Mission Radar Radome Installation - Structural Substantiation

INTRODUCTION

Static-strength substantiation of the DHC-8-100 mission-radar radome installation, with emphasis on the custom fiberglass flange that transfers external aerodynamic demand into the aircraft lower fuselage.

The assessment combines conservative aerodynamic idealization, composite joint allowables, 3D rigid-body fastener load distribution, Classical Laminate Plate Theory (CLPT), and local interface shear checks.

AIRCRAFTDHC-8-100Mid-fuselage external radome

PRIMARY INTERFACE16-ply fiberglass flangeMeasured cured thickness: 0.192 in

LOAD PATHAerodynamics → Attachments → FlangeEight-bolt bounding reaction model

SUBSTANTIATIONJoints + CLPT + ShearPositive margins / inspection checks

Aircraft station diagram with radome installation location and dimensions — Aircraft-station definition and radome installation envelope.

DESIGN ASSESSMENT

RADOMESupplier-procured aerodynamic fairing

Protects mission radar/communications equipment while maintaining the required external contour and low signal attenuation.

CUSTOM FLANGEPrimary structural interface

In-house fiberglass laminate bridges the supplier radome to the aircraft lower skin and spreads attachment load around the perimeter.

FITTING SUB-ASSEMBLYMechanical load transfer

Transfers local attachment reactions between the radome installation and surrounding fuselage structure.

RADOME FRAMESPort + starboard support

Receive the remaining installation loads through the fitting architecture.

Radome installation schematic and primary structural interfaces.

LOAD-PATH RATIONALE

Separate global aerodynamic demand from local composite-interface checks.
External lift and drag enter through the radome shell, resolve into perimeter attachment reactions, pass through the custom fiberglass flange as in-plane bearing and out-of-plane pull-through demand, and then transfer into the aircraft lower-skin/frame structure. Treating the global aerodynamics, discrete fastener reactions and local laminate response as separate calculation stages keeps the load path traceable while avoiding unnecessary detail in any one model.

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ENGINEERING RATIONALECustom flange used as the integration bridge

The flange provides a continuous load-spreading interface between the commercially supplied radome and the aircraft lower skin, while accommodating manufacturing tolerance, alignment and environmental sealing requirements.

Fiberglass flange layup and radome interface drawing — Fiberglass flange layup and radome interface detail.

Radome mock-up photograph — Radome mock-up used to support integration of the custom flange.

CURED THICKNESS0.192 in

REINFORCEMENTFiberglass Style 7781MIL-C-9084

MATRIXEPOCAST 50A + 9816Manufacturer mixing/cure schedule

LAYUP16 plies0/45/90/90/45/0/0/45/90/90/45/0/0/90/90/0°

MATERIAL-SELECTION RATIONALE

Fiberglass supports the electromagnetic function of the radome

The source identifies favorable dielectric behavior as a key integration benefit: the flange provides structural continuity without introducing the radar attenuation concerns associated with a metallic interface.

MATERIAL PROPERTIES

ENVIRONMENTAL BASIS

Hot-wet strength set

Because the fabrication condition is unknown, the source adopts the conservative 160°F / 95% RH condition as the laminate strength basis.

UNCERTAINTY ALLOWANCE

Additional 25% knockdown

The already conservative source allowables are reduced a further 25% to account for manufacturing and final-laminate uncertainty.

F_tu132.11 ksi

F_tu225.75 ksi

F_cu136.14 ksi

F_cu236.14 ksi

F_su128.44 ksi

Strength basis: [Analysis & Design of Composite & Metallic Flight Vehicle Structures by Richard Abbott], aligned by the source with the lower bound of MIL-HDBK-17-2F.

ILSS BASIS

6.5 ksi typical → 4.875 ksi design value

The source treats glass/epoxy interlaminar shear as comparable to a carbon laminate using the same resin and then applies the same 25% uncertainty reduction to the 6.5 ksi typical value.

WHY THIS IS CONSERVATIVE

Lower-bound manufacturing sensitivity retained

ILSS is sensitive to cure quality and laminate processing. Using the knocked-down value deliberately avoids crediting a best-case fabrication state that is not documented.

Composite material strength reference data — Typical composite strength data across cure and post-cure conditions [Composite Materials Handbook-MIL-HDBK-17-1F - 5F- Appendix A].

The source data below shows ILSS values spanning approximately 5.5–10.4 ksi; the adopted 4.875 ksi value remains below that range after the explicit uncertainty reduction.

JOINTS ALLOWABLE

Composite-joint capacity is treated as a failure-mode problem rather than an isolated fastener-strength check. The governing allowable is taken from the weakest applicable mechanism for the actual fastener/laminate geometry.

Joint allowable philosophy

Separate the in-plane and out-of-plane load paths, then govern each with the minimum applicable capacity.

P_{in-plane}=\min(P_{fastener\ shear},P_{laminate\ bearing})

P_{out-of-plane}=\min(P_{fastener\ tension},P_{pull-through})

COMPOSITE-JOINT METHODOLOGY Stay inside the demonstrated test envelope.

The baseline glass-laminate bearing value is not used blindly. Thickness-to-diameter, edge-distance-to-diameter, fastener diameter and reference-laminate thickness are checked first; only the applicable knockdowns are then carried into the joint allowable. This prevents the analysis from crediting bearing strength that is unsupported by the cited test configuration.

Composite laminate stacking sequence reference — Published laminate stacking-sequence reference [The Strength of Bolted Joints in Multidirectional CFRP Laminates – Fig. 1]

The source identifies five variables that control single-hole composite bearing response: [The Strength of Bolted Joints in Multidirectional CFRP Laminates – Section 7.2].

Lateral constraintControls local splitting / bearing restraint.

Ply orientationChanges the local load-carrying directions.

0° : ±45° contentInfluences bearing and shear-dominated response.

Stacking sequenceChanges the local through-thickness failure path.

Laminate thicknessDrives D/t sensitivity and bearing capacity.

Joint’s Bearing Strength:

GLASS-FIBER BASELINEF_bru = 40 ksiWorst-environment general value from the source methodology.

Geometry window0.5 ≤ t/D ≤ 2e/D > 3.0

If outside the windowUse source-chart knockdownsScale for smaller fastener diameter / thinner laminate when applicable

MODELING RATIONALEDo not credit composite bearing strength beyond the test envelope.

The source uses chart-based knockdowns when geometry departs from the published test configuration. This keeps the allowable tied to demonstrated behavior rather than extrapolating full 40 ksi bearing strength into a less favorable joint. [Analysis & Design of Composite & Metallic Flight Vehicle Structures by Richard Abbott- Section 12.2.4.2].

The published knockdown charts below are based on 0/±45° tested laminates. Applying them to the present laminate is conservative because the installed layup also contains 90° reinforcement, which adds a load-carrying direction not credited by those charts.

Laminate thickness knockdown-factor flowchart — Knockdown-factor decision process for laminate thickness.

Fastener edge-distance knockdown-factor flowchart — Knockdown-factor decision process for fastener edge distance.

Joint’s Pull-Through Strength:

No fully analytical pull-through solution is credited for the laminated joint. The source therefore uses a conservative semi-empirical method correlated to interlaminar shear strength [Analysis & Design of Composite & Metallic Flight Vehicle Structures by Richard Abbott]. A reinforcement factor of 0.75 is applied for glass fiber.

F_{pull} = \frac{2\alpha t\pi D_{f}\tau_{\max}}{2.9 - 0.018\frac{t}{D_{s}} - 0.51\left( \frac{t}{D_{s}} \right)^{2}}

t

Laminate thickness

D_s

Bolt shank diameter

D_f

Bolt head / washer diameter

\tau_{max}

Interlaminar shear stress

α

Reinforcement factor: 1.0 carbon, 0.75 glass

The source curves show a minimum shear strength at failure of 22 ksi and an ILSS value of 12.29 ksi. The analysis intentionally retains the lower knocked-down value established above: 4.875 ksi.

Bearing stress versus edge-distance ratio chart — Bearing stress at failure versus edge-distance ratio [The Strength of Bolted Joints in Multidirectional CFRP Laminates-Fig.6].

Shear strength versus edge distance chart — Shear strength versus edge distance [The Strength of Bolted Joints in Multidirectional CFRP Laminates- Fig.22].

Bearing strength versus diameter-to-thickness ratio chart — Bearing strength versus diameter-to-thickness ratio [The Strength of Bolted Joints in Multidirectional CFRP Laminates- Fig.20].

Composite joint property chart — Joint-property trends for 0/±45° laminates [The Strength of Bolted Joints in Multidirectional CFRP Laminates- Fig.33].

AN3-13A and AN525-10R16 Fasteners

AN3-13A and AN525-10R16 fastener properties [Technical Data Sheet-NASM3-20, NASM525].

P/N	Type	Head	Size (Callout)(Thread)(Length)	Material	D_s (in)	(F_su)(F_tu) for fastener material [ksi]	(f_su)(f_tu) for fastener in single shear [lb_f]	Thread Standard
AN3-13A	Aircraft Bolt	Hex	(#10‑32)(UNF‑3A)(1.40625)	Non-Corrosion Resistant Steel	0.19	N/R	(2125)(2210)	MIL-S-7742
AN525-10R16	Screw	Recessed	(#10-32)(UNF-3A)(1)	Cadmium Plated Alloy Steel	0.19	(0.6x125)(125)		MIL-S-7742

AN3-13A

Aircraft Bolt

Hex

(#10‑32)(UNF‑3A)(1.40625)

Non-Corrosion Resistant Steel

0.19

N/R

(2125)(2210)

MIL-S-7742

AN525-10R16

Screw

Recessed

(#10-32)(UNF-3A)(1)

Cadmium Plated Alloy Steel

0.19

(0.6x125)(125)

MIL-S-7742

AN3-13AGoverning in-plane: 1,109 lb_fComposite bearing governs over 2,125 lb_f fastener shear.Governing out-of-plane: 1,634.65 lb_fPull-through governs over 2,210 lb_f fastener tension.

AN525-10R16Governing in-plane: 1,109 lb_fSame laminate bearing geometry governs.Governing out-of-plane: 1,634.65 lb_fSame washer / laminate pull-through geometry governs.

AN3-13A Fastener:

AN3-13A - in-plane capacity

Fastener single-shear capacity: 2,125 lb_f. The composite bearing check below establishes whether the laminate reduces the usable joint capacity.

The total accumulated thickness of the laminate is t= 0.192 in.
- t_joint/D_s=0.192/0.19=1.01 → No knockdown factor is required.
The total flange width is 2.70 inches. Hence, the estimated edge distance of the utilized fasteners is e=2.7/2=1.35 in. Conservatively, the edge distance is further reduced by 50%.
- e/D_s=0.675/0.19=3.55 → No knockdown factor is required.
The shank diameter is 0.19 in → a knockdown factor is required.
- The knockdown factor is 0.19/0.25=0.76.
The overall thickness of the tested laminate is 0.118 in. Our incorporated laminate is 0.192 in thick; thicker than the tested laminate. This will increase the failure bearing stress. For conservatism, this minimum value is used.

After the 0.76 diameter knockdown, F_bru = 30.40 ksi and the laminate bearing capacity is 1,109 lb_f. Bearing therefore governs the AN3-13A in-plane joint allowable.

AN3-13A - out-of-plane capacity

Fastener tensile capacity: 2,210 lb_f. Pull-through of the laminate/washer interface is checked next.

The utilized AN970-3 washer has an outer diameter value of 0.875 in [Technical Data Sheet-AN970-3 Washer]. Therefore, the joint’s Pull-Through Strength can be calculated as below:

f_{pull}=\frac{2\alpha t\pi D_f\tau_{max}}{2.9-0.018\frac{t}{D_s}-0.51\left(\frac{t}{D_s}\right)^2}

f_{pull}=\frac{2\pi(0.75)(0.192)(0.875)(4.875\times10^3)}{2.9-0.018\left(\frac{0.192}{0.19}\right)-0.51\left(\frac{0.192}{0.19}\right)^2}

f_{pull}=1{,}634.65\;lb_f

The calculated pull-through capacity is lower than fastener tension, so 1,634.65 lb_f governs the out-of-plane joint allowable.

AN525-10R16 - fastener material capacities

Based on the table below, the AN525-10R16 Fastener is made from Cadmium Plated Alloy Steel which has an Ultimate Shear and Tensile Strength values of F_su=0.6x125=75 ksi [Fastener Design Manual-NASA Reference Publication 1228-Page 21] and F_tu=125 ksi [Technical Data Sheet-NASM525]. Considering the fastener cross-sectional area, the Ultimate Shear and Tensile Loads can be calculated as below:

f_{su} = F_{su} \times A_{s} = 75\pi\ \ \frac{{0.19}^{2}}{4} \times 1000

= 2,126.47\ {lb}_{f}

f_{tu} = F_{tu} \times A_{m} = 125\pi\ \ \frac{{0.1517}^{2}}{4} \times 1000

= 2,259.3\ {lb}_{f}\ \

$A_{s}$ is based on nominal shank diameter; $A_{m}$ is based on the minimum external-thread minor diameter.

The Joint Bearing strength will be the same as calculated before, namely f_bru=1,109 lb_f. Since the Ultimate Shear Load is greater than the Joint Bearing Load, the later will be considered as the in-plane joint allowable.

Moreover, the screw’s Pull-Through Strength will be the same as calculated before, namely f_pull=1,635 lb_f. Since the Ultimate Tensile Load is greater than the Pull-Through Strength, the later will be considered as the out of plane joint allowable.

Summary

Both attachment types are governed by the same composite interface capacities, summarized below.

Governing ultimate joint capacities.

Joint	In Plane Strength [lb_f]	Out of Plane Strength [lb_f]
AN3-13A Fasteners’ Joints	1,109	1,634.65
AN525-10R16 Fasteners’ Joints	1,109	1,634.65

AN3-13A Fasteners’ Joints

1,109

1,634.65

AN525-10R16 Fasteners’ Joints

1,109

1,634.65

LOADS

IDEALIZATIONInverted 2D wingRadome fairing treated as a downward-lift surface.

SEA-LEVEL DENSITYρ = 0.00238 slug/ft³

MAX OPERATING SPEEDV_MO,max = 242 kt408.45 ft/s

AERODYNAMIC CENTER1/3 chordMid-height in z

LOAD-CASE SELECTION RATIONALE

Bound the external installation with the most adverse aerodynamic operating condition used by the source.
The assessment uses sea-level density and maximum operating speed, conservatively sets the radome lift coefficient equal to the aircraft-wing cruise value, applies the cited maximum drag coefficient, and converts both aerodynamic forces to ultimate demand with a 1.5 factor. Level fuselage attitude is retained for the downward-lift bound because increasing positive angle of attack reduces the effective lift of the inverted fairing, while a significant negative fuselage angle is not considered attainable.

AERODYNAMIC RATIONALEUse the aircraft-wing cruise lift coefficient as a conservative radome upper bound.

The source assumes the radome cannot be more aerodynamically efficient than the aircraft wing. The wing cruise coefficient is therefore used directly as the maximum radome lift coefficient, then the resulting force is converted to ultimate with a 1.5 factor.

Radome fairing dimensions and assumed aerodynamic-center location.

Aerodynamic Center (in)	Dimensions (in)
X_c.g.	Y_c.g.	Z_c.g.	L	W	D
41.33	0	-13.985	124.33	65.72	27.97

X_c.g.

Y_c.g.

Z_c.g.

41.33

-13.985

124.33

65.72

27.97

Radome fairing dimensions and aerodynamic center — Radome fairing dimensions and aerodynamic-center definition.

LIFT BOUND

Horizontal fuselage attitude governs the inverted fairing

The source treats increased positive fuselage angle of attack as reducing the fairing’s effective lift. A significant negative fuselage angle is not considered attainable; level orientation therefore bounds the downward-lift case.

DRAG BASIS

C_D,max = 0.051

Maximum radome drag coefficient is taken from the cited Horner drag data [Fluid-Dynamic Drag-Horner, page 8-5, Fig. 11].

The wing cruise lift coefficient is first established from aircraft lift equilibrium:

C_{L,wing}=\frac{34,500}{\frac12(0.00238)(408.45)^2(604)}=0.2877

C_{L,rad,max}=C_{L,wing}=0.2877

Using the radome planform and the source planform factor of 0.6:

L_{max}=\frac12\rho V^2SC_L=1,944.66\ lb_f

L_{ult}=1.5(1,944.66)=2,916.99\ lb_f

The same dynamic-pressure basis is applied to drag using $C_{D,max}=0.051$ :

D_{max}=\frac12\rho V^2SC_D=129.248\ lb_f

D_{ult}=1.5(129.248)=193.87\ lb_f

ULTIMATE AERODYNAMIC LOADSBounding global demand

Downward lift 2,916.99 lb_f

Drag 193.87 lb_f

Factor 1.5

STATIC ANALYSIS

Fasteners' Reaction Forces

BOUNDING ATTACHMENT MODELCredit only eight extreme attachment bolts.

Aerodynamic force and moment are reacted at selected FWD, starboard, AFT and port extreme locations using 3D rigid-body analysis. Concentrating the global load into this sparse attachment set is intentionally conservative relative to distributing it around the full perimeter.

Radome eight-bolt attachment reaction model — Eight extreme attachment locations used for the bounding 3D rigid-body reaction model.

Shear and axial reactions [lb_f] at the assumed main attachment points.

Fastener No	Coordinates	f_x	f_z	Shear	Axial
1	0, 0, 0	-34.2	461.2	34.2	461.2
2	17.47, 16.43, 5	-24.2	426.5	24.2	426.5
3	34.94, 32.86, 10	-14.3	391.7	14.3	391.7
4	79.63, 16.43, 5	-24.2	302.8	24.2	302.8
5	124.33, 0, 0	-34.2	213.8	34.2	213.8
6	79.63, -16.43, 5	-24.2	302.8	24.2	302.8
7	34.94, -32.86, 10	-14.3	391.7	14.3	391.7
8	17.47, -16.43, 5	-24.2	426.5	24.2	426.5

GOVERNING ATTACHMENT DEMANDExtreme-bolt reaction

Max shear 34.2 lb_f

Max axial 461.2 lb_f

PASS

{M.S}_{s} = \frac{1,109}{34.2 \times 1.15} - 1

{M.S}_{t} = \frac{1,634.65}{461.2 \times 1.15} - 1

\boxed{{M.S}_{s} = 27.2\ \ }

→ PASS

\boxed{{M.S}_{t} = 2.08\ \ }

→ PASS

Flange

METHOD

Classical Laminate Plate Theory

The 16-ply flange is represented through the ABD stiffness matrix, with ply-level direct, bending and total stresses recovered through the laminate.

FAILURE CRITERIA

Tsai-Wu · Hill · Hoffman

The source spreadsheet compares ply stresses against multiple composite failure criteria, allowing the minimum margin to identify the governing interaction.

The flange input is taken at the attachment location with the highest applied load. The governing row is highlighted below.

Applied force [lb_f] and moment [in-lb_f] at each assumed attachment; $f_z$ is applied at the 2.7 in outer flange edge. The highlighted row represents the governing attachment demand.

Location No	f_x	Moment (M_y) due to f_z
1	34.2	461.2x2.70=1,245.24
2	24.2	426.5 x2.70=1,151.55
3	14.3	391.7 x2.70=1,057.59
4	24.2	302.8 x2.70=817.56
5	34.2	213.8 x2.70=577.26
6	24.2	302.8 x2.70=817.56
7	14.3	391.7 x2.70=1,057.59
8	24.2	426.5 x2.70=1,151.55

Critical flange plate element — Critical 3.7 × 2.7 in flange plate element at the governing leading-edge attachment.

CRITICAL-ELEMENT RATIONALEAnalyze the flange where local attachment demand is highest.

A 3.7 × 2.7 in plate element at the leading edge is isolated because that location carries the governing fastener reaction. Drag and the attachment moment are converted into running loads over the element dimensions for CLPT input.

The following outputs document the laminate definition and ABD stiffness, followed by ply-level stress recovery and the governing failure-criterion margin.

Laminate properties and ABD matrix — Laminate input, material properties and ABD stiffness matrix used in the CLPT analysis.

Ply stresses and laminate margins — Ply direct, bending and total stresses with the minimum laminate margins.

FLANGE GOVERNING RESULTHill criterion

Minimum margin 13%

Material knockdown 25%

PASS

The reported positive margin is obtained after the 25% laminate-property knockdown used to cover source manufacturing uncertainty.

A local 2.5 × 2.5 in interface element is then used to resolve the combined lift/drag shear transferred between the radome and flange.

Local radome-to-flange shear element — Local 2.5 × 2.5 in radome-to-flange element used for the interface shear check.

{f_{s,x} = \frac{D}{{Height}_{Radome\ }} \times Z_{Element\ } }{= \frac{193.87}{27.97} \times 2.5 = 17.33\ {lb}_{f}}

{f_{s,z} = \frac{L/2}{{Length}_{Radome\ }} \times X_{Element} }{= \frac{2,916.99/2}{124.33} \times 2.5 = 29.33\ {lb}_{f}}

f_{s} = \sqrt{{17.33\ }^{2} + {29.33}^{2}} = 34.06\ {lb}_{f}

The resultant shear is converted to average interface shear stress over the selected area:

F_s=\frac{34.06}{2.5\times2.5}=5.45\ psi

Interface shear - PASS5.45 psi demand versus 270 psi minimum EPOCAST 50-A1 lap-shear strength [Epocast 50-A1 Resin - Hardener 9816].

All other structural components

INSPECTION CHECKRemaining fitting and frame structure

The source closes the remaining components by inspection because the attachment tension and shear reactions are low; consequently, the loads passed through the tension-fitting sub-assembly into the radome frames are also small.

PASS BY INSPECTION

REFERENCES

Composite Materials & Analysis Methods

Composite Materials Handbook - MIL-HDBK-17 / CMH-17
Analysis & Design of Composite & Metallic Flight Vehicle Structures - Richard Abbott
The Strength of Bolted Joints in Multidirectional CFRP Laminates

Joint & Fastener Data

Fastener Design Manual - NASA Reference Publication 1228
NASM3-20 technical data
NASM525 technical data
AN970-3 Washer technical data

Aerodynamic Load Basis

Fluid-Dynamic Drag - S. F. Horner

Composite Fabrication & Material Data

MIL-C-9084 - Style 7781 fiberglass fabric specification
EPOCAST 50-A1 / Hardener 9816 manufacturer technical data