**2. Background**

In this section, we will discuss the evolution of the NTP concept from its theoretical beginnings in the late 1940s to present-day needs. Much of the knowledge being applied in current NTP system design is drawn from knowledge gained through a series of experimental programs beginning in the late 1950s and running through the early 1970s. However, current interests have resulted in new materials testing based on experience gleaned from earlier work combined with modern materials performance data and testing methods. A major focus of the NASA Space Nuclear Propulsion Program is in reviving NTP fuel fabrication techniques and design knowledge [6]. Hence, an overview of the history of NTP is appropriate before moving on to current testing programs. These programs provide significant insight for current research and testing programs. But first, let us revisit the motivation for nuclear thermal propulsion over chemical engines for extraterrestrial propulsion.

#### **2.1 Advantages of nuclear thermal propulsion for interplanetary travel**

The efficiency of a rocket engine design is commonly measured in terms of specific impulse. One can think of *I*sp as the miles per gallon or kilometers per liter for a car. The larger the *I*sp, the more efficient the engine. Mathematically, specific impulse is defined as the total engine thrust integrated over time per unit weight of the propellant; here, weight is defined as measured on Earth (e.g., N, or, historically, lbf) [7]). Thrust is defined as:

$$F\_{\text{thrust}} = \upsilon\_{\varepsilon} \cdot \dot{m} \tag{1}$$

where:

*F*thrust is the force (thrust) exerted by the propellant (N),

*v*<sup>e</sup> is the exit velocity of the exhaust propellant (m/s) relative to the nozzle, and *m*\_ , or d*m=*d*t*, is the mass flow rate of the propellant (kg/s).

The total impulse (*I*) of a rocket for time *t* is defined as the thrust integrated over the total time of operation (*burn time* in a chemical rocket, or time at power in an NTP engine):

$$I(t) = \int\_0^t F\_{\text{thrust}}(\tau) \, \text{d}\tau = \int\_0^t v\_\text{e} \cdot \dot{m} \, \text{d}\tau = m\_\text{ex} \cdot v\_\text{e} \tag{2}$$

Here, we have assumed that *v*<sup>e</sup> is constant, and *m*ex is the total mass expelled over the time of operation, *m*(0) � *m*(*t*).

Specific impulse is defined as the total impulse divided by the weight *W* of the propellant on Earth, i.e.:

$$W = m\_{\text{ex}} \cdot \lg\_o \tag{3}$$

*Nuclear Thermal Propulsion DOI: http://dx.doi.org/10.5772/intechopen.103895*

Hence,

$$I\_{\rm sp} = \frac{m\_{\rm ex} \cdot v\_{\rm e}}{W} = \frac{m\_{\rm ex} \cdot v\_{\rm e}}{m\_{\rm ex} \cdot \mathcal{g}\_{\rm o}} = \frac{v\_{\rm e}}{\mathcal{g}\_{\rm o}}\tag{4}$$

Rearranging the expressions in Eq. (1), in terms of *v*<sup>e</sup> and replacing *v*<sup>e</sup> in Eq. (4), we arrive at a more useful definition for *I*sp:

$$I\_{\rm sp} = \frac{F\_{\rm thrust}}{\dot{m} \cdot \mathbf{g}\_{\rm o}} \tag{5}$$

Eq. (5) shows that the *I*sp is the ratio of thrust to the product of the mass flow rate times the constant *g*o. In this form, it is clear that the *I*sp can be interpreted as the time (in s) over which 9.81 kg (or one Newton of weight on Earth) of propellant can produce one Newton of thrust. The larger the *I*sp, the longer the engine is able to operate with a given mass of fuel.

A pioneer in rocketry theory in the early 1890s, Konstantin Tsiolkovsky [8] derived a number of important relationships, including Eq. (6), which is used heavily in rocket design and is known as the ideal exhaust velocity equation, relating gas properties to the exit velocity of the propellant:

$$w\_e^2 = \frac{\frac{2kRT\_c}{M} \left(1 - \left(\frac{p\_e}{p\_c}\right)^{\frac{k-1}{k}}\right)}{k-1},\tag{6}$$

where:

*k* is the ratio of the specific heat at constant pressure (*cp*) to specific heat at constant volume (*cv*) for the propellant (i.e., *k* ¼ *cp=cv*),

*R* is the universal gas constant,

*Tc* is the reactor core exit temperature for NTP, or the combustion chamber temperature for a chemical engine,

*M* is the molecular weight of the propellant,

*p*<sup>e</sup> is the nozzle exit pressure, and

*p*<sup>c</sup> is the core exit (or combustion chamber) pressure.

*R* is a fundamental physical constant, *k* does not vary significantly between different gases (typically between 1.1 and 1.5) and *Tc*, *p*c, and *p*<sup>e</sup> depend on the engine specifications. Assuming that k, Tc, pc, and pe are known and identical between NTP and a chemical rocket, we can combine them into the constant *C*:

$$I\_{\rm sp} = \frac{C}{\sqrt{M}},\tag{7}$$

For rockets that use the chemical reaction of H2 and O2 to produce energy and release high temperature H2O, the atomic mass of the propellant, *M*, is 18 g/mole. NTP engines use high energy H2 (*M* ¼ 2 g/mole) that is discharged from a high temperature core. Comparing the theoretical specific impulses,

$$\frac{I\_{\rm sp}(\rm H\_2)}{I\_{\rm sp}(\rm H\_2O)} = \frac{\mathbf{C\_{H\_2O}}/\sqrt{2}}{\mathbf{C\_{H\_2}}/\sqrt{18}} = \sqrt{18/2} = \text{3.}\tag{8}$$

This assumes that CH2O is equal to CH2 (they are similar but not equal). Thus, based on ideal gas assumptions, H2 could provide three times the *I*sp of H2O as a propellant. However, in reality, gas is not ideal and the value of CH2O is not equal to CH2 , as the value of *k* is not the same for the two fluids. In addition, for NTP, the most significant challenge is in obtaining a high exit temperature from the core. This requires nuclear fuel materials to be able to quickly rise to and maintain very high temperatures. Chemical engine combustion chamber temperatures are on the order of 3500 K; NTP efforts currently aim for a temperature of approximately 2700–3000 K based on material limits. Together, these facts somewhat reduce the advantage in *I*sp from the ideal value of 3 to a ratio closer to 2. Nevertheless, with the *I*sp for a H2/O2 engine is on the order of 450 s, while for NTP, it would be on the order of 900 s. Hence, there remains a clear advantage to the use of an NTP engine. Heating H2 to significant outlet temperatures can be achieved using a nuclear reactor.

This advantage was recognized in the 1940s. An NTP-propelled spacecraft could significantly reduce the travel time to Mars as compared to conventional engines [9]. This would reduce astronaut radiation exposure, as well as the impact of the long-term microgravity environment.

Note that NTP engines are not intended for liftoff from Earth; they are not designed to provide sufficient thrust for launch. Chemical engines would be used to lift a full vessel (in parts) to low earth orbit (LEO), from where the vessel would be assembled and an NTP-propelled mission would be launched.

In the late 1960s, the well-known pioneer of modern rocketry, Wernher von Braun, then the director of the NASA Marshall Space Flight Center, advocated for a mission to Mars. Under his plan, NASA would launch a Mars mission in November 1981 (based on favorable planetary alignment), and land on the red planet by August 1982. Von Braun explained that "although the undertaking of this mission will be a great national challenge, it represents no greater challenge than the commitment made in 1961 to land a man on the moon" [10]. In the following subsection, we will briefly visit early NTP research and the Nuclear Engine for Rocket Vehicle Application (NERVA) rocket engines that von Braun had envisioned would take men to Mars.

#### **2.2 History of nuclear thermal propulsion**

The concept of nuclear thermal propulsion was first publicly published by the Applied Physics Laboratory in 1947 [11]. Development of NTP systems began at Los Alamos Scientific Laboratory (LASL) in 1955 as Project Rover, under the auspices of the Atomic Energy Commission (AEC). NASA was formed in 1958 in response to Russia's launch of Sputnik and the beginning of the space race, and took over the Rover project with continued collaboration with LASL and the AEC [12]. Rover later became a civilian project within NASA and was reorganized to perform research directed toward producing a nuclear powered upper stage for the Saturn V rocket. In 1961, the NERVA program was formed by NASA to develop a nuclear thermal rocket engine. The program designed, assembled, and tested 20 nuclear rocket engines through a number of experimental series, including the KIWI, PEWEE, PHOEBUS, TF, and NRX reactors. These ground-based test reactors used solid fuel, based on advanced graphite materials, and were thermal spectrum reactors. The NRX-XE rocket reactor performed 28 burns with more than 3.5 h of operation [6], demonstrating the ability to operate and restart with the high performance requirements needed for use in an NTP system.

A Nerva-type engine concept is depicted in **Figure 1**. The fuel is manufactured as solid hexagonal blocks, with holes drilled through for hydrogen flow to cool the core. Multiple elements are assembled to create the core, with criticality control through the use of control drums with a poison plate on one side of the cylindrical drum, much as has been used at the Advanced Test Reactor (ATR) for over 50 years [14]. Minimal excess reactivity is needed as the total core lifetime will be on the order of hours, and will only operate for times on the order of an hour or less

#### **Figure 1.**

*Reactor core cross section for a ROVER-type NTP engine (left) and a cutaway of a fuel assembly cluster (right) [13].*

resulting in minimal xenon buildup. These reactors were fueled using high-enriched uranium (HEU) in excess of 90% 235U.

Both the Rover and NERVA research focused on a fuel form consisting of a graphite matrix with dispersed fuel (GMWDF). Graphite fuel compacts were used with various fuel types, including UO2 and UC2 fuel particles, and as (U,Zr)C<sup>1</sup> graphite composite. The three fuel forms used with the GMWDF compact are [15]:


GMWDF compacts lead to a hard thermal spectrum [16–18]. Early designs exclusively used the graphite matrix as a moderator, but later designs starting with the PEWEE 1 experiment included ZrH sleeves in tie rods to increase the moderation ratio and reduce the core size [17]. The main issue with GMWDF compacts is that hot hydrogen corrodes the graphite matrix if they come into direct contact [15]. Therefore, all GMWDF compacts used coatings to protect the graphite matrix. The coatings must match the thermal expansion of the matrix closely to avoid excessive cracking. While still remaining a concern at the conclusion of project NERVA, corrosion rates were reduced by more than a factor of 10 [17]. GMWDF was used in

<sup>1</sup> This notation denotes a solid solution where C sits on one lattice and U and Zr share the second lattice.

the shape of fuel plates (KIWI-A) and cylindrical fuel elements in a graphite module (e.g., KIWI-A<sup>0</sup> and KIWI A3), but most often as hexagonal fuel elements connected via tie tubes [17], as illustrated in **Figure 1**.

While not used in most early testing, CERamic-METallic (CERMET) fuels were evaluated during the NERVA program. The technology was too new and not well understood in the early 1960s, but was being investigated in parallel to the NERVA experiments. CERMET compacts consist of ceramic fuel particles embedded in a refractory metal matrix [19, 20]. The choice of matrix and fuel material influences thermal stability, thermal conductivity, structural integrity, and neutronics performance of the CERMET compact. Concurrent with the NERVA program, ANL and General Electrics (GE) developed separate CERMET NTP concepts. In a simplistic sense, CERMET fuels are particles of ceramic fuel (i.e., UO2 or UN) encapsulated in a metal matrix, typically, but not limited to, tungsten, rhenium, or molybdenum. The research conducted by ANL and GE included the development and testing of the CERMET fuel and the design of the ANL-200, ANL-2000 [20, 21], and the GE 710 reactors [21, 23]. These CERMET programs focused entirely on HEU fuel kernels and fast reactor concepts. In contrast to GMWDF, the GE CERMET concepts did not undergo prototypical irradiation testing, nor did either concept undergo engine testing. Therefore, prior to the twenty-first century, the technology readiness of CERMET compacts trailed that of the GMWDF compacts.

The matrix material of a CERMET usually makes up about 30–60% of the compact volume [23], so its properties are both neutronically and structurally important. The ANL and GE programs focused mostly on natural tungsten as matrix material [20]. Among the available matrix materials, tungsten provides the largest fracture strength and temperature stability [6]. However, tungsten is brittle at low temperatures, causing issues with cracking. All isotopes of tungsten have strong ð Þ *n*, *γ* resonances between 1 eV and 5 keV, thereby making tungsten neutronically challenging, except for fast reactor applications.

Fuel kernels also make up a significant fraction of the volume, so the materials properties and performance must be evaluated. Some work was performed in this area under the GE and ANL engine design programs for UO2 and UN fuel types, as described below:


The GE experiments were at temperatures significantly lower than the NTP requirements, but provided much data on materials behavior and failure mechanisms [20, 26]. ANL focused on the production of CERMET fuels; different fuel fabrication procedures were employed with mixed success. Non-nuclear testing of samples was performed in flow loops of hydrogen heated to 2770°C to understand the fuel loss rates. Nuclear tests on the ANL CERMET samples were run in the Transient Reactor Test Facility (TREAT) located at Idaho National Laboratory
