R.S. Heller

Massachusetts Institute of Technology, Cambridge, Massachusetts

G.A. Landis and A.F. Hepp

Glenn Research Center, Cleveland, Ohio

A.J. Colozza

QinetiQ North America Corporation, Cleveland, Ohio

Heated-Atmosphere Airship for the Titan

Environment: Thermal Analysis

NASA/TM—2012-217639

July 2012
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R.S. Heller

Massachusetts Institute of Technology, Cambridge, Massachusetts

G.A. Landis and A.F. Hepp

Glenn Research Center, Cleveland, Ohio

A.J. Colozza

QinetiQ North America Corporation, Cleveland, Ohio

Heated-Atmosphere Airship for the Titan

Environment: Thermal Analysis

NASA/TM—2012-217639

July 2012

National Aeronautics and

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Glenn Research Center

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NASA/TM—2012-217639 1 

Heated-Atmosphere Airship for the Titan Environment: 

Thermal Analysis 

R.S. Heller 

Massachusetts Institute of Technology 

Cambridge, Massachusetts 02139 

G.A. Landis and A.F. Hepp 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

A.J. Colozza 

QinetiQ North America Corporation 

Cleveland, Ohio 44135 

Abstract 

Future exploration of Saturn’s moon Titan can be carried out by airships. Several lighter-thanatmosphere gas airships and passive drifting heated-atmosphere balloon designs have been studied, but a 

heated-atmosphere airship could combine the best characteristics of both. This work analyses the thermal 

design of such a heated-atmosphere vehicle, and compares the result with a lighter-than-atmosphere 

(hydrogen) airship design. A design tool was created to enable iteration through different design parameters 

of a heated-atmosphere airship (diameter, number of layers, and insulating gas pocket thicknesses) and 

evaluate the feasibility of the resulting airship. A baseline heated-atmosphere airship was designed to have a 

diameter of 6 m (outer diameter of 6.2 m), three-layers of material, and an insulating gas pocket thickness of 

0.05 m between each layer. The heated-atmosphere airship has a mass of 161.9 kg. A similar mission 

making use of a hydrogen-filled airship would require a diameter of 4.3 m and a mass of about 200 kg. For a 

long-duration mission, the heated-atmosphere airship appears better suited. However, for a mission lifetime 

under 180 days, the less complex hydrogen airship would likely be a better option. 

Introduction 

Titan, the largest moon in the solar system, and the only moon with an atmosphere, is a high priority 

target for future science missions. The Cassini mission to Saturn landed a probe, Huygens, on the surface of 

Titan, but this probe only had a short lifetime on the surface, and had no mobility after landing. Future 

missions to Titan are desired with the capability to explore the nature of the moon and its environment in 

greater detail. 

Several options have been considered as follow-on to the Huygens-Cassini mission. Of these, the main 

consideration for a Titan mission has been the type of vehicle used to collect data. A follow-on joint 

proposal between National Aeronautics and Space Administration (NASA) and European Space Agency 

(ESA) has mentioned three desired vehicles. One was an orbiter similar to the Cassini orbiter. Another was a 

lake lander similar to Huygens, but with the emphasis of studying a lake instead of the atmosphere. The 

final vehicle was an atmospheric balloon. This proposed mission was called the Titan Saturn System 

Mission (TSSM) (TSSM 2009). The TSSM atmospheric balloon was planned to passively collect 

atmospheric data at an altitude of about 10 km. The power source of the electronics as well as the heat 

source for the balloon would come from a Radioisotope Thermoelectric Generator (RTG) (TSSM 2009). 

However, the TSSM would not be well suited to study the organic molecules that have accumulated around 

the moon. This has led to proposals to use controlled aerial vehicles that could move around the moon and 

land at various sites to collect samples. Several types of aerial vehicles could be used. These include: 

balloons, airships, vertical takeoff, and landing vehicles (such as helicopters), airplanes, gliders, and rockets. 
NASA/TM—2012-217639 2 

Previous Work 

Two earlier studies (Colozza et al. (Colozza 2003) and Lorenz (Lorenz 2000a)) compared different 

aerial vehicles for use on Titan. Wright et al. (Wright 2004) concluded that airships are the best for initial 

missions, with vertical takeoff and landing vehicles being considered for future missions due to the 

improved landing and takeoff ability. Lorenz et al. (Lorenz 2005) proposed an airship that would float and 

act as a balloon if propulsion were to fail either permanently or temporarily, providing an inherent 

redundancy to the flight capability for the mission. Designs for airplanes and vertical takeoff and landing 

vehicles would not have this capability, and hence failure of the propulsion system would likely be 

mission ending. Lorenz (Lorenz 2000a) considered a reference concept of a vertical takeoff and landing 

vehicle, selected for ease of landing at desired sites. Of these two design possibilities, the airship was the 

vehicle type analyzed in this report. 

Airships are very similar to balloons, but instead of passively relying on winds to move (such as 

winds used for drift discussed in Lorenz et al. (Lorenz 2010)), they have an active propulsion system. 

Various airship designs are also commonly referred to as dirigibles, Zeppelins, and blimps. They could 

either use lighter-than-atmosphere gas or heated-atmosphere for lift. Studies have been conducted to 

understand the performance of Titan airships using lighter-than-atmosphere gas as the lifting gas. A paper 

by Duffner et al. (Duffner 2007) analyzes a hydrogen gas airship with a rover appendage for surface 

studies. Lorenz (Lorenz 2000b) discussed how an airship could maintain continuous flight with dutycycled propellers for conserved power propulsion. The paper suggested that the surface could be studied 

by drop sondes, or by possibly landing the airship. Elfes et al. (Elfes 2005) and the ESA Titan Saturn 

System Mission report (TSSM 2009) proposed that excess heat from an RTG could be used for lift. 

Duffner et al. (Duffner 2007), Lorenz (Lorenz 2000b), and Wright et al. (Wright 2004) required an RTG 

for the electrical power needs. Additionally, the TSSM planned to use an RTG heated-atmosphere 

balloon; however, an RTG heated-atmosphere airship could make use of the waste heat of an RTG, saving 

launch weight by providing the initial lifting gas. It would also have the maneuverability of an airship for 

possibly greater science return. 

The material for the lifting gas envelope is also an important consideration. This is due to the 

atmospheric conditions as well as the necessary constraints for a working airship. The material must be 

nonporous, foldable for storage to Titan, resistant to tearing, and maintain these properties at cryogenic 

temperatures. Hall et al. (Hall 2008) proposed a combination of polyester and Mylar for an effective 

cryogenic balloon material. Their paper demonstrated that this material satisfactorily meets the necessary 

criteria discussed. 

Radioisotope Generator 

A Radioisotope Thermoelectric Generator (RTG), uses a nuclear isotope heat source and a 

thermoelectric (thermal-to-electrical) transducer to convert heat produced from radioactive decay to 

electricity. A commonly used fuel is Pu238, which has a half-life of about 87 years, and therefore can 

support a mission for several decades. The plutonium is packaged into a standard general purpose heat 

source (GPHS) block for use in spacecraft applications. Each block produces 250 W of thermal power at 

its beginning of life and has a mass of 1.44 kg. The primary disadvantage of implementing RTGs is the 

low level of electrical power that they supply. Large power requirements would necessitate many RTG 

units, which would consequently increase the system’s complexity, mass and cost. RTGs have been used 

in many space vehicles and satellites such as the Voyager spacecraft discussed in Freedman et al.

(Freedman 1966). This demonstrates that RTGs are a viable option so long as large power requirements 

are not needed. Higher-efficiencies can be achieved through the use of a dynamic conversion technique 

instead of a thermoelectric element, for example, a Stirling converter. This has been demonstrated in the 

form of the Advanced Stirling Radioisotope Generator (ASRG) (Wong 2008), which can achieve 

conversion efficiencies near 30 percent. 
NASA/TM—2012-217639 3 

For the following analysis the Multi-Mission Radioisotope Thermoelectric Generator (MMRTG) was 

assumed as the power and thermal source. This RTG is used on the Mars Science Laboratory. At the time 

of arrival at Titan, it will produce about 100 W of electrical power for about 1700 W of thermal power 

(TSSM 2009). The mass of the MMRTG along with a battery system is assumed to be 100 kg. 

Buoyancy Calculations for Heated Gases 

A balloon or airship for operation on Titan achieves its lift by use of a confined envelope of gas that 

is lighter than the surrounding atmosphere. This can be achieved either by using a lifting gas of lower 

molecular weight than the ambient (optimally hydrogen), or by filling the envelope from the surrounding 

atmosphere and heating the interior gas to a higher temperature, thus reducing its density (“heated 

atmosphere”). The heated-atmosphere approach has the advantage of not requiring fill gas to be brought 

from Earth. Hence, it is relatively insensitive to the gas leakage rate, and can achieve long duration flight 

as long as a heat source is available, without replenishing the fill gas. However, the approach required the 

gas to be at a higher temperature than the surrounding atmosphere, and hence the lift achieved depends on 

the thermal design. This work is intended to calculate the thermal analysis of such a design, in order to 

compare the heated-atmosphere and lifting-gas approaches for lift in the Titan environment. 

In order for an airship to float in an atmosphere, the buoyancy force must overcome the force of 

gravity. Buoyancy comes from the Archimedes principle that the weight of fluid displaced equals upward 

buoyancy force. Because the airship uses heated atmosphere, the density of the internal gas can be closely 

approximated from the ideal gas law. For the buoyancy calculation, we can approximate the airship as a 

sphere with internal diameter D, calculated from the sphere’s volume. Thus, Equation (1) relates 

buoyancy to the mass of the airship, where P is the gas pressure, Rs is the specific gas constant (J/kgK), 

T is the temperature of the gas, atm is the density of the atmosphere, and m is the total mass that can be 

lifted. In order for the vehicle to achieve flight, its mass must be equal to or less then this mass. 

3

atm 6 s

D P

m

R T

  

     

 (1) 

Thermal Balance 

The buoyancy (Eq. (1)), contains three unknown variables: the airship diameter, the average 

temperature inside the airship, and the mass that can be lifted. The pressure within the envelope (P) is 

held near atmospheric pressure in order to minimize the stress on the envelope material. The relationship 

between the airship diameter and the average inside temperature is then calculated by analyzing the heat 

transfer from the warm environment inside of the airship to the colder environment of Titan’s atmosphere. 

The heat source (from the MMRTG) provides a constant flux of waste heat to the gas within the envelope. 

The heat loss to the environment is proportional to the surface area of the gas envelope: the larger the 

diameter, the lower the inside temperature will be. Figure 1 shows the thermal circuit diagram for a 

three-layer example. 
NASA/TM—2012-217639 4 

The equivalent thermal resistor for a three-layer airship is simply the sum of the resistors from 

Figure 1. The equation for the internal temperature is given by Equation (2), where Tw is the temperature 

within the airship, and Tt is the temperature of Titan’s atmosphere. 

T T QR R R R R R R w t         1234567  (2) 

The temperature of Titan’s atmosphere is assumed to be 95 K and the thermal load Q is an input 

variable set by the amount of isotope used in the MMRTG. The value of each of the thermal resistors 

must be determined. The thermal resistance for conduction through the material layers is given by 

Equation (3) and the thermal resistance for convection from the inner and outer surfaces is given by 

Equation (4), where ro is the outer radius, ri is the inner radius, and rs is the radius of the surface where 

convection is taking place. 

cond 4

o i

i o

r r

R

kr r

  

 (3) 

conv 2

1

4 s

R

hr  

 (4) 

Two quantities are needed to solve for the thermal resistance in the conduction case: the thickness of 

the material and thermal conductivity of the material. For the convection cases, only the heat transfer 

coefficient (h) needs to be solved for. 

Iteration Program 

The effects of the three main design parameters (the diameter, number of layers, and the thickness of 

the insulating gas pockets) on the mass of the vehicle as well as the mass that can be lifted were 

evaluated. An iteration program was utilized since thermal resistance depends on the temperature, and the 

temperature affects the thermal resistance. Table 1 shows the assumed values and constants that were 

used in the iteration program. 

The wind speed here is the relative speed difference between the airship and the actual wind velocity. 

For example, a completely unpowered airship would move with the wind and would have a relative wind 

speed of zero. A temperature of 95 K was assumed for the atmosphere, representative of the highest (i.e., 

worst case) atmospheric temperature. Since the lower the atmospheric temperature, the greater its density 

and therefore buoyancy force, operation at a lower temperature would result in the ability to carry a 

slightly higher payload mass. 

TABLE 1.—CONSTANT PROPERTIES WITH ASSUMED VALUES 

Property Value 

Envelope material thickness 0.00003 m 

Cryogenic envelope material weight (Hall 2008) 94 g/m2

Estimated thermal conductivity of the cryogenic envelope material 0.4 W/mK 

Titan pressure (Friedlander 1986) 160 kPa 

Titan gravity 1.352 m/s2

Titan temperature 95 K 

Wind speed 5 m/s 

Specific gas constant for the atmospheric nitrogen 297 J/kgK 

RTG heat (TSSM 2009) 1700 W 

Mass of RTG and Battery System (Duffner 2007) 100 kg 

Mass of instruments (Duffner 2007) 28.9 kg 

Prandlt number of atmosphere (TSSM 2009) 0.796 
NASA/TM—2012-217639 5 

Results 

A number of parameters, including diameter, insulating gas pocket thickness, and number of layers as 

well as the lifting mass, deadweight, and extra payload, can be adjusted to affect the design output. In 

order to narrow down the list of choices, consideration was given to difficulty. For an example, larger 

numbers of layers are more difficult to build, and a heavier deadweight means more fuel would be 

expended for launch. With these considerations in mind, a satisfactory set of feasible parameters was 

chosen for the baseline airship. The parameters are presented in Table 2. 

Figures 2 and 3 show the effects of modifying the number of layers and the thickness of the insulating 

gas pocket between the layers, respectively. It is seen that there is a significant increase in payload mass 

(over 250 kg) by adding two extra envelope layers compared to using a single layer, but with more 

additions, there is only a small increase in the payload mass. Figure 3 shows that, as the insulating gas 

pocket thickness increases, the payload mass increases. However, it should be noted that as the insulating 

gas pocket thickness increases, the outer diameter also increases. 

The heated-atmosphere airship was designed primarily to interact with and explore the surface of 

Titan. As the airship’s altitude increases, the payload mass capability diminishes until the airship reaches 

an altitude where neutral buoyancy takes place. From the surface to an altitude of 20 km, both the 

temperature and the density of the atmosphere decrease approximately linearly. At the 20 km altitude, the 

temperature was estimated to be 78 K and the density of the atmosphere was estimated to be 2 kg/m3

(Jennings 2011, Justus 2004). Figure 4 shows the effect that the operating altitude has on the payload for 

different internal diameters. The baseline airship with a diameter of 6 m would be able to reach a 

maximum altitude of 6 km. Different designs would be required to reach higher altitudes, such as an 

airship with a larger diameter. An airship with an internal diameter of about 8 m is shown to have the 

highest payload mass at an altitude of 6 km (about 75 kg). At 6 km, the range of feasible airships with 

three-layers and an insulating gas pocket thickness of 0.05 m are those with an internal diameter between 

5.2 and 10.1 m. 

TABLE 2.—BASELINE CASE DESIGN 

Parameter Value 

Inner diameter 6 m 

Outer diameter 6.2 m 

Number of layers 3 

Insulating gas pocket thickness 0.05 m 

Average internal temperature 187.7 K 

Deadweight 161.87 kg 

Lift mass 343.5 kg 

Extra payload mass 181.63 kg 

 
NASA/TM—2012-217639 6 

Effects of the Heat Source on Sizing 

As mentioned previously, an ASRG unit could be substituted for the MMRTG unit used for the 

calculations in preceding sections, increasing the amount of electrical power produced from a given 

amount of isotope. Multiple ASRG units could be combined to add additional thermal or electrical power. 

Each ASRG unit has a mass of about 20 kg and a thermal output of about 500 W (Wong 2008). Figure 5 

shows the effect that additional ASRG units would have on the resulting payload mass. Increased payload 

can be achieved by increasing the number of ASRG units used. For example, with an internal diameter of 

10 m, 20 ASRG units could support about 1200 kg of payload more than a single ASRG unit could. 

However, the additional 19 ASRG units weigh 380 kg. The additional weight would make it much more 

expensive to launch from Earth. 
NASA/TM—2012-217639 7 

Conclusion 

While a hydrogen airship is an inherently simpler design for several missions, the thermal airship is 

superior for others. Table 3 presents the results of the major design parameters for both the heatedatmosphere airship and a comparable hydrogen airship. 

The hydrogen airship has a heavier mass (~200 kg) compared to the heated-atmosphere airship 

(~162 kg) in the deployed state. Additionally, the heated-atmosphere airship has a longer possible flight 

life of up to 154 years compared to 180 days for the hydrogen airship. This was determined by the time it 

takes the MMRTG unit to decay to the point of neutral buoyancy at the surface, compared to the leak rate 

for the hydrogen airship. For a desired long-term mission, the heated-atmosphere airship appears better 

suited. However, for a desired mission under 180 days, the less complex hydrogen airship would likely be 

a better option. 

TABLE 3.—MAJOR BASELINE DESIGN PARAMETER RESULTS 

Parameter Heated-atmosphere airship Hydrogen airship 

Mass (kg) 162 200 

Diameter (m) 6 4.3 

Number of layers 3 1 

Insulating gas pocket thickness (m) 0.05 N/A 

References 

Colozza, A., Landis, G., and Lyons, V. (2003), “Overview of Innovative Aircraft Power and 

Propulsion Systems and Their Applications for Planetary Exploration,” NASA/TM—2003-212459, 

July 2003. 

Duffner, J., Liu, M., Mandy, C., Panish, R., and Landis, G.A. (2007), “Conceptual Design of an 

Airship Mission to Titan,” AIAA Space Conference & Exposition, Long Beach, California, September 18-

20, 2007. 

Elfes, A., Montgomery, J.F., Hall, J.L., Joshi, S.S., Payne, J., and Bergh, C.F. (2005), “Autonomous 

Flight Control for a Titan Exploration Aerobot,” 8th International Symposium on Artificial Intelligence, 

Robotics and Automation In Space, Munich, Germany, September 5-9, 2005. 

Freedman, S., Friedman, J., and Schmieder, J. (1966), “RTG Parametric Study Report for the RTG 

Study Voyager Task C (U),” JPL Report, JPL-C00458, September 29, 1966. 

Friedlander, A.L. (1986), “Buoyant Station Mission Concepts for Titan Exploration,” 

Acta Astronautica, Vol. 14, pp. 233-242, 1986. 

Hall, J.L., Jones, J.A., Kerzhanovich, V.V., Lachenmeier, T., Mahr, P., Pauken, M., Plett, G.A., 

Smith, L., Van Luvender, M.L., and Yavrouian, A.H. (2008), “Experimental Results for Titan Aerobot 

Thermo-Mechanical Subsystem Development,” Advances in Space Research, Vol. 42, Issue 10, pp. 1641-

1647, November 2008. doi:10.1016/j.asr.2007.02.060. 

Jennings, D.E., Nixon, C.A., and Cottini, V. (2011), “Seasonal Changes in Titan’s Surface 

Temperatures,” Fifth Workshop on Titan Chemistry – Observations, Experiments, Computations, and 

Modeling, Kauai, Hawaii, 12 April 2011. 

Justus, C.G., Duvall, A., and Keller, V.W. (2004), “Engineering-level model atmospheres for Titan 

and Mars,” Proceeds of the International Workshop Planetary Probe Atmospheric Entry and Descent 

Trajectory Analysis and Science, February 2004. 

Lorenz, R.D. (2000a), “Optimizing Science Return from Titan Aerial Explorers,” Aerospace 

Conference, IEEE, Vol. 7, pp. 493-499, Big Sky, MT, 2000. doi: 10.1109/AERO.2000.879317. 

Lorenz, R.D. (2000b), “Post-Cassini Exploration of Titan: Science Rationale and Mission 

Concepts,” Journal of the British Interplanetary Society, Vol. 53, pp. 218-234, 2000. 
NASA/TM—2012-217639 8 

Lorenz, R.D., Lunine, J.I., Zimmerman, W. (2005), “Post-Cassini Exploration of Titan: Science 

Goals, Instrumentation and Mission Concepts,” Advances in Space Research, Vol. 36, Issue 2, pp. 281-

285, 2005. doi:10.1016/j.asr.2005.03.080. 

Lorenz, R.D., Tokano, T., and Newman, C.E. (2010), “Winds and Tides of Ligeia Mare, With 

Application to the Drift of the Proposed Time (Titan Mare Explorer) Capsule,” Planetary and Space 

Science, 2010. doi:10.1016/j.pss.2010.12.009. 

“TSSM In Situ Elements,” (2009), ESA Contribution to the Titan Saturn System Mission, 

ESA-SRE(2008)4, February 12, 2009. 

Wong W.A., Wood J.G., and Wilson K. (2008), “Advanced Stirling Convertor (ASC)—From 

Technology Development to Future Flight Product,” Space Technology and Applications International 

Forum (STAIF–2008), Albuquerque, New Mexico, February 10–14, 2008; NASA/TM—2008-215282.

Wright, H.S., Levine, J.S., Croom, M.A., Edwards, W.C., Qualls, G.D., and Gasbarre, J.F. (2004), 

“Measurements from an Aerial Vehicle: A New Tool for Planetary Exploration,” Proceedings of SPIE, 

the International Society for Optical Engineering, Vol. SPIE-5660, No. 23, 2004. 
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14. ABSTRACT

Future exploration of Saturn’s moon Titan can be carried out by airships. Several lighter-than-atmosphere gas airships and passive drifting 

heated-atmosphere balloon designs have been studied, but a heated-atmosphere airship could combine the best characteristics of both. This 

work analyses the thermal design of such a heated-atmosphere vehicle, and compares the result with a lighter-than-atmosphere (hydrogen) 

airship design. A design tool was created to enable iteration through different design parameters of a heated-atmosphere airship (diameter, 

number of layers, and insulating gas pocket thicknesses) and evaluate the feasibility of the resulting airship. A baseline heated-atmosphere 

airship was designed to have a diameter of 6 m (outer diameter of 6.2 m), three-layers of material, and an insulating gas pocket thickness of 

0.05 m between each layer. The heated-atmosphere airship has a mass of 161.9 kg. A similar mission making use of a hydrogen-filled 

airship would require a diameter of 4.3 m and a mass of about 200 kg. For a long-duration mission, the heated-atmosphere airship appears 

better suited. However, for a mission lifetime under 180 days, the less complex hydrogen airship would likely be a better option.

15. SUBJECT TERMS

Titan; Titan atmosphere; Airships; Methane; Hydrogen; In situ resource utilization; Mission planning 

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