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Starship-class 2000kg satellite

This section defines a baseline “Falcon-class” LiDAR Earth-observation satellite: a smallsat in the few-hundred-kilogram regime, sized to be compatible with Falcon 9 rideshare economics and manufacturable by multiple commercial smallsat vendors. The goal is not to present a final design, but to establish a credible reference architecture that can be used consistently in the laser/coverage/economic equations that follow (and then stress-tested against larger “Starship-class” options).

The architecture is intentionally conservative and is aligned with the feasibility framing in Hancock et al. (2021), which derives the key lidar-coverage equations and evaluates plausible mission performance from in-orbit technology characteristics.

Reference form factor and mission class

Source: https://www.exolaunch.com/news_76.html

Smallsat size class (≈250 kg)

We assume a spacecraft in the ~250 kg class, broadly comparable to current high-cadence commercial Earth-observation satellites. Planet’s next-generation Pelican spacecraft is in this general regime (around ~200 kg class) and operates in low Earth orbit; its public documentation and third-party mission descriptions support the idea that this is a realistic “state-of-the-art” size band for capable EO platforms.

Why this matters: this class is large enough to host meaningful optics, power, and pointing stability - while still small enough to be launched frequently at low marginal cost via rideshare.

Launch: Falcon 9 rideshare (“Falcon-class”)

SpaceX’s Smallsat Rideshare Program provides published pricing that scales with mass (a base $325,000 price for 50 kg with an additional $/kg rate), making the economics of placing a few-hundred-kilogram satellite into SSO/LEO relatively predictable.

Orbit and coverage assumptions

Orbit type

We assume a sun-synchronous LEO (SSO) in the ~450 km altitude band for the baseline architecture, consistent with modern EO constellation orbits (Pelican is described as operating around ~475 km SSO and planned specs cite ~350–450 km bands).

Why this matters: altitude drives (i) required pulse energy via range loss, (ii) ground spot size for a given divergence and optics, (iii) accessible swath/coverage geometry, and (iv) revisit cadence.

Payload concept: split-aperture LiDAR instrument

“2×Pelican” conceptual layout (transmit + receive separation)

The baseline concept is a simple split between transmit and receive functions: one “side” of the spacecraft hosts the laser transmitter optics, the other hosts the receiver telescope and detector chain. This is conceptually similar to two small EO satellites fused together to allocate volume and thermal/power separation for a demanding active payload.

Rationale:

• Separates thermal and electrical noise sources (laser driver) from sensitive detection electronics
• Simplifies optical packaging and baffle design
• Enables a larger receiver telescope in a smallsat envelope

Receiver telescope

Baseline assumption: ~0.5 m class receiver aperture (circular equivalent), consistent with the kind of optics you can plausibly package into a few-hundred-kg satellite (and aligned with the kind of feasibility trade studies conducted in the global lidar literature). Hancock et al explicitly parameterize coverage as a function of aperture, altitude, efficiency, and resolution in their framework.

Bus selection: commercially available 200kg+ class platform

A realistic way to anchor power and subsystem budgets is to reference an existing commercial bus family. Dragonfly Aerospace’s bus line is publicly specified as supporting missions up to the few-hundred-kilogram scale, with published mass/payload accommodation language and a stated intent to support more power-intensive payloads.

For smaller members of the family (ηDragonfly), publicly listed electrical specs include orbit-average power ranges and a payload power allocation band, which is useful as a conservative baseline for what “continuous” power might look like in the smallsat regime.

Subsystem requirements driven by LiDAR

Even before detailed laser sizing, the architecture implies several non-negotiable spacecraft capabilities:

Attitude control and pointing

A nadir-pointing lidar needs:

• Stable pointing to keep the beam on-target
• Knowledge/control sufficient to geolocate returns consistently
• Agility to gate firing over target regions (e.g., forests) while conserving power elsewhere

High-agility EO satellites (like Pelican) emphasize rapid slewing and 3-axis stabilization in public documentation, which supports the plausibility of this control performance in the same size class.

Electrical power system

Active lidar is power-hungry because the spacecraft must supply:

• Laser electrical input power (wall-plug → optical)
• Detector chain and timing electronics
• Onboard processing and data handling

Published orbit-average power figures for smallsat buses in this general ecosystem provide a realism check that continuous, kilowatt-class payload power is not typical in the smallest buses, motivating duty-cycle strategies (burst operation) in the Falcon-class concept. For 1500W of solar array peak power, we will assume 400W of orbit average power is available for the LiDAR payload.

Thermal management

A pulsed laser plus its power electronics impose a thermal load that is more like a “radar-class” payload than an optical camera. The Falcon-class architecture therefore assumes:

• Dedicated radiators for payload electronics
• Thermal isolation between laser driver and detector
• Burst-mode operation to manage peak dissipation

Data handling and downlink

Even a narrow-swath lidar can produce large volumes of waveform/return data. The baseline architecture assumes:

• High-speed payload interface to onboard storage
• A high-rate downlink (e.g., X-band or optical, depending on mission constraints)

(We treat downlink as a design dimension to be traded later, because lidar operational concepts can range from full waveform storage to onboard feature extraction and compression.)

Cost breakdown

We consider a 3 year design life, in line with IceSat-2. Based on this, we can explore the full mission cost below. But to summarise, we estimate that a 250kg EO LiDAR mission can be executed for <$20M recurring cost, excluding R&D and non-recurring engineering. This is made possible by adopting a production-oriented payload design, protoflight test philosophy, lean operations (”fly by exception”), and self-insurance.

This is an aggressive cost target for this size of satellite. This works only if all of the following are true:

1. Payload design is mature and frozen
2. Protoflight testing is accepted
3. No launch or on-orbit insurance
4. Lean ops model (no science-mission staffing, no 24/7 ops)
5. Mission failure is tolerated

This is how modern EO constellations achieve cost control.

Spacecraft & payload hardware

The LiDAR payload is the key cost driver here. We assume the cost of a production instrument, not a science payload.

Assembly, integration & test (AIT)

Here we allow reduced AIT testing, assuming a risk-tolerant approach.

Launch & integration

Ground segment & operations (3 years)

We assume no need for 24/7 ops here. Automation is a must, with human-in-the-loop operations only for exceptional situations.

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Regulatory, compliance & programmatic

Costs here are assumed to be shared across a constellation, not just one single satellite.

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Insurance & contingency

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Baseline architecture summary