Thomas Haag

Glenn Research Center, Cleveland, Ohio

Mechanical Design of Carbon Ion Optics

NASA/TM—2005-213977

October 2005

AIAA–2005–4408
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Thomas Haag

Glenn Research Center, Cleveland, Ohio

Mechanical Design of Carbon Ion Optics

NASA/TM—2005-213977

October 2005

National Aeronautics and

Space Administration

Glenn Research Center

Prepared for the

41st Joint Propulsion Conference and Exhibit

cosponsored by the AIAA, ASME, SAE, and ASEE

Tucson, Arizona, July 10–13, 2005

AIAA–2005–4408
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NASA/TM—2005-213977 1

Mechanical Design of Carbon Ion Optics 

Thomas Haag 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

Carbon Ion Optics are expected to provide much longer thruster life due to their resistance to sputter erosion. 

There are a number of different forms of carbon that have been used for fabricating ion thruster optics. The 

mechanical behavior of carbon is much different than that of most metals, and poses unique design challenges. In 

order to minimize mission risk, the behavior of carbon must be well understood, and components designed within 

material limitations. Thermal expansion of the thruster structure must be compatible with thermal expansion of the 

carbon ion optics. Specially designed interfaces may be needed so that grid gap and aperture alignment are not 

adversely affected by dissimilar material properties within the thruster. The assembled thruster must be robust and 

tolerant of launch vibration. The following paper lists some of the characteristics of various carbon materials. 

Several past ion optics designs are discussed, identifying strengths and weaknesses. Electrostatics and material 

science are not emphasized so much as the mechanical behavior and integration of grid electrodes into an ion 

thruster. 

Nomenclature 

E elastic modulus 

σf stress at failure 

ρ material density 

CTE coefficient of thermal expansion 

H Poisson’s ratio 

r grid radius 

d grid gap 

t grid thickness 

p electrostatic pressure 

δ axial deflection 

I. Introduction 

Electrostatic ion propulsion is being considered for a number of demanding space missions by NASA and other 

organizations. The high specific impulse provided by these thrusters is viewed as an enabling technology, without 

which some missions might be impossible. Some of the first efforts to fabricate carbon fiber ion optics were made 

by the Jet Propulsion Laboratory. These were relatively small ion optics assemblies, typically 10 to 15 cm in 

diameter (refs. 1 and 2). Circular and slit-type apertures were investigated, and implied major consequences on 

carbon fiber integrity (refs. 3 and 4). Success of the NASA Solar Electric Propulsion Technology and Readiness 

(NSTAR) program focused interest on 30 cm ion optics, and JPL fabricated carbon fiber ion optics intended as an 

alternative to the molybdenum grids that had been baselined (ref. 5). The more recent Carbon Based Ion Optics 

(CBIO) program directed significant resources toward large-diameter carbon optics, including both carbon fiber 

composites and pyrolytic graphite (refs. 6 and 7). NASA Glenn Research Center (GRC) successfully pursued 

pyrolytic graphite (PG) ion optics development under the Energetics Program, demonstrating good laboratory 

performance and structural integrity (refs. 8 and 9). Lessons learned in the CBIO program were used as a starting 

point for the higher power Nuclear Electric Xenon Ion System (NEXIS) thruster (ref. 10). In a similar way, NASA 

GRC designed and tested large PG ion optics designs under the High Power Electric Propulsion (HiPEP) program 

(ref. 11). There has also been significant international success with ion propulsion. The Artemis spacecraft, 

sponsored by the European Space Agency (ESA), was salvaged from a misplaced orbit through use of its RITA and 

EITA ion propulsion system (ref. 12). The Hayabusa mission is moving forward with the MUSES-C success 

(ref. 13). 
NASA/TM—2005-213977 2

Sputter erosion of the grid electrodes will limit the functional life of an ion thruster. Dimensional changes to the 

screen and accelerator apertures will ultimately lead to the onset of electron backstreaming. The erosion rate of the 

grid electrodes depends on the voltage of thruster operation, and the material from which they are constructed. Since 

the 1950’s, molybdenum has been commonly used to fabricate ion thruster grids, due to its low thermal strain, 

resistance to high temperature, and relatively low sputter erosion rate. Since the 1990’s, carbon has received 

increased attention as a candidate material because it is significantly superior to molybdenum in all three of these 

areas. Carbon, however, is a non-metallic material, and introduces many manufacturing challenges. The thermal 

strain, elastic modulus, and ultimate strength of carbon can vary drastically, depending on its form. These properties 

are often inconsistent with commonly used metals within a thruster assembly. Special design considerations must be 

employed in order to accommodate carbon based materials. This paper will discuss some of the challenges that have 

surfaced in attempts to incorporate carbon materials into ion thrusters. The successes and failures of various projects 

will be discussed, spanning from the early 1990’s through to the present. 

II. Forms of Carbon 

A. Pyrolytic Graphite 

Structural carbon can take on different forms. Pyrolytic graphite is formed by chemically depositing carbon onto 

a precisely shaped mandrel. Hydrocarbon gas is admitted into a low pressure furnace, in excess of 1800 °C. The gas 

chemically cracks upon contact with the hot mandrel, leaving a stratified lattice of carbon that builds up over time. 

After cooling, the PG layer is dislodged from the mandrel and may undergo extensive post-machining. The material 

properties of PG are very anisotropic. Thermal conductivity and electrical conductivity exceed that for most metals 

in directions parallel to the deposition plane. Conversely, these qualities are quite poor in the direction normal to the 

deposition plane. PG is vulnerable to split or delaminate between planes, and care should be taken in designing parts 

to avoid this structural weakness. Its strength and elastic modulus are much higher parallel to the deposition plane, 

and produce excellent membrane-like shapes. Material density is typically about 2200 kg/m3

, which is near the 

theoretical maximum for graphite. 

B. Carbon Composites 

Carbon composites are fabricated with a preform of carbon fiber, consisting of filament tape or woven fabric. 

Multiple layers are often used in order to tailor strength properties within the part. Since internal forces within the 

part are carried primarily by the fibers, fiber density is kept to a maximum. Bonding between fibers is necessary in 

order to resist compressive and shear loads. Bonding is achieved by depositing additional carbon in between 

individual fibers, filling voids, and densifying the fiber matrix. Density is increased either by chemical vapor 

infiltration, or resin pyrolysis. The properties of carbon composites can be manipulated through careful selection of 

carbon fiber, and preform configuration. Care must be taken to avoid delamination between layers, and trapped 

regions of low density. Fibers are commercially available with different elastic modulus, tensile strength, and cost. 

Fiber properties can be modified by performing a post-densified heat treatment of the entire part. The combinations 

of fabrication techniques are endless. The maximum stiffness and strength of carbon composites are aligned with the 

fiber direction, and are significantly less in the cross-grain direction. Raw carbon fiber is available with a strength 

much higher than steel, however densified parts are seldom this strong because various ply directions are needed to 

distribute strength in all three axes. Thermal strain and conductivity are also aligned with fiber grain, with CTE

values that vary from slightly negative, to moderately positive. Material density can vary widely depending on the 

fabrication process, but 1800 kg/m3

 is typical. 

C. Polycrystalline Graphite 

A relatively isotropic form of carbon can be produced by processing carbon powder in an isostatic press. 

Material properties vary depending on the powder grain size, and final density of the compacted block. Grain size is 

typically on the order of 10 microns. Graphite components are mechanically machined from blocks of material, with 

directional orientation being arbitrary. Material density varies from 1750 to 1920 kg/m3

, depending on the specified 

grade. The thermal strain rate varies slightly with density and grain size. Generally, CTE is uniform in all directions, 

and is comparable to that of molybdenum. Ultimate strength varies significantly with material grade. Flexural 

strength of 138 Mpa (20 ksi) are available in small sizes, but elastic modulus is typically a fraction of that compared 

to PG. 
NASA/TM—2005-213977 3

III. Material Behavior 

A. Sputter Erosion 

The sputter erosion rate for graphite is roughly an order of magnitude less than that of molybdenum. The exact 

rate is somewhat variable, depending on particle energy, incident angle of particle, surface finish, etc. Chargeexchange (pit and groove) erosion occurs on the downstream surface of an ion thruster accelerator grid. It is easily 

seen as a hexagonal honeycomb pattern among the apertures, and can ultimately lead to structural failure of the 

accelerator grid. Aperture barrel erosion is less visible, and occurs on the cylindrical wall of the accel apertures. As 

the diameter of the apertures increase, they will ultimately enlarge to the point where electron back-streaming cannot 

be prevented. For most space missions, aperture enlargement is viewed as the most likely failure scenario. 

Recent data has indicated that carbon fiber composites and PG have roughly the same sputter erosion rate. The 

erosion rate for polycrystalline graphite is roughly 40 percent higher than carbon composites or PG (ref. 14). If 

material density is considered, PG is about 25 percent more dense than the others. There is concern that anisotropic 

forms of carbon will wear irregularly, leading to jagged surfaces, enhanced electric field, and more frequent intergrid arcs. While plausible, this issue has not been conclusively proven. There are numerous wear tests presently 

underway that could support or discount this concern. 

B. Thermal Strain 

Thermal strain affects the dimensional stability of ion optics. The CTE of most materials is a very small number, 

and results in linear expansion that is of little consequence. It is the differential expansion of various components 

that must considered. Choosing materials with the same CTE is not good enough, because temperature variations 

exist within an ion thruster that will cause differential expansion. The radial temperature distribution within most ion 

thruster grids is a clear example of this effect. If radial expansion within a flat disk exceeds hoop expansion, internal 

stress will develop. Radial compressive stress leads to buckling, and buckling dramatically amplifies geometric 

distortion. The flat disc with barely any height will become dish shaped. Instead of dimensions changing by only a 

tenth of a percent as predicted by linear theory, buckling amplifies this effect and height may increase by a factor of 

ten or more. 

One way to reduce the chance of buckling is to select a material with a lower CTE. The problem diminishes as 

CTE is reduced, and perfect stability is achieved when CTE is identically zero. There are no common metals with a 

CTE of zero, and invar-type materials are only effective over a narrow temperature span. Fortunately, there are 

carbon-based materials with very low CTE, however they often meander with temperature between slightly negative 

and slightly positive. In the case of flat thin ion thruster grids, a slightly negative CTE works out well with a radial 

temperature distributions. Assuming the perimeter of the grid is clamped tightly to a rigid stiffening structure, the 

thin part of the grid will draw into tension. As with the skin of a drumhead, the grid surface will remain flat as long 

as the perimeter can tolerate the compressive hoop stress. This design can be implemented with certain types of 

carbon fiber composite or pyrolytic graphite, which have negative CTE. While this may represent a thermal design 

solution, flat ion thruster grids remain vulnerable to structural issues of electrostatic deflection and launch vibration. 

Another solution to the problem of buckling is to design-in spherical curvature throughout the radius of the grid. 

Almost all molybdenum optics are domed in order to permit controlled and predictable buckling as the thruster goes 

through each thermal cycle. Since temperature gradients are time dependent, some grid gap variation is inevitable 

with metallic optics. Fabrication complexity is significantly increased with domed optics in that curvature must be 

carefully matched, and aperture locations compensated for the diverging beam. Chemical etching of metallic grids is 

a very cost effective approach to form screen and accelerator apertures. Tens of thousands of apertures can be 

formed simultaneously, during an etching process that may take only an hour. Chemical etching can be performed 

on either flat or domed geometry, with no burs or strain-hardening of the metal. Unfortunately, carbon grids cannot 

be chemically etched with existing technology, and apertures must be drilled hole by hole. This is a time consuming 

and costly procedure, requiring high-precision 5-axis CNC machines. 

Thermal strain can affect the structural integrity of an ion thruster. It is common practice in flight systems to 

maximize first mode resonant frequency by increasing structural stiffness. However, if a rigid optics frame is mated 

to a rigid ion thruster, there is risk of severe cyclical stress. Unless components are made from similar material, and 

operate at similar temperatures, it is advisable to design transition joints with some degree of strain isolation. What 

fits together at room temperature may see significant stress at operating temperature. It may take numerous thermal 

cycles for evidence to appear, but thermally induces structural fatigue can lead to system failure. Flexures need to be 

carefully designed to accommodate thermal strain of normal operation, yet survive launch vehicle dynamic loads. 
NASA/TM—2005-213977 4

C. Material Robustness 

There are a number of ways in which the ion optics assembly might be damaged. Areas-of-risk include 

manufacturing and assembly, launch vehicle vibration, and electric arcs. Under some circumstances, failure occurs 

at the onset of plastic deformation, where permanent distortion goes beyond geometric specifications. Under other 

circumstances, failure is not reached until the point of material rupture. The two words “Strength” and “Toughness” 

are sometimes confused as being interchangeable, but in reality they define different qualities. The following three 

scenarios distinguish different types of stress-induced effects. 

1) A high strength material will tolerate high stress before failure. Silica glass has fairly high strength, but it 

also has a high elastic modulus. Stress levels increase at a very steep rate, and the material will shatter when 

impacted by a similar object. There is no ability to absorb impact energy via plastic deformation, and failure 

occurs as brittle fracture. 

2) Silicone rubber has a low failure stress, and a very low elastic modulus. When impacted, the material 

elastically deforms, converting kinetic energy into strain energy. Large deformation may take place because 

the low modulus delays the buildup of stress, and fracture will not take place until ultimate stress is exceeded. 

Fairly high specific energies can be absorbed before material failure occurs. Under normal conditions, the 

strain energy will re-bound back to kinetic energy, with complete recovery to the original state. This type of 

response is typical of materials with high toughness. 

3) Annealed copper has a low yield strength, and fairly high elastic modulus. Under impact, the high modulus 

quickly results in elevated stress levels within the material. Very little strain energy can be stored before 

plastic deformation begins. Because of its ductility, much cold-working may take place, where kinetic energy 

is irreversibly converted to heat. While the original geometry was lost very early, fracture is delayed by the 

cold-working process until strain-hardening reaches the ultimate stress limit. Ductile materials can absorb 

tremendous punishment before fracture occurs, and are often employed where human safety is a concerned. 

For ion optics, failure occurs when yield strength is exceeded, even though there is no fracture. 

The maximum amount of strain energy that is storable in a material can be predicted by evaluating the strainenergy equation at the point of failure. If the material is linearly elastic, 

Maximum Strain Energy

(per unit volume)

1

2

f

2

E (1) 

where σf represents stress at failure, and E is the elastic modulus (ref. 15). The more a material can store strainenergy, the more it can resist damage. When a material exceeds its Maximum Strain Energy, it either plastically 

deforms, or fractures. Launch loads are an inertial effect in which the entire spacecraft is subjected to mechanical 

vibration. Dynamic stress is induced proportional to material density. As might be expected, the lightest materials 

will see the lowest dynamic stress. Conversely, a structure fabricated with high density material will experience 

higher dynamic stress. The Maximum Strain Energy equation can be recast into units of mass, such that: 

Maximum Strain Energy

(per unit mass)

1

2

f

2

E (2) 

where ρ is material density. This number can be calculated for many different types of material in order to compare 

their relative resistance to dynamic loads. In table 1, strength is defined as either maximum yield strength or point of 

material fracture. 

TABLE 1.—MAXIMUM STRAIN ENERGY OF VARIOUS MATERIALS. 

Material Strength, Mpa Elastic Mod., Gpa Density, kg/m3 Max Energy, J/kg 

Stainless Steel 207, yield 193 8,030 13.8 

Molybdenum 655, yield 324 10,220 64.8 

Titanium 483, yield 103 4,850 233 

C-C Comp. 345, fracture 206 1,830 158 

Pyrolytic C 89.6, fracture 32.4 2,220 55.8 

Polycrystalline C 68.6, fracture 12.7 1,900 97.5 
NASA/TM—2005-213977 5

As seen in the table, high-strength materials are not necessarily the most damage resistant. Molybdenum has a very 

high yield strength, and is presently the most common metal used in ion optics. It has a relatively high density, 

resulting in higher stress under dynamic loads. The high elastic modulus results in good rigidity, but this also 

increases stress under shock loads. Pyrolytic graphite has an order of magnitude lower strength and stiffness, yet it 

has a comparable strain energy storage limit. The low density and elastic modulus of PG are an advantage, and 

permit significant storage of strain energy before stress induced failure occurs. Titanium shows very high strain 

energy storage, and has been considered as a possible replacement for molybdenum. Polycrystalline graphite has a 

very low elastic modulus, and can endure extreme deformation before failure occurs. Carbon-carbon composites 

show excellent potential for toughness, due to its high tensile strength. Anisotropic materials must be considered 

with caution because subtle vulnerabilities may exist. The strain energies tabulated above represent macroscopic 

generalizations, and may be overly optimistic at small length scales. Flakes or chips are sometimes dislodged from 

larger components, posing an arc hazard. Both pyrolytic graphite and carbon composites are prone to inter-laminar 

separation, requiring careful design consideration. 

D. Electrostatic Attraction 

There is a mutual attraction force between the screen and accelerator grids due to applied voltage across the gap. 

Even the open area of the screen grid is bridged over by fringe effects, and contributes to the pull of the electric 

field. A first-order estimate of the electrostatic pressure p on the grids can be approximated by (ref. 16) 

p

V2

2d2 (3) 

where ε is permittivity of free space (8.842×10–12 C2

N–1m–2). For a flat circular grid, this force deflects the surface in 

a way similar to hydrostatic pressure, and can be estimated as (ref. 17) 

 3

16 pr 4 (1 )(5 ) /(Et 3

) (4) 

Deflection δ must be calculated for both the screen and accelerator grid in order to determine constriction of the grid 

gap. Material macro-properties may be chosen to account for high open-area, because the equation assumes a 

uniform disk, of radius r, and thickness t. If the change in grid gap is significant, flat grids may not be a workable 

design. As the gap is drawn together, the attraction force increases. Ultimately, a run-away situation is reached 

where the grids come into contact and temporarily short. Either the grids must be strengthened, or the electric field 

must be reduced. A shallow spherical dome has tremendous stiffness relative to a flat disk. Manufacturing a dome, 

however, is factors more complex. 

IV. Design of Carbon-Metallic Interface 

The screen grid and accelerator grid of an ion thruster can be made of carbon in order to take advantage of the 

sputter resistant properties of this material. The thermal strain rate of carbon is often very low, resulting in good 

geometry stability of the aperture alignment. The majority of the spacecraft, however, is often metallic, with 

considerably higher thermal strain rate. Tubing, wire, magnetic components, and general structure are often made of 

aluminum, stainless steel, or titanium. Eventually, there must be an interface between the carbon ion optics and the 

metallic assemblies, and this interface needs to accommodate differing strain rates. 

A. Ground Based Applications 

A number of industrial ion sources simply support the carbon optics components with oversized clearance holes. 

The accelerator grid is loosely held, with as much as a millimeter of free radial movement within the grid plane. The 

screen is tightly coupled to the accelerator grid via ceramic insulators. Several insulators are positioned around the 

perimeter, to establish grid gap and aperture alignment. Since they are of the same material, differential expansion is 

not an issue. The pair are permitted to float relative to the ground structure, as the thruster goes through each thermal 

cycle. For industrial applications, this is a very simple and functional design. A loose optics assembly, however, 

might be expected to rattle quite violently under launch vehicle vibration, making this unacceptable for a flight 

design. 
NASA/TM—2005-213977 6

B. Early 30 cm Carbon-Carbon 

In the mid 1990s, JPL fabricated a C-C composite ion optics assembly under contract with NASA (ref. 5). The 

design was intended to be compatible with the NSTAR 30 cm thruster, which was baselined with molybdenum 

optics. Shallow domed screen and accelerator grids were mounted onto a C-C composite mounting ring via ceramic 

insulators. The optics assembly was completely non-metallic, except for fasteners (see fig. 1). It was self-contained, 

permitting grid gap and aperture alignment prior to installation on the thruster. The screen grid and accelerator grid 

were made from six plies of prepreg fiber ribbon. The layers were symmetrically stacked, with fiber direction 

staggered at 60° intervals. Blank panels were carbonized, graphitized, and Chemical Vapor Infiltration (CVD) 

densified over a domed mandrel. Apertures were laser machined into the blank panels to complete the screen and 

accelerator grids. A mounting ring was made from carbon fiber fabric using a design very similar to the titanium 

mounting ring in the NSTAR thruster. The screen grid attached directly to the mounting ring via machine screws. 

The accelerator grid was suspended away from the screen on ceramic insulators identical to those used by NSTAR. 

Precise grid gap was to be adjusted by placing shims between the accelerator grid and the insulator mounting face. 

Unfortunately, the spherical curvature of the two grids did not match, resulting in severe gap variation along the 

radius. The carbon optics assembly was rigidly mated to the metallic NSTAR thruster through 6 ceramic insulators. 

There was no specific design feature intended to accommodate the different strain rates between the carbon optics 

and the thruster. The severe non-uniform grid gap prevented the optics from being operated at full power, so the 

carbon-metallic interface was never tested under design conditions. 

C. Eight cm Pyrolytic Graphite 

An 8 cm pyrolytic graphite optics assembly was tested at NASA GRC in 2002 (ref. 8). The flat grids had 

aperture geometry similar to that of NSTAR, and were aligned on a C-C composite mounting ring (see fig. 2). The 

self contained optics assembly was mated to the metallic ion thruster via 4 ceramic insulators, similar to the JPL 

approach. The insulators were long relative to the grid diameter, which may have eased the strain mismatch. The 

thruster discharge chamber was oversized relative to the optics assembly, and represented a laboratory 

configuration. It ran well, with no consequences from the simple interface. 

D. Thirty cm Pyrolytic Graphite, GRC 

Success of the flat 8 cm PG ion optics was quickly followed by fabrication of domed 30 cm PG optics (ref. 9). 

The screen grid, accelerator grid, screen stiffener, and accelerator stiffener were designed to be similar to 

molybdenum components used on NSTAR. The grids were attached to the stiffeners using small machine screws, 

and then installed on a titanium mounting ring of NSTAR design. Radial strain accommodation for the screen grid 

was achieved by flexures integral with the original NSTAR mounting ring. Radial strain accommodations for the 

accelerator grid was absent from the original NSTAR design, and needed to be addressed. New mounting posts were 

fabricated for the accelerator grid which incorporated face-to-face leaf spring flexures. At room temperature, the 

facing leaf springs were fully compressed, and nearly bottomed out. As the temperature of the titanium mounting 

ring increased, it expanded radially relative to the accelerator grid and began to draw the leaf springs apart. The 

flexure design was axially rigid, but very compliant in the radial direction. The flexures were also rigid in the 

tangential direction, allowing for precise alignment between the screen and accelerator apertures. The new flexures 

were compact enough to be installed without any modification to the NSTAR insulators or mounting ring. Thruster 

performance with PG ion optics was similar to that of NSTAR, and easily covered the entire operating envelope. 

There was some grid gap variation due to non-spherical aberration in the screen grid. Some performance differences 

might have been due to cylindrically machined apertures instead of chemically etched apertures. The cusp normally 

present from the etch process would behave as an exceptionally thin screen grid, improving screen grid 

transparency. The mechanical drilling of the carbon screen left a more blunt aperture edge, reducing transparency. 

Upon completion of performance measurements, the 30 cm pyrolytic ion optics began structural dynamic testing. 

Modal testing revealed a first mode resonance of 504 Hz with the screen grid, and 664 Hz with the accelerator grid 

(see fig. 3). The ion optics were then installed on NSTAR EMT1c for vibration testing. Random vibration level was 

gradually stepped up, and eventually reached DS-1 acceptance level for 60 seconds. Grid-to-grid contact was 

initially detected at 2 Grms, and was very intense at 9.3 Grms. The ion optics assembly showed no sign of damage 

after the test, although a small amount of fine carbon dust was collected under the grids. 

E. CBIO Program 

The CBIO program consisted of parallel efforts with the intention of developing large scale carbon ion optics. A 

carbon-carbon composite approach was pursued by the Jet Propulsion Laboratory, and a pyrolytic graphite approach 

was pursued by Boeing EDD (refs. 6 and 7). Both approaches were sized to a modified NSTAR geometry, and 
NASA/TM—2005-213977 7

intended to be mountable as a substitute for the original molybdenum optics. 

The pyrolytic graphite approach used a spherically domed accelerator grid and modified dome screen grid. The 

original NSTAR titanium mounting ring was essentially unchanged from NSTAR, and began as the basis for the 

new design. The mounting ring was naturally compatible with the rest of the NSTAR thruster, which was essentially 

spin formed titanium. Ceramic insulators were attached to the mounting ring, which electrically isolated the 

accelerator grid from the screen grid, and the screen grid from the anode. Both grids were mounted on titanium 

stiffeners and held in place with titanium clamping rings. Each pyrolytic graphite grid was loosely sandwiched 

between the stiffener and the clamp, but allowed to slide radially as materials expanded at different rates. Graphite 

foil was inserted to help dampen and lubricate radial movement. Short radial slots in the grid flange were to follow 

pins in the stiffener to prevent tangential motion and allow uniform radial expansion. Unfortunately, there were a 

number of problems with carbon grid fabrication, and particularly so with the screen grid. The first set of carbon 

grids experienced stress relief distortion upon machining, resulting in mismatch between the spherical curvature. 

The flange surrounding both carbon grids was not flat and became canted by a few degrees. The screen grid 

developed cracks along rows of aperture holes that spanned inches in length. Many of these flaws were corrected by 

implementing an extra stress relief process to the furnace cycle, and through staged machining. Once successfully 

assembled, the pyrolytic graphite ion optics was installed on NSTAR EMT1c and vibration tested to DS-1 levels. No 

grid-to-grid contact was detected with the CBIO PG design, no physical damage was detected, and aperture 

alignment was unchanged. Performance measurements of the PG ion optics were conducted at Boeing EDD. Very 

high accelerator currents were measured, suggesting that there may have been poor aperture alignment, or distorted 

gap. Visual observation of the thruster under operation revealed a red-orange glow on the accelerator grid, leaving 

little doubt that direct accelerator impingement was occurring. It is conceivable that jammed movement in the 

stiffener/grid interface could have adversely affected grid alignment or gap. 

The carbon-carbon composite ion optics developed under CBIO followed a much different development path. A 

fresh design was started, which was made almost completely of carbon fiber. Both screen and accelerator grids were 

intended to have uniform spherical curvature, with heavy stiffening rings around each grid perimeter. Both grids 

consisted of layered uni-directional fiber tape, with each layer clocked 60° relative to the next. The layups were 

carbonized and densified over a spherical mandrel. Apertures were laser machined into both grids, and a final CVD 

deposition was coated over all surfaces. The curvature of the mandrels was adjusted to account for heat treatment 

and laser machining effects. Adjustments were empirical, and solved the grid gap variation problem experience in 

the early 30 cm work. A double-walled cylinder mounting ring was made from carbon fiber, and provided extreme 

rigidity to the optics assembly (see fig. 4). The carbon fiber screen grid was bonded directly to the mounting ring. 

Ceramic insulators were recessed within the walls of the mounting ring, on which the accelerator grid was installed. 

A flat carbon fiber stiffener was bonded around the perimeter of the accelerator grid, and shimmed to the correct 

grid gap via the insulators. Since the entire optics structure was made from carbon fiber, thermal strain was nearly 

zero for all components. While the thruster mounting ring was electrically isolated from the anode with rigid 

ceramic insulators, strain isolation at this transition was not addressed. The extreme stiffness of the mounting ring 

likely dominated the interface. Vibration tests were conducted at various levels of assembly, beginning with 

individual grids. The ion optics were then tested as an assembly, and later installed on EMT1c. Ultimately, the 

carbon-carbon composite ion optics were tested on the thruster, meeting the 9.1 Grms acceptance level. There was 

no observed contact between grids, and no damage was detected following the test. Ion thruster performance was 

mapped out over operating conditions defined at the start of the program. Thruster performance was good, and only 

fell short at the maximum beam current operating point. Initial rate of recycle was much lower than expected. The 

thruster was wear tested for 1028 hours without serious incident; however this fell short of the intended 2000 hour 

test. The ion optics were in fairly good condition following this test. There were some minor delaminations and 

broken webs, but some of this was traceable to initial fabrication. 

F. NEXIS 

The ion optics used in the 20 kW NEXIS DM is essentially a carbon fiber assembly scaled up from CBIO 

(ref. 10). A double walled cylinder formed the thruster mounting ring, and the screen grid was bonded directly to the 

ring (see fig. 5). Insulators were recessed into the hollow volume of the ring, and isolated the accelerator from the 

screen. Shims were placed between the insulator and accelerator grid in order to adjust grid gap. Except for 

fasteners, the entire NEXIS optics assembly was made of carbon fiber, which reduced thermal strain to near zero. 

Since the thruster chamber was a metallic assembly, strain isolation was necessary across this interface. An early 

design consisted of an array of 18 “C” flexures distributed around the interface from the carbon mounting ring to the 

thruster chamber. Elevated stress levels were predicted, so the design was replaced with an array of 9 bipod flexures. 

Each flexure pair spanned from the back of the carbon mounting ring, to a metallic bulkhead surrounding the 
NASA/TM—2005-213977 8

thruster chamber. The bipods were flexible in the radial direction, but self-bracing for stiffness in the axial and 

tangential directions. Electrical isolation between the bipod and the thruster completed the design. The NEXIS 

thruster performed well, with few structural issues. The thruster began what is planned to be a 2000 hour wear test to 

determine the effect of sputter erosion on optics geometry, and thruster performance. 

G. HiPEP 

The 20 kW HiPEP Gen1b thruster had flat, rectangular, PG grids (ref. 11). They were the largest single-piece PG 

components that could be commercially manufactured. The screen grid and accelerator grid were independently 

mounted off a downstream frame, which was at ground potential. PG stiffeners were mounted directly to the 

perimeter of both grids, and each set of stiffeners was supported by an array of 22 flexures. The flexures 

accommodated thermal expansion of the discharge chamber and optics mounting frame. All flexures were rigid in 

the axial direction. However, they were all flexible radially outward from a single point at the centroid of the 

rectangle. The fixed end of the flexures were anchored to individual ceramic insulators. The insulators were 

mounted to a forward frame structure, in-place of what is commonly referred to as the front mask. This laboratory 

design allowed for relatively easy grid gap adjustment and aperture alignment. Although Gen1b operated very well, 

the downstream oriented optics frame was an issue of concern. It was felt that beamlet divergence in excess of 

45° might become a life limiting impingement issue, so an extensive redesign was pursued. 

A second generation HiPEP thruster used the same PG grid electrodes, but with a completely different mounting 

arrangement. The screen and accelerator grids were independently mounted off an upstream frame in order to 

provide a much greater plume clearance angle. Grid extensions were necessary in order to straddle the discharge 

chamber. One hundred stainless steel extension fingers surrounded the perimeter of each grid (see fig. 6). The 

fingers were designed to pivot at both ends, allowing differential expansion between the PG grids and metallic grid 

stiffeners. Four centerline fingers were rigid, and keyed to prevent drifting of screen to accelerator alignment. The 

stainless steel stiffeners were supported by a series of tandem flexures, which allowed movement orthogonal to that 

of the extension fingers. The extension finger and flexure arrangement was very complex, but was intended to 

provide similar behavior if a large enough sheet of PG were available. The design worked well as a laboratory 

thruster, but larger sheets of PG or an integral carbon frame extender would be necessary for a vibration tolerant 

flight design. The ion optics assembly was mated to the thruster chamber through a series of tie rods, connecting the 

two frames. Thruster performance matched predictions very well (ref. 18). Grid to grid arcs were relatively frequent 

at first, but subsided to about one per hour after about 100 hours of operation. The thruster has recently completed 

2100 hours of wear testing, but condition of the electrodes is not known as of this writing. Structural dynamic tests 

were performed at Aerojet Corporation using identical pyrolytic graphite grids (ref. 19). Aluminum grid extenders 

were installed around the perimeter of each grid to simulate how a larger flight representative thruster might behave. 

The optics set was vibration tested at DS-1 levels for 60 seconds. Heavy grid-to-grid contact occurred, but no 

significant structural damage was observed. A 2.0 g extended sine sweep was also preformed on the grid set, with 

similar results. 

V. Conclusion 

Carbon Ion Optics are expected to provide much longer thruster life due to their resistance to sputter erosion. Ion 

thruster optics can be fabricated from carbon fiber composites, pyrolytic graphite, or polycrystalline graphite. All 

three of these carbon materials can store significant amounts of strain energy, and should be resistant to impact 

loads. The mechanical behavior of carbon is much different than that of most metals, and poses unique design 

challenges. In order to minimize mission risk, the behavior of carbon must be well understood, and components 

designed within material limitations. Thermal expansion of the thruster structure must be compatible with thermal 

expansion of the carbon ion optics. Specially designed interfaces may be needed so that grid gap and aperture 

alignment are not adversely affected by dissimilar materials within the thruster. The assembled thruster must be 

robust, and tolerant of launch vibration. Carbon ion optics hardware has improved significantly over the past 

10 years. 
NASA/TM—2005-213977 9

References 

1. Garner, C.E. and Brophy, J.R., “Fabrication and Testing of Carbon-Carbon Grids for Ion Optics,” AIAA 

Paper 92–3149, 28th Joint Propulsion Conference, Nashville, TN, July 6–8, 1992. 

2. Mueller, J., et al., “Performance Characterization of 15-cm Carbon-Carbon Composite grids,” AIAA Paper 

94–3118, 30th Joint Propulsion Conference, Indianapolis, IN, June 27–29, 1994. 

3. Brophy, J.R., et al., “Carbon-Carbon Ion Engine grids with Non-Circular Hole Apertures,” AIAA Paper 

95–2662, 31st Joint Propulsion Conference, San Diego, CA, July 10–12, 1995. 

4. Meserole, J.S., and Rorabach, M.E., “Fabrication and Testing of 15-cm Carbon-Carbon grids with Slit 

Apertures,” AIAA Paper 95–2661, 31st Joint Propulsion Conference, San Diego, CA, July 10–12, 1995. 

5. Mueller, Juergan, et al., “Design, Fabrication and Testing of 30-cm Dia. Dished Carbon-Carbon Ion Engine 

Grids,” AIAA Paper 96–3204, July 1996. 

6. De Pano, M.F., et al., “Fabrication and Vibration Results of 30-cm Pyrolytic Graphite Ion Optics,” AIAA 

Paper 2004–3615, 40th Joint Propulsion Conference, Fort Lauderdale, FL, July 11–14, 2004. 

7. Snyder, J.S., et al., “Performance Characterization and Vibration Testing of 30-cm Carbon-Carbon Ion 

Optics,” AIAA Paper 2004–3959, 40th Joint Propulsion Conference, Fort Lauderdale, FL, July 11–14, 2004. 

8. Haag, T.W., and Soulas, G., “Performance of 8-cm Pyrolytic Graphite Ion Thruster Optics,” AIAA Paper 

2002–4335, 38th Joint Propulsion Conference, Indianapolis, IN, 2002. 

9. Haag, T.W., and Soulas, G., “Performance and Vibration of 30-cm Pyrolytic Ion Thruster Optics,” AIAA 

Paper 2003–4557, 39th Joint Propulsion Conference, Huntsville, AL, 2003. 

10. Monheiser, J., et al., “Conceptual Design of the Nuclear Electric Xenon Ion System (NEXIS) Engine,” AIAA 

Paper 2004–3614, 40th Joint Propulsion Conference, Fort Lauderdale, FL, July 11–14, 2004. 

11. Williams, G., et al., “Performance Characterization of HiPEP Ion Thruster Optics,” AIAA Paper 2004–3627, 

40th Joint Propulsion Conference, Fort Lauderdale, FL, July 11–14, 2004. 

12. Killinger, R., et al., “Orbit Raising with Ion Propulsion on ESA’s ARTEMIS Satellite,” AIAA Paper 2002–

3672, 38th Joint Propulsion Conference, Indianapolis, IN, 2002. 

13. Hayakawa, Y., et al., “Endurance Test of C/C grids for 14-cm Xenon Ion Thrusters,” AIAA Paper 2002–

3958, 38th Joint Propulsion Conference, Indianapolis, IN, 2002. 

14. Williams, John, “Differential Sputtering Behavior of Pyrolytic Carbon and Carbon-Carbon Composites,” 

September 7, 2004. 

15. Rivello, R.M., Theory and Analysis of Flight Structures, McGraw-Hill, 1969, page 113. 

16. Lorrain, P., and Corson, D., Electromagnetic Fields and Waves, W.H. Freeman and Company, 1988, page 

111. 

17. Lindeburg, M., Mechanical Engineering Reference Manual, Professional Publications Inc., 1994, chapter 14. 

18. Williams, G., et al., “Performance Characterization of a 25 kW, High-Specific Impulse Ion Thruster,” AIAA 

Paper 2005–4248, 41st Joint Propulsion Conference, Tucson, AZ, 2005. 

19. Polaha, J., et al., “Random and Sine-Spectrum Vibration Testing of Pyrolytic graphite Ion Optics,” AIAA 

Paper 2005–4395, 41st Joint Propulsion Conference, Tucson, AZ, 2005. 
NASA/TM—2005-213977 10

Figure 1.—Early 30 cm carbon-carbon ion optics assembly. 

Figure 2.—Eight centimeter pyrolytic graphite ion optics assembly. 

Figure 3.—Thirty centimeter pyrolytic graphite ion optics assembly. 
NASA/TM—2005-213977 11

Figure 4.—CBIO carbon-carbon ion optics assembly. 

Figure 5.—NEXIS carbon-carbon screen grid and mounting ring. 

Figure 6.—HiPEP Gen2 laboratory PG ion optics assembly. 
This publication is available from the NASA Center for AeroSpace Information, 301–621–0390.

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17

Mechanical Design of Carbon Ion Optics

Thomas Haag

Space propulsion; Ion thruster; Electric propulsion

Unclassified -Unlimited

Subject Categories: 20 and 24

Prepared for the 41st Joint Propulsion Conference and Exhibit cosponsored by AIAA, ASME, SAE, and ASEE,

Tucson, Arizona, July 10–13, 2005. Responsible person, Thomas Haag, organization code RPP, 216–977–7423.

Carbon Ion Optics are expected to provide much longer thruster life due to their resistance to sputter erosion. There are

a number of different forms of carbon that have been used for fabricating ion thruster optics. The mechanical behavior of

carbon is much different than that of most metals, and poses unique design challenges. In order to minimize mission risk,

the behavior of carbon must be well understood, and components designed within material limitations. Thermal expansion of the thruster structure must be compatible with thermal expansion of the carbon ion optics. Specially designed

interfaces may be needed so that grid gap and aperture alignment are not adversely affected by dissimilar material

properties within the thruster. The assembled thruster must be robust and tolerant of launch vibration. The following

paper lists some of the characteristics of various carbon materials. Several past ion optics designs are discussed, identifying strengths and weaknesses. Electrostatics and material science are not emphasized so much as the mechanical behavior

and integration of grid electrodes into an ion thruster.