Advanced CMC Leading Edge Materials
and Processing for Hypersonic Vehicles
AFOSR SBIR Phase I Contract FA9550-14-C-0051
Matthew Wright and Jerry Brockmeyer Ultramet
Pacoima, California
Materials for Extreme Environments Program Review 2015
Arlington, Virginia May 18–22, 2015
• Background and Objectives
• Material Selection
• Thermal and Structural Analysis
• Test Article Fabrication
• Testing and Test Results
• Prior Related Work
• Summary
Agenda
2
Background and Objective
Note: Because of the open nature of this meeting, export-
controlled details are neither included nor quantified.
• Hypersonic vehicle leading edge structures require lighter
weight, higher performance, and improved manufacturability
relative to current materials, including CMCs containing
silicon and boron.
• Melt infiltrated ultrahigh temperature capability CMCs have
potential to meet these requirements.
• Objective: to demonstrate feasibility of ultrahigh temperature,
melt infiltrated CMCs for hypersonic vehicle leading edges
• Design and analysis (MR&D), based on prime contractor
mission needs, to define requirements
• Materials selection, fabrication, and testing under simulated
operational conditions to demonstrate capability and identify
needs for further development 3
• Initial application is relatively short duration, single cycle
• Materials must be durable in high temperature, oxidizing
operational environment
– Carbon fiber beneficial for density, cost, and mechanical
properties
– Even for short duration, matrix and / or interface must provide
oxidation protection of reinforcement fibers
• Aerodynamic heating and thermal gradient-induced
stresses lead to need for materials with thermostructural
durability to ~2500°C (4500°F)
– Hafnium and zirconium compounds and selected binaries
feasible
– HfC (r = 12.2 g/cm3) provides highest temperature capability
but in situ formed HfO2 only marginally higher temperature
capability than ZrO2 formed on ZrC (r = 6.6 g/cm3)
Material Selection
4
• Materials must be shape stable and damage tolerant
and resistant to particle impact and rain erosion.
– CMCs preferable to monolithic UHTCs
• Cf / HfC initially selected for design and analysis
• User interest for selected application led to switch to
Cf / ZrC for test article fabrication and testing
Material Selection—cont’d
5
Thermal and Structural Analyses
• Thermal analysis performed using mission scenario
• Radiation equilibrium temperatures and heat loads calculated
for various leading edge radii using “worst case” assumptions
at selected key times during mission
• Optimum performance achievable with sharp leading edge radii
drives need for ultrahigh temperature materials
0
1000
2000
3000
4000
5000
6000
0 0.2 0.4 0.6 0.8 1 1.2
Rad
iati
on
Eq
uili
bri
um
Te
mp
era
ture
s (°
C)
Leading Edge Radius (in)
Summary of Radiation Equilibrium Temperatures ( C)
150 sec
250 secs
280 secs
0
500
1000
1500
2000
2500
3000
3500
4000
4500
5000
0 0.2 0.4 0.6 0.8 1 1.2
He
at L
oad
s (B
TU/f
t2-s
ec)
Leading Edge Radius (in)
Summary of Heat Loads (BTU/ft2-sec)
150 sec
250 secs
280 secs
Time 1
Time 2
Time 3
Time 1
Time 2
Time 3
Sharp LE Blunt LE Sharp LE Blunt LE
Low
High
6
Thermal and Structural Analyses—cont’d
• Transient temperatures calculated
assuming sinusoidal heat flux
distribution and radiation on all
faces
• Example shown is blunt radius
leading edge with balanced (1:1)
0/90 layup fiber architecture
Time 7
• Predicted thermal
gradient-induced loads
calculated for Cf/HfC
and compared with
computed CMC
properties to identify
potential issues (red)
and determine preferred
fiber architectures.
• Example is balanced
(1:1), [0/90] blunt radius.
• Other radii, imbalanced
(3:1) and quasi-isotropic
architectures also
examined
• Lower peak
temperatures with 3:1
fiber architecture
8
Thermal and Structural Analyses—cont’d
• Initial plan was to test at Air Force Laser Hardened
Materials Evaluation Laboratory (LHMEL) under
conditions (heat flux levels) defined by analysis for
specified durations.
• LHMEL scheduling conflicts led to test in Spytek
Aerospace Inc. combustor rig at specified
temperatures and durations.
• Test articles fabricated using available (1:1 [0/90])
carbon fiber reinforcement in panel configuration
consistent with Spytek capabilities.
• Test articles were melt infiltrated Cf/ZrC processed
using methods previously developed at Ultramet.
9
Test Article Fabrication
Step 1: Obtain carbon fiber preform (potentially net or near-net shape)
Thick section 2D and 3D C/C Braided near-net shape
Test Article Fabrication—cont’d
Step 2: Apply interface
coating(s) by CVD or UVCVD
ZrC HfC
Step 3: Infiltrate preform with sacrificial
carbon and infiltrate with molten metal
M (liquid) + C (solid) → MC (solid) 10
Test Article Fabrication—cont’d
Left, as-fabricated melt infiltrated
Cf/ZrC test panels for Spytek
combustor rig testing: each
~5×5×1.3 cm (2×2×0.5") and 130 g
11
Below, button and wedge (leading
edge configuration) articles
fabricated for future arcjet testing
Testing and Test Results
• Exposures conducted in Spytek Aerospace burner rig
• JP10-air augmented with acetylene-O2 for maximum
temperature of 2982°C (5400°F)
• Temperature monitored with two optical pyrometers
• Rapid insertion / removal port for test article insertion at
temperature and removal after defined exposure duration
• Test 1: duration 1, 2204°C (4000°F); Test 2: duration 2, 2204°C
(4000°F); Test 3: duration 1, 2482°C (4500°F)
• Test articles cooled rapidly in air after test 12
• Test articles inserted into rig
after preheat to target
temperature, held for
specified duration, then
removed and cooled in air
• Upon removal, all test
articles appeared intact
• Material loss and possible
delaminations observed
during cooling
4000 F, Duration 1
13
Testing and Test Results—cont’d
Testing and Test Results—cont’d
Impinged test article faces after test:
Left, 2204°C (4000°F), duration 1,
average loss rate 0.022 g/cm2∙s;
Center, 2204°C (4000°F), duration 2 (~8× duration 1),
average loss rate 0.0064 g/cm2 ∙ s
Right, 2482°C (4500°F), duration 1 (planned, actual ~1.3X),
average loss rate 0.0093 g/cm2 ∙ s
• Potential near-surface delamination observed
(consistent with analysis)
• White oxide evident on edges, less prevalent on flat faces
14
• Cf/ZrC flexural strength >50 ksi
• 2D and 3D fiber reinforcement
possible
• Various UVCVD and carbon
fiber interface coatings
available to meet application
requirements
for use temperature,
oxidation resistance,
and lifetime
• Hot-gas tested at
various facilities at
ultrahigh temperatures
15
Prior Related Work
Melt infiltrated Cf /ZrC and Cf /Zr-Si-C composite panels
with UVCVD ZrN interface coatings were tested by laser
heating with Mach 0.7 airflow. A heat flux up to 600 W/cm2
was generated that
induced a surface
temperature as high
as 5200 °F.
Nominally 1.2" diameter laser spot area
16
Cf/Zr-Si-C Cf/ZrC
Prior Related Work—cont’d
A B
SEM images of
ZrN coating on
YSH fabric
SEM images of
ZrN interface
coating on
Hi-Nicalon
fiber filaments 17
A B
UVCVD deposited ZrN interfaces
Prior Related Work—cont’d
Gases in
(H2) (Cl2) (O3)
Substrate
ZrCl4 (g)
Coating
UV Lights
UV Radiation
O3 + hn O2 + O**
MCl4 + hn MCl4*
MCl3* + Cl*
18
• Similar to
conventional
(thermal) CVD but
lower deposition
temperatures
• Not limited to
line of sight
• Single oxide
(ZrO2) deposition
shown for
reference
Prior Related Work—cont’d
UVCVD
process
schematic
UVCVD
Carbon interface coatings also effective depending on use
temperature, duration, and gaseous environment requirements.
19 Before test After test During test
Flow
Prior Related Work—cont’d
Test
no.
Specimen
ID
Max temp,
F ( C)
Heat flux,
Btu /ft2·sec Test time
No. of
tests
Weight
difference, g
001 34 2D ZrC 4000 (2204) 415 12 min @ temp 3 1.2290
002 38 2D ZrC 4000 (2204) 439 4 min @ temp 1 —
003 40 3D ZrC 4030 (2221) 443 4 min @ temp 1 0.2110
004 41 3D Zr-Hf-C 4042 (2227) 500 4 min @ temp 1 −0.7950
005 39 3D Zr-Si-C 3816 (2102) 456 8 min @ temp 1 −0.4230
006 37 2D ZrC 3891 (2144) 452 12 min @ temp 1 −1.8300
Arcjet Test Results (NASA LaRC HYMETS Facility)
Thick-section (6×2") Cf /ZrC composite
leading edges were fabricated in single
infiltration cycle and tested at LHMEL under
high heat flux/short-exposure and low heat
flux/long-exposure conditions. Recession
measurements confirmed that outer mold
line remained unchanged.
20
Prior Related Work—cont’d
Before test During test After test
T300 carbon fiber/melt infiltrated ZrC matrix composite
thruster with UVCVD ZrO2 fiber interface coating subjected to
multiple 30-sec O2/H2 hot-fire tests at NASA GRC to >4300 F
with minimal mass or dimensional change. First ply was
consumed to form protective ZrO2 layer.
21
Prior Related Work—cont’d
13" diameter Cf/ZrC CMC-lined C/C
nozzle hot-fire tested for 120 sec
using JP10 to >3000 F peak
temperature at ATK. No damage
was evident (SBIR Phase II for
NASA GRC)
22
Prior Related Work—cont’d
Summary
• Design and analysis for selected mission profile
indicated that ultrahigh temperature CMCs have
potential for hypersonic vehicle leading edges.
• Potential limitations principally related to low matrix
dominated properties for 2D fiber reinforcement
architectures
• Representative test conditions defined
• 2D panel specimens tested at targeted temperatures
in combustion rig environment
• Weight loss was greater than expected but no
catastrophic failure
• Stable, adherent oxide layer did not appear to form.
23
Summary—cont’d
24
• Supplemental test article fabrication shows ability to
fabricate leading edge shapes.
• Technology appears feasible with further work.
Future Work Recommendations
• Reconcile differences between Spytek combustor rig tests
and prior tests in LHMEL, arcjet, and JP10 exhaust nozzle
environments that showed adherent oxide formation and
little or no weight loss.
• Validate benefits of enhanced fiber architectures
(imbalanced vs. balanced, quasi-isotropic vs. orthogonal,
3D vs. 2D).
• Consider alternative matrices and / or surface treatments
for improved oxidation resistance and / or delamination
resistance.
• Enhance database to better support design and analysis.
• Update and scale up demonstrator designs.
• Fabricate and test “optimized” demonstrators in
simulated operational environment. 25
Acknowledgments
• Air Force Office of Scientific Research for funding this effort SBIR Phase I contract FA9550-14-C-0051
• Dr. Ali Sayir of Air Force Office of Scientific Research
for program and technical guidance and support
• Dr. Larry Matson of Wright-Patterson AFB for technical
support and insight
• Brian Sullivan, Mike Dion, Leslie Weller, and Derek Caputo
of MR&D for design and analysis
• Chris Spytek of Spytek Aerospace for combustor rig testing
26
Thanks to all for your time and consideration
