the shelf dominates the structure mass. 

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4.3 Thermal Control 

Requirement s 

The probe must maintain proper internal temperatures during the 9-year 

cruise to Saturn, 6 to 20 day coast from spacecraft separation to Titan entry, 

and descent into the cold Titan atmosphere. 

temperature requirements. In addition to these ranges, the batteries should 

remain below 2OoC for long-term storage. 

Table 4.3-1 summarizes the operating 

Cruise and Coast Thermal Control 

Figure 4.3-1 shows the thermal control concept. The approach follows that 

of the Galileo Probe, with 4 cm thick blankets surrounding the descent module 

that contains the units. However, €or the Titan Probe, the radioisotope heater 

units (RHUS) that keep the probe warm during cruise and coast are placed inside 

the descent module thermal blanKets where they can also help to warm the module 

during descent. 

to deploy without affecting descent module thermal control. 

A separate parachute blanket and RHU for the parachute allow it 

Because the Probe is designed for a thermal balance while coasting alone 

after separation from the Carrier spacecraft at Saturn, it requires shade from 

the sun near Earth and isolation from the Carrier. This isolation may include 

shading from the Carrier RTG thermal radiation, if the small view factor does not 

sufficiently minimize the effect. The Probe-Carrier attachments may also require 

conductive isolation; exact values of acceptable radiation and conduction require 

definition and are beyond the scope of this study. 

The RHUS must maintain the Probe between -loo and 2OoC with effective blanket 

emissivity from 0.02 to 0.04 (Figure 4.3-2). Since variations in blanket installation and penetration cause the range of emittances, tests are required to determine the exact number of RHUS. A baseline value of 40 watts falls in the expected 

range. 

Descent Thermal Control 

Table 4.3-2 compares the three key parameters of the descent thermal environment to previous probe missions. The conductivity of Titan's atmosphere, about 

ijne-sixth of Jupiter's, reduces the rate of heat loss. Free convection is approximately the same as on Pioneer Venus. Although the cold Titan atmosphere tends 

to cool the Probe, it also results in low conductivity of blankets and metal 

parts compared to the Galileo descent. 

Figures 4.3-3 and 4.3-4 show the variation in Probe temperature during 

descent for initial temperatures of +loo and -1OOC. The profiles show the effects 

of 4cm and 8 cm (1.5 and 3.0 inch) thick insulation, with up to 65 watts of heater 

power added to the internal dissipation as described by the power budget in the 

following section. Heaters would primarily warm the batteries because they are 

most sensitive to low temperatures. Appropriate heater power and insulation 

thickness can maintain internal temperatures for any of the descent times considered. 

initial temperature and no supplemental heater power; modest heater power ensures 

maximum battery performance for longer missions or cooler initial temperatures. 

The baseline 4 cm thick insulation supports a 3-hour descent with +loo 

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4.4 Power Subsystem 

Operat ion 

The Probe must internally supply 60 mW, 32 W, and 153 W of regulated 

battery power to the units during coast, pre-entry, and descent, respectively. 

It also must distribute Carrier power supplied through two Probe power interface units (PPIUs) for checkout during cruise. In addition, the power subsystem 

must provide firing current for up to six simultaneous separation and deployment 

pyrotechnic initiators. 

Figure 4.4-1 shows the subsystem operation. The subsystem power interface 

unit (SPIU) diode-ORs and regulates the power lines from four Lis02 battery 

modules It also contains reconditioning circuits to depassivate the batteries 

before entry. Orbiter power supplied by the PPIU passes through the SPIU for 

switching and distribution. 

The SPIU switches power to the command and data processor (CDP), the CDP 

The IPIU in turn provides overvoltage 

coast timer, the instrument power interface unit (IPIU), the communications equipment, and the pyro control unit (PCU). 

regulation, fusing, and switching of science instrument power. The PCU contains 

the arming and hybrid driver circuits for the pyrotechnic initiators. 

measured performance of the Galileo ?robe LIS02 batteries, battery power can drive 

the squibs directly; no thermal batteries are required. 

Based on 

Table 4.4-1 lists the subsystem unit masses and required modifications from 

the Galileo design. The added battery module supports the longer Titan mission 

and provides the desired margin for this stage of the program. Repackaging the 

Galileo Probe electronics eliminates tk hermetically-sealed, pressure tolerant 

(20 bar) cases to save mass. 

Table 4.4-2 snows the power budgeL for a 3-hour descent. All loads carry a 

10% contingency. In addition, allocated battery shelf losses of 0.1 A-hr/module/ 

:;ear consume 3.6 A-hr, although Galileo Probe has not observed losses this high. 

I’hd science instruinents drive the checkout power requirement; sequencing the 

checkout would reduce the maximum power if necessary. 

Carrier Power Interface 

Two ?robe power interface units (PPIUs) on the Carrier supply power for 

Probe operation during the cruise to Saturn (Figure 4.4-2). 

switched, regulated, current-limited DC lines from the Carrier. Two DC-to-DC 

converters shift the voltage from the +15V input to the desired +33 V and +39 V 

values. Like Galileo Probe, three busses pass from the converters to the probe 

SPIIJ. The first powers the command interface functions and contains a Shuttle 

lockout and bleed resistor circuit for safety. The second bus powers the CDP; 

the third bus provides current and voltage limited checkout power. The second 

PPIU redundantly supplies the same interfaces. 

Each PPIU accepts 

126 
Battery Selection 

Table 4.4-3 summarizes the characteristics of batteries considered for 

the Probe. Safety, mass, cost, shelf life, and transition from the coast 

timing phase of the mission to the higher power required just before entry 

primarily affect the battery selection. Although other systems offer some 

advantages, the overall tradeoff favors the Galileo probe LIS02 batteries 

because of the low cost and mass, and developed status. 

capacity margin overcomes both the temperature and shelf life disadvantages 

of this type. The entire power subsystem therefore retains most of its 

Galileo Probe heritage. 

Carrying sufficient 

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4.5 Communications Subsystem 

To simplify the communications analysis the transmitter power was set 

at 2.5W, S-Band frequency (2295 MHz) was assumed, and atmospheric absorption 

effects and galactic and synchrotron noise were ignored. . A bit error 

rate of was assumed based on the Galileo Probe requirement, and receiver/ 

demodulation losses, cable loss, and noise temperature (437%) were set at 

the values measured for the Galileo Probe. 

disk noise of 90%. 

Link budgets assume a worst case 

Section 2.4 describes the tradeoff of other communications parameters 

and derives a baseline Probe antenna beamwidth, Receive antenna beamwidth, 

and BPSK receiver design employing a 2nd order tracking loop (see Section 5.3). 

The Probe communication subsystem employs a parallel string configuration with no cross-strapping, sidhr to that used on Galileo Probe (see Figure 

4.5-1). The two phase modulated S-band channels are transmitted to the Carrier 

at orthogonal polarization and offset frequency. The two strings transmit 

nearly redundant telemetry information from the CDP unit. 

Ultra Stable Oscillators (USO) generate the RF carriers with sufficient 

frequency stability to allow determination of atmospheric wind from doppler 

measurement. An exciter in each string biphase modulates the carrier with 

the composite convolutionally encoded data stream from the associated CDP. A 

power amplifier delivers 2.5 watts to the antenna. All the units have their 

heritage in existing hardware. Table 4.5-1 shows mass, power, and dimensions 

for each unit. 

Inclusion of radio science irr,,:tts both the Probe and Carrier. If radio 

science is deleted the Ultra Stable Oscillators can be replaced by temperature 

controlled crystal oscillators decreasing mass, power and dimensions (see 

table 4.5-2). Eliminating radio science also allows similar simplification 

of the oscillator in each receiver, a small receiver data rate decrease 

( 250 bps), and some reduction of receiver AID hardware. 

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136 
4.6 Command and Data Subsystem 

Read remen t s 

The Titan Probe Command and Data Subsystem (OS) must communicate with 

the Carrier during cruise to perform Probe checkout. After the Probe separates, 

the CDS must time the coast period to ensure pre-entry warm-up of the stable 

oscillator. The g-switches give the CDS an indication of the forces experienced 

by the Probe as it enters Titan's atmosphere. The CDS must use this information 

to derive the timing of parachute deployment, decelerator separation, and commands 

to the science instruments. 

As the Probe descends, the CDS must deploy sensors, command instruments 

to take measurements, format science and engineering data, and send the 

composite data to the communications subsystem for transmission to the Carrier. 

The estimated number and type of command and telemetry channels are nearly the 

same as Galileo Probe. 

Hardware Description 

The Probe command and data handling subsystem comprises the two halves 

of the CDP, two PCU's (Pyrotechnic Control Unit), and four g-switches. 

Table 4.6-1 lists the mass, power, and dimensions of each of these units. 

The mass and power column has the total for all units on that line, while the 

dimensions apply to individual units. 

Interfaces 

Figure 4.6-1 shows the Carrier command interfaces with the Probe, receiver, 

and separation mechanism. Commands come to the Probe over the Carrier serial 

bus. Note that the serial bus interface passes through the cable cutter. 

The Bus Interface Unit (BIU) handles Carrier bus protocol and decodes commands 

for the Probe. An identical command interface exists between the Carrier and 

the receiver. The Carrier provides a number of drivers to fire NASA standard 

initiators in the cable cutter and separation mechanism. 

Data interfaces between the Carrier and Probe-related hardware are shown 

in figure 4.6-2. All Probe data interfaces also pass through the cable cutter. 

The Probe, separation mechanism, and the receiver have a number of passive 

analog and bilevel transducers that are sampled by the Carrier CDS. Sampling 

of these telemetry points does not require the participation of the receiver 

or Probe electronics. 

CDP-formatted checkout data pass through the Probe exciters and travel 

at RF to the receivers over coaxial cables. The receiver then processes the 

data as if they were received through the antenna and sends them to the 

Carrier serial bus via its BIU. This scheme provides an end-to-end checkout 

of both the Probe and receiver. An option exists whereby the Probe returns 

data directly to the Carrier through its BIU. This option would allow checkout 

of the Probe without involving the receiver or Probe RF equipment. The baseline 

does not employ this option because the increased CDP complexity outweighs 

the potential checkout power savings. 

137 
CDP Redundancy Options 

Four redundancy options for the Command and Data Processor (CDP) have been 

identified. 

redundancy and complexity. In Option A, shown in figure 4.6-3, the unit simply 

comprises two identical and independent halves. The Carrier selects the prime 

half prior to Probe separation, and the other half remains off for the remainder 

of the mission. This option has the advantage of simplicity, but is vulnerable 

to single failures after separation. 

In the following descriptions, the options are ordered by increasing 

Option B (figure 4.6-4) offers increased reliability by using a self-test 

analyzer to test both CDP halves and select the prime one just before entry. The 

analyzer monitors prime timing signals from then on, and selects the backup half 

should the timing go out of tolerance. If the backup half is viable, its data 

function remains synchronized with that of the prime half, and allows redundant 

data collection and transmissicl. VLally, the Carrier can disable the self-test 

analyzer should it fail prior to separation. This option was selected for the 

Galileo Probe, because volume, mass, and schedule precluded the following options. 

As shown in figure 4.6-5, the two CDP halves of Option C operate independently to provide active redundancy. The identical halves are not synchronized, 

and hence, require asynchronous serial data interfaces with the non-redundant 

science instruments. The asynchronous data interfaces can be implemented by each 

instrument or by a separate unit (see following paragraphs). 

enabling, coupled with judicious command output circuit assignment, gives protection against single-failure false commands. CDP oscillator fault protection 

precludes premature command sequencing as a result of a single failure (see 

following paragraphs). 

and eliminates the synchronization hardware and self-test analyzer. 

Critical command 

Option C provides increased reliability over Option E, 

Option D, figure 4.6-6, is the most reliable and complex of the four options. 

The data function is essentially the s'pme as that of Option C. The command 

function, however, employs three sequencers and majority voting hardware. As with 

Option C, no self-test analyzer or synchronization hardware is required. Premature 

sequencing by one of the three command sequencers is not mission critical to this 

option. Coarse synchrony among the sequencers is necessary, however, and demands 

a certain stability of the CDP oscillators. 

Base 1 i ne Subsystem 

Option C has been chosen for the baseline subsystem because it offers the 

necessary reliability with lowest complexity and cost. Figure 4.6-7 shows the 

independent CDP halves and the remainder of the subsystem. Carrier commands are 

routed to both halves, either of which responds if powered and properly addressed. 

G-switch inputs provide timing cues to the CDP for descent sequencing. Each 

critical command requires an enable to prevent single-failure false commands. In 

addition, a latching g-switch enable within the PCU prevents execution of critical 

commands before entry. Non-critical commands do not depend on enables and execute 

directly if sent by either half. Each CDP half provides asynchronous telemetry 

timing to the Probe and science instruments, which then return separate, asynchronous data streams. The CDP halves independently format data for parallel 

transmission by the communication subsystem. 

138 
Pyrotechnic Control Unit 

The redundant PCU provides isohtion diodes, arming relays, and driver 

hybrids to meet requirements for mission reliability and shuttle safety. 

Hybridization of the driver circuits has proven to significantly reduce the 

mass and volume of this unit on the Galileo Probe. These same advantages 

apply to the Titan Probe. The arming circuit requires a command from the 

CDP in addition to g-switch activation before the drivers are connected to 

the power bus. Finally, each fire command from the CDP must be preceded by 

an enable command. 

Oscillator Fault Protection 

The baseline subsystem requires that sequencer oscillators be protected 

from single failures that cause oscillation at some higher frequency, since 

the resulting premature sequencing is potentially mission catastrophic. 

This type of failure is charact.3ri7ed by oscillation either at a harmonic 

frequency or in an undesired mode. In either of these cases, a bandpass 

filter centered about the fundamental frequency would block the undesired 

oscillation and halt the faulty sequencer. 

Asynchronous Science Data Interfaces 

The asynchronous data acquisition of the baseline CDP halves places 

special requirements on the serial interfaces between the CDP and the 

science instruments. Implementation of these interfaces requires either a 

second data output circuit from each instrument or a separate data splitter 

unit. Figure 4.6-8 shows one option in which the science instrument design 

remains essentially unchanged. The instrument provides asynchronous data 

outputs by simply replicating the cirwitry from the A/D converter to the 

output connector. Figure 4.6-9 shows an ytion in which the science instrument retains the single data interface typical of Galileo Probe instruments. 

This option requires a separate data ;plitter circuit, like the one shown, 

for each instrument. Intermediate options require that the science instruments provide some of the buffering and/or control functions. 

In the baseline design, science instruments provide asynchronous data 

outputs to the CDP. The implementation of these outputs is an instrument 

design issue whose impact depends on data handling hardware complexity. 

Convolutional Encoding 

The tradeoff regarding the implementation of convolutional encoding 

for the Probe-receiver link is essentially hardware vs. software. The 

hardware option requires an estimated 100 equivalent parts, which is about 

2.5% of a Galileo CDP. Software encoding requires about 25 instructions 

per bit, which occupies 8.3% of the baseline processor's time. 

design uses software encoding because it imposes an acceptably low demand 

on processor time, and requires no additional hardware. 

The baseline 

139 
Data Rates 

Table 4.6-2 lists the factors that contribute to the rate of data transfer 

from the receivers to the Carrier for both the high and low Probe data rates. 

The factors comprise: Probe formatting and convolutional encoding, receiver 

engineering data, radio science data, receiver packetization, and a factor of 

2 to account for receiver redundancy. Packetization follows the CCSDS recommendation, uses 1024-bit packets, and does not employ error control fields. 

The highest data rate seen by the Carrier is around 45,000 bps - well within 

the capability of the serial bus. 

140 
TABLE 4.6-1 - UNIT MASS, POWER, AND DIMENSIONS 

- UNIT QUANTITY MASS (KG) POWER (W) DIMENSIONS PER UNIT (CM) 

L W H 

CDP HALF 2 14.4 20 30.5 15.2 8.9 

PCU 2 3.45 0 27.9 10.8 8.3 

G- S WITCH 4 0.55 0 3.8 3.8 3.8 

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151 
5. CARRIER INTERFACE 

5.1 Separation 

Reauirements 

The requirements for Probe separation from the Carrier derive from the 

basic requirement that the separation velocity be sufficient to assure 

separation, Furthermore, the Probe angular momentum must be large enough to 

assure a nose first atmospheric entry and sufficiently small to permit 

satisfactory Probe precession to zero angle of attack during entry. 

Any candidate separation design must be within Carrier requirements for 

disturbance impulse and must meet mission requirements for separation precision. 

Imprecision in the separation velocity vector produces targeting errors and 

imprecision in the angular momentum vector produces errors in the angle of 

attack at entry. 

Design Selection 

The requirements can be met with three separation concepts. Each is 

characterized by the way that the angular energy is applied. 

The first is the Galileo approach: ejecting the probe from a spinning 

carrier. The carrier provides the angular velocity through its spin, allowing 

the separation hardware to provide only a translational impulse. The gyro 

control limit (2O/sec) of the 3-axis stabilized Mariner Mark I1 Carrier 

eliminates further consideration of this approach. 

The second approach is a gradual buildup of angular momentum with a Carrier 

mounted spin up system. The separation impulse is also only translational. 

The third concept is simultaneous translational and rotational impulse. 

A wide variety of hardware implementations are possible. The difficulty is 

to meet spin and delta V requirements without sacrificing separation precision. 

Table 5.1-1 shows the hardware implementations examined for this study. All 

three designs can provide the required separation energy but differ greatly in 

design complexity. 

The spin table system applies a gradual spinup through a motor driven 

platform with springs on the platform separating the Probe when the proper angular 

velocity is reached. All shuttle launched HS-376 satellites use this approach. 

Application to the Titan Probe mission necessitates costly scaling down of complex 

and expensive hardware. Furthermore, the volume of the hardware pushes the Probe 

into the Shuttle envelope, significantly reducing the allowable size of the fixed 

decelerator. 

The centerline spring system and the three canted spring system are two 

different methods of applying a simultaneous spin up and ejection. The centerline approach incorporates a single spring mechanism attached at the Probe 

centerline. The spring is mounted in a helix-slotted cyclindrical housing which 

152 
A . 

attaches to the Carrier. Guide pins in an outer jacket force the Probe to 

rotate as the spring extends, constraining the Probe's motion during separation. The angle of the helix determines the ultimate angular velocity 

of the Probe at release. A single separation bolt initiates separation 

which occurs in approximately 0.5 seconds. 

Some sample design characteristics for this system are given below. 

The assumed separation parameters are consistent with the requirements 

listed above. 

probe mass - 141 kg 

probe inertia - 44 kg-m2 

probe spin rate - 10 rpm 

probe velocity - 0.2 m/sec 

helix angle - 41.3O 

torque radius - 16.9 cm 

spring constant - 12.1 kn/nl 

spring diameter - 5.7 cm 

spring compressed length - 3.0 cm 

spring length at release - 8.0 cm 

release time - 0.44 

The primary disadvantage of this system is the complexity of the design 

for Probe attachment. The centerline application of force significantly 

complicates nosecap design. It also requires a separate attachment structure 

to carry launch loads. 

By incorporating the springs into the load bearing structure the three 

canted spring approach minimizes the structural complexity. This system uses 

three springs mounted near the radius of gyration. Canting the springs at 

the proper angle provides the proper ratLn of angular to translational momentum. 

Unlike the centerline system, the three spring approach requires the simultaneous release of three mechanisms foi proper Probe deployment, introducing 

additional separation errors due to non simultaneity. Analysis of a similar 

three spring separation for Galileo Probe indicates a time uncertainty of 

release of 2 msec. 

Concepts for a three canted spring design range from a guided release 

resembling the centerline spring to an unguided release in which the mass and 

inertia of the Probe preserve the alignment during ejection. The unguided 

three canted spring design has the advantage of simplicity and the disadvantage 

of less precision. 

as low cost of the three spring design results in its selection as the baseline 

separation scheme for this report. 

angle of attack are acceptable, the baseline further assumes unguided separation. 

A discussion of separation parameters and dynamics for the baseline design 

follows the hardware description in the next section. 

The relative structural and mechanical simplicity as well 

Since the predicted targeting error and 

153 
. 

Baseline Separation Hardware - Three Canted Spring Design 

Figures 5.1-1, 5.1-2, and 5.1-3 show the Carrierhrobe mechanical interfaces. 

The three support structure designs (monopod, bipod, and tripod) accommodate 

the different Carrier mounting locations along the bulkhead and equipment 

ring (Figure 5.1-1). 

separation nuts. 

All three have plates holding the spring cartridges and 

(Figure 5.1-2). A ring links the three supports together. 

Three elliptical penetrations in the Probe decelerator nose allow the 

spring cartridges to mate to the main Probe structure (Figure 5.1-3). The spring 

cartridge tube carries loads from the Probe to the support plate. It houses a 

10 cm compressed spring which extends to 15 cm for Probe release. A bolt running 

along the centerline of the spring ties the support to the Probe. The ball-in 

socket at the cartridge end bears loads in the direction of the Probe's 

centerline. 

Insulation around the bolt catrher r-atects the equipment module from concentrated 

heating loads during atmospheric entry. 

A bolt catcher within the Probe attaches to the equipment shelf. 

The umbilical cord to the Probe runs along one of the three struts and enters 

the Probe through another penetration. A cable cutter and retraction springs 

separate the cable prior to Probe release. Section 4.6 discusses the signals 

carried by the umbilical cable. 

Firing the three explosive nuts forces the bolts entirely into the Probe and 

frees the springs for Probe release. Tabs in the cartridges impede any further 

spring extension beyond 15 cm. Crushables at the end of the bolt catcher 

prevent any damage to the Probe. The Probe separates with a corkscrew motion - 

simultaneoiisly spinning and translating away from the Carrier. The dynamics of 

the separation are discussed below. 

Separation Dynamics 

The following design parameters separate a probe with a mass of 141 kg and 

inertia of 44 Kg-m 2 at 0.2 m/s and 10 rpm spin rate. 

cant angle: 

radius to spring attach points: 

spring constant: 

spring compressed length: 

spring stroke: 

release time: 

angular impulse: 

linear impulse: 

75.2 deg 

0.45 m 

7184 N/m (3 places) 

10 cm 

5 cm 

0.15 s 

46 Nms 

28.2 

The resulting impulses applied to the Carrier are within JPL's limiting 

values for such impulses:. 100 Nms angular, 60 Ns linear. A wide range of 

separation velocity and spin rate is allowable. Figure 5.1-4 sketches the bounds. 

The orientation of each spring provides both a tangential (spin-up) and 

axial (separation) force to the Probe. There is no radial component. The Probe 

contact surface is a ball-in socket, as is used €or Leasat frisbee ejection. 

By attaching the spring to the support structure with a pin the cant angle can 

change slightly during ejection. With the example design, the spring cant angle 

154 
rotates from 75.2 deg to about 72 deg during ejection. The springs apply 

their force at 0.45m, close to the radius of gyration of 0.56m. Mounting 

the springs further from the Probe radius of gyration would increase the 

amount of spring rotation. A spring guide prevents excessive spring bending 

in the Probe radial direction. 

The key concern of the three spring spinlejection scheme is the possibility of undesired torques perpendicular to the intended spin direction 

causing the Probe to tumble. Tumble will not occur if there is a fast enough 

spin up -- or large enough inertia. 

To help understand the action of this system we can consider the dynamic 

response of classes of systems to an angular impulse about a transverse axis. 

For a spinning body this impulse induces nutation about an offset angular 

momentum vector. For a non-spinning body this impulse induces tumbling (if 

given enough time). For a body undergoinx spin-up this impulse can start the 

body tumbling but as long as the motion doesn't prevent proper application of 

the spin-up torque, the motion ~?CPW- nutation as the body is spun up. The 

response approaches the spinning body response as the spinup time approaches 

zero. 

An estimate for the maximum undesired transverse angular impulse comes 

from previous analysis of the Galileo Probe ejection. In this case, the 

effect of all misalignments, asymmetries, mistimings, and other separation 

torques in the transverse direction is calculated to be less than the equivalent of 1 Nms angular impulse applied at the start of separation. 

If a 1.0 Nms transverse angular impulse is applied to the Titan Probe at 

the start of spin/eject and assume separation within 0.2~~ the symmetry axis 

tips less than 0.3 deg, giving displacements less than 5% of the spring 

stroke. Also, after 0.07s the Probe has acquired half of its final angular 

momentum. At this time, the symmetry axla has tipped only 0.1 deg and the 

displacements are less than 1.5% of the spring stroke. Since the displacements 

are small, they do not prevent the spi!ngs from spinning up the Probe, and 

the Probe achieves a stable spinning condition although the initial condition 

has no angular momentum. Figures 5.1-5 and 5.1-6 demonstrate successful 

separation for this case based on dynamic simulation. 

A perfect ejection would give the Probe angular momentum and separation 

velocity vectors with the desired magnitude and direction. 

Probe angular momentum and separation velocity come directly from two sources: 

1) Carrier orbit and attitude errors, and 2) asymmetric ejection effects. 

A two body simulation program has been developed to study asymmetric separation 

effects. The simulation uses JPL-supplied mass properties for the Mariner Mark 

I1 carrier and Hughes estimates for Probe mass properties: 

spin inertia, and 24 KG-m2 transverse inertia. The effects of spring mismatch, 

Carrier reaction, release mechanism mistiming, and Probe center of mass offset 

were studied separately. 

The errors in the 

141 Kg mass, 44 KG-m2 

Simulation dl assumes an inertially fixed Carrier, matched symmetric springs, 

simultaneous separation device release, and perfect Probe balance and alignment. 

Figure 5.1-7 shows that the Probe acquires the desired 0.2 m/s separation velocity 

and 10 rpm spin rate following an ejection that takes about 0.15 s. 

155 
Simulation 82 shows the effect of the Carrier reaction to the separation 

forces. The separation axis does not pass through the Carrier center of 

mass and does not lie along a principal axis. As a result, the Carrier 

rotates about all three axes as the Probe is ejected. Figure 5.1-8 resembles figure 5.1-5. This similarity indicates small errors in magnitude: 

about 1% in spin angular velocity and 0.5% in axial velocity. Figure 5.1-9 

shows the resulting components of separation velocity and angular velocity 

in the directions perpendicular to the desired direction. 

Simulation W3 shows the effect of a large spring mismatch. Spring 

mismatch is expected to be the largest asymmetric effect associated with 

separation device asymmetry. This simulation assumes one spring has a spring 

constant 110% of nominal while the other two have spring constants 90% of 

nominal. In reality, this mismatch can be reduced by making a special effort 

to select matched springs. The conservative simulation shows about 2% change 

in the magnitudes of both the separation velocity and the angular momentum. 

Figure 5.1-10 shows the lateral velocity and transverse angular velocity 

profiles. 

Simulation #4 shows the effect of a mistiming in one release mechanism. 

The 0.002 s maximum mistiming estimated for Galileo is used here. The simulation predicts a 1.2% error in separation velocity and a 0.2% error in 

angular momentum caused by the mistiming. Figure 5.1-11 shows the lateral 

velocity and transverse angular velocity profiles. Note that the final 

lateral velocity is much smaller than the lateral velocity in the middle of 

the ejection. 

Simulation 85 shows the effect of Probe center of mass offset. A 1 cm 

offset from the release system axis ot zynmetry would cause a 1 cm/s lateral 

velocity for a 10 rpm release and guided separation. In the unguided system, 

figure 5.1-12, the final lateral velocity is less than this because of the 

action of the springs during the ejection. The simulation predicts a 1.2% 

error in separation velocity and a 0.2% error in angular momentum. 

Imprecision in Probe angular momentum vector direction causes angle of 

attack errors at entry. 

lateral velocity) causes targeting errors at entry. Table 5.1-2 summarizes 

the effect of the asymmetries studied in lateral velocity and angle of attack. 

The exact relationship between separation velocity error and targeting error 

depends on the Probe trajectory. If there were no gravitational effects, 

each mm/s of lateral velocity would cause 1.7 Km B-plane targeting error. 

Imprecision in separation velocity (particularly 

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5.2 Receive Antenna 

Figure 5.2-1 shows a possible Carrier mechanical integration of the baseline 

0.75m diameter receive antenna. 

below the Earth pointing high gain antenna. 

indicate that an antenna up to 1 m diameter can mount in this area, but the 

integration of any antenna requires a detailed examination by the Mariner Mk I1 

configuration engineers at JPL. 

may be required to point anywhere from -5O to -140O from the anti-Earth line. 

An antenna positioner is necessary to all-ow this mission flexibility and, as 

shown in section 2.4, considerably increases communications performance if used 

to track the Probe position during Carrier over flight. The receive antenna 

may also mount to the scan platform which eliminates the need for a separate 

pointing mechanism, but requires a longer cable run to the receiver. A thorough 

tradeoff of antenna integratiov ontlris requires refinement of Carrier design 

constraints and link performance requirements. 

The antenna mounts on the top of the spacecraft 

Preliminary configuration studies 

Depending on the final mission plan the antenna 

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OREGIlPliAk PAGE IS 

W POOR QUALITY 

5.3 Receiver 

The mission analysis in section 2.4 shows two ways that the receiver' 

design is a determining factor in the performance of the system. 

design sets both the requirement for Eb/No (energy to noise level ratio) as 

well as the tolerable level of doppler rate. These requirements limit the 

time that at a given rate can be supported. 

The receiver 

The key driver in the receiver design is the requirement for two data 

rates: one at 100 bps and the other at 10 Kbps. Both data rates must be 

received with an error rate better than one in a thousand. In addition, the 

receiver must detect when the transition between the two data rates occurs. 

The purpose of the higher rate is to allow for an imaging device which requires a data rate of at least 2 kbps and preferably 10 Kbps. The preliminary 

strategy is to initiate the imaging at an altitude of 10 km with the Probe 

clock triggering the switch to -Lc :.:gner data rate. As a point of reference, 

the Galileo Probe has a simplification of only one data rate of 128 bps and a 

complication of a lower C/No (carrier to noise ratio) at acquisition. 

Modulation Formats 

The first step in the receiver design is to determine the means of 

transmitting digital data from the Probe to the receiver. Given that there 

are no bandwidth constraints, three modulation formats are applicable: (1) 

non-coherent frequency shift keying (NC FSK), (2) non-coherent differentially 

encoded phase shift keying (NC DPSK) and (3) coherent binary phase shift 

keying (BPSK). NC FSK provides the least complex receiver structure while 

BPSK is the most complex. If the Eb/No requirements for the different modulation formats were identical, then IX PSK would be the obvious choice for 

implementation; however, Table 5.3-1 demonstrates that this is not the case. 

Table 5.3-1 shows the theoretical (implement tion losses are not included) 

Eb/N required to achieve a bit error rate of lo-'. Requirements are listed 

for goth uncoded data and rate 1/2, constraint length 7, convolutionally encoded data. Constraints on the receiver's output data rate preclude the use 

of soft decision information (see section 4.6). The requirements for the 

convolutionally encoded data assume that the error rate on the Carrier to 

Earth link is not significant. Note that while elimination of coding would 

seem to reduce receiver complexity, this is not really the case. Removal of 

the convolutional code would require the inclusion of Manchester encoding to 

ensure enough bit transitions for proper signal synchronization. 

For coded data, BPSK clearly provides the best performance while NC FSK 

is the worst performer. The issue is whether the reduced complexity of NC 

FSK offsets the 5.2 dB penalty in the Eb/No requirement. 

requirment lowers the time that can be supported at a given data rate (section 

2.4); however, if both NC FSK and BPSK provide sufficient time, then NC FSK is 

the proper choice. To quantitatively evaluate the differences in receiver 

complexity, specific receiver designs must be examined. NC DPSK was not 

included in this study because it was marginally less complex than BPSK. 

This higher Eb/No 

173 
Receiver Design Tradeof f - BPSK 

The simple NC FSK mechanization comes at a cost of higher required Eb/No 

to achieve a given bit error rate. BPSK requires significantly less link 

power at some increased complexity. In BPSK, the transmitter modulates the 

carrier with a signal which generates one of two phases. Typically, the two 

phases are 0 and 180 degrees; or equivalently, the modulating signal is plus 

or minus 1. To distinguish between the two transmitted phases, the receiver 

must generate a coherent phase reference. This process is known as carrier 

recovery and is typically accomplished with a phase-locked-loop (PLL). The 

PLL accounts for the additional complexity of the BPSK receiver. 

Since coherent BPSK is used on the Galileo Probe, one conceptual approach 

for Titan Probe is to modify the Galileo receiver. Galileo uses a microprocessor based digital receiver designed to support a single data rate of 128 

bps. There are two ways the G~Ifl?n iesign can be modified to accomodate 

higher rates: (1) replace the existing components with faster equivalents 

and/or (2) use the Galileo receiver concept, but map as many software functions 

as possible into dedicated hardware (to aleviate microprocessor loading). 

The fastest spacequalified microprocessor (3 MHz clock) applied to the Galileo 

design supports a data rate of only 660 bps. Mapping software functions into 

dedicated hardware yields a maximum supportable data rate of only 1300 bps. 

The Galileo design, therefore, can not be modified to meet the 10 kbps requirement. Though excellent for the low rate Galileo application, the all digital 

receiver is too computationally intensive in its implementation of carrier 

recovery and symbol synchronization loops. 

The alternative BPSK receiver with analog carrier recovery and symbol 

synchronization loops (Figure 5.3-4) i.?ndles a data rate of 10 kbps without 

difficulty. The feedback loop at the frbA;t end recovers the phase of the 

carrier and demodulates the incoming BPSK signal. The acquisition logic and 

lock detector estimates carrier frequcncy to recover the phase of the carrier. 

The acquisition logic sweeps the center frequency of the voltage-controlled 

oscillator (VCO) until the lock detector indicates that the loop is in lock. 

Once the carrier frequency has been located, the data are demodulated. The 

output of the low pass filter (LPF) is the plus or minus one signal which, 

integrated over a bit time, estimates whether a "1" or a "0" has been received. 

As mentioned previously, the data rate decision and symbol synchronization 

functions are common to both the BPSK and NC FSK receivers. 

The doppler performance of the BPSK receiver depends on the ability of the 

carrier recovery loop to track the phase and frequency variation of the carrier. 

Conceptually, the recovery loop is a feedback loop which reacts to the phase of 

the input signal. 

doppler variation. 

tracking ability. 

is dependent upon the functional form of the doppler variation. 

Hence, the l~op must be at least second order to track the 

A third order loop (as chosen for Galileo) provides improved 

The degree to which the loop can track the frequency variation 

Figure 5.3-5 shows a doppler rate profile for a typical mission. With 

an exponential approximation, the doppler limitations of the carrier recovery 

loop can be determined both by simulation and analysis. 

depend on the order and bandwidth of the loop. 

a loop bandwidth of 20 Hz satisfies the constraints on phase degradation, mean 

The performance limits 

For the second order loop, 

174 
t 

time to cycle slip, and loop sweep rate. Simulations show that the second 

order loop can track doppler rates up to 2 Hz/s with a bandwidth of 20 Hz. 

Simulations also show that the third order loop can track doppler rates 

up to 110 Hz/s with a loop bandwidth of only 5 Hz. This is well above any 

mission requirements. In fact, the third order loop can probably be designed 

with a larger loop bandwidth, allowing even higher doppler rates to be tracked. 

The difficulty with using a third order loop is that under certain conditions, 

the loop can become unstable. Consequently, the time to design a third order 

loop is significantly longer than for a second order loop. 

The ability to take radio science measurements is also dependent on the 

order of the carrier recovery loop. With a second order loop, an estimate of 

the carrier frequency is available. A third order loop provides estimates of 

both carrier frequency and frequency rate. 

required to digitize the measuremen+;. 

In any case, D/A converters are 

Tradeof f Summarv 

Table 5.3-2 summarizes the investigation of NC FSK and BPSK receivers. 

Recalling from the mission analysis of section 2.4 that a doppler rate limit 

of 8 Hz/s is not a severe constraint, we conclude the additional hardware 

required by a four filter receiver is not justified. Likewise, the second 

order recovery loop provides adequate doppler rate protection for the BPSK 

receiver. The additional design time required for the third order loop is not 

warranted. 

The tradeoff between NC FSK and BPSY is power efficiency versus implementation complexity. Section 2.4 sbvs that NC FSK supports inadequate time 

at the high data rate. Another consideration is the fact that BPSK provides 

radio science measurements while NC FSK does not. The choice for a baseline 

design is BPSK with a 2nd order carrifsr recovery loop. 

Data Rate Detection 

Although the baseline design has been chosen, there are still two more 

receiver issues which must be pinned down: (1) data rate change detection 

and (2) radio science frequency estimation. The change in data rate must be 

detected so that the symbol synchronization loop can be properly adjusted. 

There are two ways this can be done. The first technique simply uses a clock 

onboard the carrier to indicate the switch from the low data rate to the high 

data rate. Given that the carrier and probe clocks are synchronized just 

before separation, it is possible to compute the uncertainty of timing the 

transition to the high data rate. Assuming a 20 day coast period and clock 

stabilities on the order of 20 parts per million, the timing uncertainty is 

approximately 69 seconds. 

Hypothesis testing yields a higher degree of accuracy. The output of 

the LPF in the BPSK receiver (see Figure 5.3-4) is the sum of the data signal 

plus noise. The objective is to distinguish the high rate from the low rate. 

After integrating the signal plus noise over a certain period of time if the 

output of the integrator exceeds a threshold, then the low rate is present; 

otherwise, the high rate is present (Figure 5.3-5). 

175 
e 

To understand how this works, recall that the high data rate is an 

integral multiple of the low data rate. Hence, an integer number of "fast" 

bits occupy the same time period as one "slow" bit (See Figure 5.3-7). Let 

T be the time for one slow bit. Assume there is a 50 percent bit transition 

probability. If the high data rate is present, then integrating the signal 

over T will yield zero. However, if the low rate is present, integrating 

the signal over T will yield a non-zero value (see Figure 5.3-8). 

to Figure 5.3-6, if no noise is present, then integrating over T will indicate 

which data rate is present. 

Returning 

The addition of noise does not really complicate the problem. Instead 

of being zero or non-zero, the output of the integrator is a gaussian 

process with zero mean, (Figure 5.3-9). If the low rate is present, there 

is some probability that the output of the integrator exceeds the threshold. 

Thus, if the low rate is present, then the output of the integrator will be 

a non-zero mean gaussian process. Tf the high rate is present, then the 

gaussian process will be zero mean. 

Based on noise statistics, the performance of the hypothesis test gives 

a 0.005 second mean time to correctly detect the high data rate. The mean 

time to incorrectly decide high when the low rate is present is 1015 hours. 

The mean time to incorrectly decide low when the high rate is present is also 

d5 hours. 

estimate of the data transition time. 

Thus, hypothesis testing provides an extremely quick and accurate 

Since using an onboard clock to predict the time of rate transition is 

much less complex than hypothesis testing, it was chosen as the baseline. 

Should more accuracy be required, the use of hypothesis testing is a viable 

alternative. 

Frequency Estimation 

The output of the VCD in the 2nd order PLL is a noisy estimate of the 

carrier frequency. By periodically sampling this output and then smoothing 

it, a more accurate estimate of the carrier frequency is obtained. This is 

exactly the technique that was used on Galileo to obtain radio science measurements. Although Galileo used a third order loop, the parameter which sets 

the performance of the estimator is proportional to it. Galileo used a loop 

bandwidth of 8 Hz. This yielded a frequency estimation accuracy of 0.25 Hz, 

assuming the doppler rate is not significant over the integration time of 213 

second. To accomodate the doppler rate, the baseline BPSK receiver requires 

a loop bandwidth of 20 Hz. Hence, we would expect a frequency estimation 

accuracy of about 0.40 Hz and a corresponding wind measurement accuracy of 0.29 

m/s at initial descent and 0.71 m/s at impact (Table 5.5-3). The actual 

performance should be better since the C/No for Titan is higher than for 

Galileo. 

176 
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6. OONCLUSIONS 

This study has demonstrated that some new design concepts can enable a 

Titan Probe to meet new science objectives while maintaining acceptable risk 

and cost. 

Table 6-1 summarizes the results of the major tradeoffs of this study. 

The preceding sections have addressed each of these in detail and defined a 

baseline selection. 

The baseline mission and probe design demonstrates that haze layer sampling 

can be accomplished together with an adequate descent time for atmospheric sampling and a high data rate for surface imaging. However, the atmospheric wind 

uncertainty compromises the data link. If wind bounds can be established the 

analysis in section 2.4 should be repeated. Fortunately, the automated computer 

simulations will facilitate this aiialYSiS during a future study. 

The fixed decelerator configuration with attachment to the Carrier through 

the probe nose offers the lowest cost and risk design while still providing 

adequate drag to facilitate high altitude measurements. 

heat sink decelerator nose with the descent module eliminates any nose separation 

risk. Differential drag enables low risk aft separation of the skirt. The 

only complication inherent to the baseline design is the need to deploy instrument sensors. As instruments become better defined, further study can define 

the deployment mechanisms. 

Integrating a beryllium 

Though somewhat more complicated than an FSK receiver, a BPSK receiver 

is required to support high rate data transmission. A 2nd order loop receiver 

provides adequate tracking considering that antenna geometry constrains the 

mission timing away from periods of peak doppler. Antenna sizing depends on 

wind model and Carrier pointing capability and is an appropriate subject for 

further study refinement. 

The selected redundancy implementation has no single point failures but 

does require redundant and asynchronous instrument data interfaces. After 

instruments are somewhat better defined, further study can compare the impact 

of this approach on instrument complexity as compared with other approaches 

which maintain a single interface with each instrument and use internal spacecraft switching to handle redundancy. 

189 
6.1 Technology Readiness 

Part of this Titan Probe study has been to define any needed technology 

breakthrough so that appropriate enabling research can be undertaken before 

program start in the 1988 time frame. Although a number of further studies 

have been described throughout this report (summarized in table 6.1-l), the 

study has identified no need for enabling technology before hardware go-ahead. 

There are some desirable early activities to be accomplished before the 

development decision. These activities include verification of entry stability 

based on 6 DOF simulations, consideration of less costly heat sink designs than 

beryllium, definition of detailed instrument integration and sensor deployment 

concepts, and refinement of the relay link analysis and design based on a more 

detailed wind model including uncertainties. More detailed decelerator design 

definition and analysis may show advantages of a ribbed or semi monocoque 

design over the baseline honeycomb construction. 

validate candidate heatshield materials. 

A coupon test or tests would 

Finally, the new separation technique of splitting the decelerator and 

allowing differential drag to pull the skirt aft of the descent module nose 

would benefit from a flight demonstration similar to the PAET demonstration 

of parachute extraction. It is suggested that such a demonstration be based 

on an Earth re-entry vehicle and balloon/rocket mechanization simulating the 

desired properties and staging sequence. 

190 
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192 
APPENDIX A - Symbols and Abbreviations 

A/ D 

A-hr 

AS1 

BIU 

BP SK 

BTU 

C 

CCSDS 

CDP 

CDS 

cg 

CM 

C/ No 

dB 

DC 

deglhr 

DO F 

DP SK 

E-20 

Eb/No 

ESM 

EX C 

FIFO 

FSK 

ft 

g 

GC 

HS-376 

HZ 

HYB 

IPIU 

K 

kg 

kH z 

km 

kbps 

Li SO2 

LP F 

JP L 

m 

MHZ 

mm 

N2 

NC 

NeP 

NFR 

Nms 

NMS 

Ns 

PAET 

bPS 

Analog to Digital 

Amper e hour 

Atmospheric Structure Instrument 

Bus Interface Unit 

bit per second 

Binary Phase Shift Keying 

British Thermal Unit 

Ce 1s ius 

Consultative Committee on Space Data Systems 

Command and Data Processor 

Command and Data Subsystem 

center of gravity 

Centimeter 

Carrier-to-Noise ratio 

Decibel 

Direct Current 

degree per hour 

Degree of Freedom 

Differentially Encoded Phase Shift Keying 

Entry minus 20 minutes 

Energy-to-Noise Level ratio 

Elastomeric material 

Exciter 

First in First Out 

Frequency Shift Keying 

feet 

gram, force of gravity at Zqrth sea-level 

Gas Chromatograph 

Hughes Spacecraft Model 376 

Hertz 

hybrid 

Instrument Power Interface Unit 

Kelvin 

ki 1 ogr am 

kildHer tz 

k i lome t er 

kilobits per second 

Lithium Suphur Dioxide 

Low Pass Filter 

Jet P r opu 1s ion Labor at o ry 

meter 

MegaHe r t z 

millimeter 

Diatomic Nitrogen 

non-coherent 

Nephelometer 

Net Flux Radiometer 

Newton meter second 

Neutral Mass Spectrometer 

Newton second 

Planetary Atmospheric Entry Test 

C-3 A- 1 
a 

PCU 

PLL 

PPIU 

P SF 

RF 

RHU 

r Pm 

RS S 

RTG 

RTV 

SP IU 

VCO 

VOIR 

voe 

W 

Pyro Control Unit 

Phase Locked Loop 

Probe Power Interface Unit 

pound per square foot 

Radio Frequency 

Radioisotope Heater Unit 

revolutions per minute 

Root Sum Square 

Radioisotope Thermal Generator 

Room Temperature Vulcanizing 

Subsystem Power Interface Unit 

Voltage-controlled oscillator 

Venus Orbiting Imaging Radar 

Approach Velocity 

Watt 

P 

e 

A- 2 
0 . 

REFERENCES 

Conrath, 85 - Conrath, B.J., Titan atmosphere model presented at the Cassini 

Science Working Group meeting of March 20-22, 1985. 

Flasar, 81 - Flasar, M., Nature, 292, 693, 1981. 

Hughes, 82 - Hughes Aircraft Company, Saturn/Titan Probe Final Report, 

Contract NAS 2-10841, June 1982 

Hunten, 82 - Hunten, D.M., "A Review of Titan," 1982. 

Lindal, 83 - Lindal et al, "The Atmosphere of Titan: An Analysis of the 

Voyager 1 Radio Occultation Measurements," lcarus, 53, 348-363 

(1983). 

Lunine, 83 - Lunine, J.I., Stevenson, D.J., and Yung, Y.L., "Ethane Ocean 

on Titan," Science - 222, 1229 (1983). 

Swenson, 84 - Swenson, B.L., et al, "A New System concept for a Titan Atmospheric Probe," AIAA paper 84-0456, January 1984. 

Tiwari, 81 - Tiwari, S.N., Chow, H., Moss, J.N., "Analysis of Aerothermodynamic Environment of a Titan Aerocapture Vehicle," AIAA 

paper 81-1128, June 1981. 

A- 3 
1. Report No. 2. Govanmmt Accr*on No. 

NASA CR-177394 

4. Title and Subtitle 

STUDY OF NEW SYSTEMS CONCEPTS FOR A TITAN 

ATMOSPHERIC PROBE. Final Technical Report. 

'. Author(r) Doug Bernard, Todd Citron, Robert Drean, 

Scott Lewis, Martin Lo, John McCarthy, Robert IP Staff11 Tina Varoac Mar+v Wnl ff ,, UI 

9. Performing Organization Name and Address 

3. Recipient's CIt~log No. 

5. Report Date 

6. Performing Orglniznion cod. 

April 1986 

8. Morming Orglniration Report No. 

10. Work Unit No. 

Hughes Aircraft Company 

Space and Communications Group 

P.O. Box 92919 Air ort Station 

Los Angeles, CA 98009 , 

Point of Contact: Contract Monitor, Byron L. Swenson, Ames Research Center, 

Moffet Field, CA 94035 (415) 694-5705 

6. Abstract The second Mariner Mk I1 mission scheduled for launch in the mid 1990's 

11. Contract or Gnnt No. 

NAS2-12049 

13. Type of Reporl and Period Covcrd ~ 

2 Sponsoring Agency Name and Address 

National Aeronautics and Space Administration 

Washington, DC 20546 

Contractor Report 

14. Sponsoring Amy Coda 

186-30-2 1 

17. Key Worck (Suggartcd by Authods)) 

Titan Probe, gas sampling, communi- 

*For sale by the National Technical Information Service, !Springfield. Virginia 22161 

18. Distribution Statement 

Unclassified - Unlimited 

A-4 

cations relay, beryllium decelerator, 

surface imaging. Subject 88 

19. Security aaoif. (of thir report) 20. Security c*rrif. (of this plea) 21. No. of Pagm 

Unclassified Unc 1 ass i f i ed 192 

22. Rice'