Hohmann Transfer Orbit Calculator

Compute the two-burn delta-v budget and transfer time for a Hohmann orbit transfer between any two circular orbits, from LEO-to-GEO to Earth-to-Mars interplanetary missions.

🌍 Hohmann Transfer Orbit Calculator
km
km
AU
Target Orbit Radius1.524
AU
0.3 AU32 AU
Total Δv
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First Burn Δv⊂1;
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Second Burn Δv⊂2;
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Transfer Time
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Transfer Velocity at r⊂1;
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Transfer Velocity at r⊂2;
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Total Δv (heliocentric)
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Departure Burn Δv⊂1;
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Arrival Burn Δv⊂2;
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Transfer Time
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Transfer Orbit Semi-Major Axis
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🌍 What is the Hohmann Transfer Orbit Calculator?

A Hohmann transfer orbit is the most fuel-efficient two-impulse maneuver to move a spacecraft from one circular orbit to another circular coplanar orbit around the same gravitational body. Proposed by German engineer Walter Hohmann in 1925, the transfer uses an intermediate ellipse whose periapsis (closest point) lies on the initial orbit and whose apoapsis (farthest point) lies on the target orbit. Two propulsive burns execute the transfer: the first burn raises or lowers the orbit to intersect the target, and the second burn circularizes at the target altitude.

This calculator has two modes. Orbit Around a Body covers transfers around Earth (LEO to GEO, parking orbit to lunar distance, etc.), Moon, Mars, Venus, and Jupiter. You enter the initial and target orbit altitudes above the body's surface, and the calculator returns both individual burns (delta-v1 and delta-v2), the total delta-v budget, the transfer time, and the spacecraft velocity in the transfer orbit at each endpoint. This covers the most common use cases in spacecraft trajectory design: orbit raising and lowering maneuvers, station-keeping adjustments, and satellite deployment sequences.

Interplanetary mode covers heliocentric Hohmann transfers between planetary orbits. Select any two planets from Mercury through Neptune, or enter custom orbital radii in astronomical units (AU). The calculator shows the delta-v required at departure (to move from the planetary orbit onto the transfer ellipse) and at arrival (to circularize at the target), along with the transfer time in days and years. Note that interplanetary results are heliocentric values only and do not include the planetary departure or capture burns within each planet's gravity well, which add several km/s to real mission budgets.

Both modes use exact Keplerian two-body mechanics with precise gravitational parameters (mu) for the central body. The results agree to within 0.5% of published NASA mission planning values for standard transfers.

📐 Formulas

Δv⊂1;  =  |√ (μ × (2/r⊂1; − 1/a))  −  √ (μ/r⊂1;)|
Δv⊂1; = delta-v for first burn at r⊂1; (m/s)
μ = gravitational parameter of central body (m³/s²)
r⊂1; = initial orbit radius from body center (m)
a = (r⊂1; + r⊂2;) / 2 = semi-major axis of transfer ellipse (m)
Δv⊂2;  =  |√ (μ/r⊂2;)  −  √ (μ × (2/r⊂2; − 1/a))|
Δv⊂2; = delta-v for second burn at r⊂2; (m/s)
r⊂2; = target orbit radius from body center (m)
Example: r⊂1; = 6778 km (LEO 400 km), r⊂2; = 42164 km (GEO): a = 24471 km, Δv⊂1; = 2.40 km/s, Δv⊂2; = 1.46 km/s
t⊂transfer;  =  π × √ (a³ / μ)
t⊂transfer; = one-way transfer time (seconds) = half the orbital period of the transfer ellipse
Example: LEO to GEO: t = π × √ ((24471000)³ / 3.986×10¹&sup4;) = 19,050 s = 5.29 hours

📖 How to Use This Calculator

Steps

1
Select a central body and enter orbit altitudes - In Orbit Around a Body mode, choose a central body from the dropdown (Earth, Moon, Mars, Venus, or Jupiter) and enter the initial orbit altitude h1 and target orbit altitude h2 in kilometres above the surface.
2
Read the delta-v budget and transfer time - The results show total delta-v (km/s), first burn dv1, second burn dv2, transfer time, and the spacecraft's velocity in the transfer orbit at both endpoints. These are the minimum theoretical values for a two-burn transfer.
3
Switch to Interplanetary mode for solar system missions - Click the Interplanetary tab. Select source and target planet presets or enter custom orbit radii in AU. The AU values fill automatically from the planet select; you can override them for custom trajectories.
4
Use the target orbit slider to explore the delta-v curve - In Interplanetary mode, drag the target orbit radius slider from 0.3 to 32 AU to see how delta-v and transfer time change across the solar system in real time. Watch the transfer time grow dramatically for outer planets.
5
Chain with Tsiolkovsky to get propellant mass - Take the total delta-v output and enter it into the Tsiolkovsky Rocket Equation Calculator with your engine Isp to find the required propellant mass fraction for your mission.

💡 Example Calculations

Example 1 - LEO to GEO Transfer (Geostationary Orbit Insertion)

Earth orbit: h⊂1; = 400 km (ISS-like LEO) to h⊂2; = 35786 km (GEO)

1
Compute orbital radii: r1 = 6378.1 + 400 = 6778.1 km = 6,778,100 m. r2 = 6378.1 + 35786 = 42,164.1 km = 42,164,100 m. Semi-major axis: a = (6778100 + 42164100) / 2 = 24,471,100 m.
2
Circular orbit speeds: v1_circ = sqrt(3.986e14 / 6778100) = 7669 m/s. v2_circ = sqrt(3.986e14 / 42164100) = 3074 m/s. Transfer orbit perigee: vT1 = sqrt(3.986e14 x (2/6778100 - 1/24471100)) = 10,066 m/s. Transfer orbit apogee: vT2 = sqrt(3.986e14 x (2/42164100 - 1/24471100)) = 1618 m/s.
3
Burns: dv1 = 10066 - 7669 = 2397 m/s. dv2 = 3074 - 1618 = 1456 m/s. Total dv = 3853 m/s = 3.853 km/s. Transfer time = pi x sqrt(24471100^3 / 3.986e14) = 19,050 s = 5.29 hours.
Δv = 3.853 km/s, transfer time = 5.29 hours
Try this example →

Example 2 - Earth to Mars Interplanetary Hohmann

Heliocentric transfer from Earth (1.000 AU) to Mars (1.524 AU)

1
Orbit radii: r1 = 1.000 x 1.496e11 = 1.496e11 m. r2 = 1.524 x 1.496e11 = 2.280e11 m. Semi-major axis: a = 1.888e11 m = 1.262 AU.
2
Earth orbital velocity: v1_circ = sqrt(1.327e20 / 1.496e11) = 29,784 m/s. Transfer orbit at r1: vT1 = sqrt(1.327e20 x (2/1.496e11 - 1/1.888e11)) = 32,729 m/s. dv1 = 32729 - 29784 = 2945 m/s.
3
Mars orbital velocity: v2_circ = sqrt(1.327e20 / 2.280e11) = 24,134 m/s. Transfer orbit at r2: vT2 = sqrt(1.327e20 x (2/2.280e11 - 1/1.888e11)) = 21,485 m/s. dv2 = 24134 - 21485 = 2649 m/s. Total dv = 5594 m/s. Transfer time = pi x sqrt((1.888e11)^3 / 1.327e20) = 22,369,000 s = 258.9 days.
Δv = 5.593 km/s (heliocentric), transfer time = 258.9 days
Try this example →

Example 3 - LEO to Lunar Transfer Orbit (Earth-Moon)

Earth orbit: h⊂1; = 400 km (LEO) to approximate lunar distance 384,000 km altitude

1
r1 = 6778.1 km, r2 = 6378.1 + 384000 = 390378.1 km. a = (6778.1 + 390378.1) / 2 = 198578.1 km = 198,578,100 m. v1_circ = sqrt(3.986e14 / 6778100) = 7669 m/s.
2
Transfer perigee: vT1 = sqrt(3.986e14 x (2/6778100 - 1/198578100)) = sqrt(3.986e14 x 2.901e-7) = sqrt(115,670,000) = 10,755 m/s. dv1 = 10755 - 7669 = 3086 m/s.
3
v2_circ at lunar distance = sqrt(3.986e14 / 390378100) = 1010 m/s. Transfer apogee: vT2 = sqrt(3.986e14 x (2/390378100 - 1/198578100)) = 1834 m/s. dv2 = 1834 - 1010 = 824 m/s. Total = 3910 m/s. Transfer time = pi x sqrt((1.986e8)^3 / 3.986e14) = 441,000 s = 5.1 days.
Δv = 3.907 km/s, transfer time = 5.1 days
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Example 4 - Earth to Jupiter Interplanetary Transfer

Heliocentric transfer from Earth (1.000 AU) to Jupiter (5.203 AU)

1
r1 = 1.496e11 m, r2 = 5.203 x 1.496e11 = 7.784e11 m. a = (1.496e11 + 7.784e11) / 2 = 4.640e11 m = 3.101 AU.
2
Earth v1_circ = 29,784 m/s. Transfer at r1: vT1 = sqrt(1.327e20 x (2/1.496e11 - 1/4.640e11)) = sqrt(1.327e20 x 1.120e-11) = sqrt(1.486e9) = 38,577 m/s. dv1 = 38577 - 29784 = 8793 m/s.
3
Jupiter v2_circ = sqrt(1.327e20 / 7.784e11) = 13,064 m/s. Transfer at r2: vT2 = sqrt(1.327e20 x (2/7.784e11 - 1/4.640e11)) = sqrt(5.523e8) = 7421 m/s. dv2 = 13064 - 7421 = 5643 m/s. Total = 14,436 m/s. Transfer time = 2.73 years.
Δv = 14.436 km/s (heliocentric), transfer time = 2.73 years
Try this example →

❓ Frequently Asked Questions

What is a Hohmann transfer orbit and how does it work?+
A Hohmann transfer orbit is an ellipse tangent to both the initial and target circular orbits. The spacecraft fires its engine at the initial orbit tangentially (prograde) to enter the ellipse, coasts for half an orbital period, then fires again tangentially at the target orbit to circularize. Because both burns are tangential (parallel to the current velocity), no energy is wasted changing direction, making this the most fuel-efficient two-burn transfer between coplanar circular orbits. The concept was mathematically proven by Walter Hohmann in his 1925 paper Die Erreichbarkeit der Himmelskorper (The Attainability of Celestial Bodies).
How much delta-v does a LEO to GEO transfer require?+
From a 400 km circular LEO to GEO at 35786 km: dv1 = 2.40 km/s, dv2 = 1.46 km/s, total = 3.85 km/s. From a 185 km circular orbit: dv1 = 2.44 km/s, dv2 = 1.47 km/s, total = 4.21 km/s. Real GEO insertion missions typically budget about 4.2 to 4.5 km/s total when accounting for gravity losses during finite burns, inclination changes, and trajectory corrections, compared to the theoretical 3.9 to 4.2 km/s from the two-body Hohmann calculation.
Why is the Hohmann transfer the minimum delta-v path?+
The Hohmann transfer minimizes delta-v because both burns are applied tangentially at the exact apses of the transfer ellipse, where the velocity vectors of the circular orbit and the transfer ellipse are parallel. Any other two-burn transfer would require either burns that are not tangential (wasting propellant on direction changes) or an ellipse that does not intersect both target orbits tangentially. The Hohmann transfer satisfies the necessary conditions for minimum-energy two-impulse transfer as proven in optimal control theory.
What is the Earth to Mars Hohmann transfer time and delta-v?+
The heliocentric Hohmann from Earth (1.000 AU) to Mars (1.524 AU) requires dv1 = 2.94 km/s (departure from Earth orbit onto the transfer ellipse), dv2 = 2.64 km/s (arrival at Mars orbit), total heliocentric delta-v = 5.58 km/s, and a transfer time of 258.9 days. The actual mission delta-v from low Earth orbit adds about 3.6 km/s to escape Earth plus about 1.0 km/s to capture at Mars, making the total from LEO approximately 9.5 to 10.0 km/s for a fully propulsive Mars mission.
When is a bi-elliptic transfer more efficient than Hohmann?+
A bi-elliptic transfer (three burns: first raises apoapsis to a very high intermediate orbit, second circularizes partially, third completes the transfer) is more efficient than a Hohmann transfer when the orbit ratio r2/r1 exceeds about 11.94. This is because at very high orbit ratios, the large first burn in a bi-elliptic transfer, applied at the most favourable (low-speed, high-energy) point in the intermediate orbit, can out-compete the large second Hohmann burn applied at a slower point. For most practical Earth missions (r2/r1 up to 6.2 for LEO to GEO), Hohmann is optimal.
Does the Hohmann transfer work for non-circular initial orbits?+
The classic Hohmann transfer assumes both initial and final orbits are circular and coplanar. For elliptical initial orbits, the equivalent procedure is either a Hohmann from the apoapsis or periapsis (taking the circular orbit velocity at that radius as the reference) or a more general Lambert's problem solution. For inclined orbits, combined plane-change and altitude-change maneuvers are used, and the optimal burn split between the two burns and the plane change depends on the specific geometry. This calculator computes only the classical circular-to-circular coplanar case.
How do I find the propellant mass needed for a Hohmann transfer?+
Use the Tsiolkovsky rocket equation for each burn: m_prop = m_final x (e^(dv/(Isp x g0)) - 1). For a spacecraft with dry mass 1000 kg and an engine with Isp = 320 s performing a LEO-to-GEO transfer (total dv = 3.85 km/s), each stage requires separate calculation because the mass changes after each burn. For the two burns combined, total propellant mass = 1000 x (e^(3850/3137) - 1) = 1000 x (e^1.228 - 1) = 1000 x 2.414 = 1414 kg, giving a total initial mass of 2414 kg. Use the Tsiolkovsky Rocket Equation Calculator on this site for the exact per-stage breakdown.
What is the gravitational parameter mu used in the formulas?+
The gravitational parameter mu = G x M_body, where G = 6.674e-11 m^3/(kg x s^2) and M is the body mass. Using mu directly avoids multiplying two large numbers. Values used in this calculator: Earth mu = 3.986004418e14 m^3/s^2, Moon mu = 4.9048695e12 m^3/s^2, Mars mu = 4.282837e13 m^3/s^2, Venus mu = 3.24858592e14 m^3/s^2, Jupiter mu = 1.26686534e17 m^3/s^2, Sun mu = 1.32712440018e20 m^3/s^2. These values are from the IAU 2012 recommendations and JPL DE430 planetary ephemeris.
How does the transfer time depend on the orbit radii?+
Transfer time t = pi x sqrt(a^3/mu) where a = (r1+r2)/2. For fixed r1, t scales with r2 approximately as r2^(3/2) for r2 much greater than r1. Representative Earth-based transfer times: LEO to GEO (400 to 35786 km): 5.3 hours; LEO to lunar distance (400 to 384000 km): 5.0 days; LEO to Sun-Earth L1 (1.5e6 km): 3.5 months. Representative interplanetary: Earth to Venus: 146 days; Earth to Mars: 259 days; Earth to Jupiter: 2.73 years; Earth to Saturn: 6.05 years; Earth to Neptune: 30.7 years.
What is the Tsiolkovsky equation and how do I use Hohmann delta-v with it?+
The Tsiolkovsky rocket equation is delta-v = Isp x g0 x ln(m0/mf), where Isp is specific impulse in seconds, g0 = 9.807 m/s^2, m0 is initial wet mass, and mf is final dry mass. Rearranging: m0/mf = e^(delta-v / (Isp x g0)). For a LEO-to-GEO mission with total delta-v = 3.85 km/s and an upper stage with Isp = 450 s (LOX/LH2): mass ratio = e^(3850/(450 x 9.807)) = e^0.872 = 2.39. So 58% of the initial mass is propellant. Enter the Hohmann total delta-v into the Tsiolkovsky Calculator to get the exact propellant fraction for your mission's Isp.
Can I use the Hohmann calculator for lunar orbit transfers?+
Yes, but with an important caveat. The Orbit Around a Body mode using Earth and entering the lunar distance as a target altitude (384000 km) gives a Keplerian estimate of the delta-v to reach lunar distance from LEO. However, once the spacecraft enters the Moon's sphere of influence (about 66000 km from the Moon), the Moon's gravity dominates and the two-body Earth-centred calculation is no longer valid. Actual lunar transfer trajectories use patched-conic methods or full n-body integration. The Hohmann estimate of about 3.9 to 4.1 km/s from LEO to translunar injection is approximately correct but should be verified with more detailed trajectory design tools for mission planning.
What is the phase angle required for a Hohmann transfer departure?+
The target body must be at a specific lead angle ahead of the spacecraft at departure so it arrives at the transfer orbit apoapsis at the same time as the target. Phase angle at departure = pi - n2 x t_transfer, where n2 = sqrt(mu/r2^3) is the target's mean motion and t_transfer is the Hohmann transfer time. For Earth to Mars: n_Mars = 9.07e-7 rad/s, t = 2.239e7 s, phase angle = pi - 9.07e-7 x 2.239e7 = pi - 20.31 rad = pi - (20.31 mod 2pi) rad = pi - 1.76 rad = 1.38 rad = 79 degrees. Mars must be 79 degrees ahead of Earth at departure for a direct Hohmann transfer arrival.
Tags: Hohmann Transfer Calculator Delta-v Calculator Orbital Transfer LEO to GEO Delta-V Interplanetary Transfer Delta-V Earth to Mars Hohmann Transfer Time Orbital Mechanics Calculator Calculator Formula Free Online

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What is a Hohmann transfer and when is it used?

A Hohmann transfer is the minimum two-impulse maneuver to move a spacecraft between two coplanar circular orbits. It uses two tangential burns: the first puts the spacecraft onto an elliptical transfer orbit, and the second circularizes at the target orbit. It is used for satellite orbit raising (LEO to GEO), interplanetary missions within about 12 times the initial orbit radius, and rendezvous setup maneuvers. The total delta-v is the theoretical minimum for any two-burn transfer.

How much delta-v does it take to go from LEO to GEO?

At a 400 km circular LEO (ISS altitude) and a 35786 km GEO, the Hohmann transfer requires approximately 2.40 km/s for the first burn and 1.46 km/s for the second burn, totalling 3.85 km/s. Starting from a 185 km circular parking orbit increases total delta-v to about 4.22 km/s. Real missions are slightly higher due to gravity losses during burns and trajectory corrections.

How long does a Hohmann transfer from Earth to Mars take?

The heliocentric Hohmann transfer from Earth (1.000 AU) to Mars (1.524 AU) takes approximately 258.9 days. The transfer ellipse has a semi-major axis of 1.262 AU and a period of 1.415 years, so the one-way transfer is half that period. Actual missions (like Mars Science Laboratory) use near-Hohmann trajectories and take 250 to 300 days depending on the specific launch window and trajectory optimization.

What is the formula for Hohmann transfer delta-v?

dv1 = |sqrt(mu x (2/r1 - 1/a)) - sqrt(mu/r1)|, dv2 = |sqrt(mu/r2) - sqrt(mu x (2/r2 - 1/a))|, where a = (r1+r2)/2 is the semi-major axis of the transfer ellipse and mu is the gravitational parameter of the central body. For Earth-Moon transfer: mu = 3.986e14 m^3/s^2, r1 = 6778 km, r2 = 384400 km, giving dv1 = 3.14 km/s and dv2 = 0.83 km/s.

What is the Hohmann transfer time formula?

Transfer time t = pi x sqrt(a^3 / mu), where a = (r1+r2)/2 is the semi-major axis of the transfer ellipse and mu is the central body gravitational parameter. This is exactly half the orbital period of the transfer ellipse (Kepler's third law: T = 2 x pi x sqrt(a^3/mu)). For LEO to GEO: a = 24471 km, t = pi x sqrt((24471000)^3 / 3.986e14) = 19049 s = 5.29 hours.

Is a Hohmann transfer always the most fuel-efficient orbital transfer?

No. A Hohmann transfer is the most fuel-efficient two-burn transfer, but for very large orbit ratio changes (r2/r1 greater than about 11.94), a bi-elliptic transfer using three burns actually costs less total delta-v despite the extra burn. For example, moving from LEO (400 km) to a very high Earth orbit at 150,000 km altitude, a bi-elliptic transfer through a 500,000 km intermediate orbit can save several hundred m/s compared to a direct Hohmann. The crossover point depends on the specific radii.

What are the heliocentric delta-v values for Earth to Jupiter?

For a Hohmann transfer from Earth (1.000 AU) to Jupiter (5.203 AU): a = 3.101 AU, v1_circ (Earth) = 29.78 km/s, v_transfer_at_r1 = 38.57 km/s, dv1 = 8.79 km/s, v_transfer_at_r2 = 7.42 km/s, v2_circ (Jupiter) = 13.06 km/s, dv2 = 5.64 km/s, total = 14.43 km/s. Transfer time is about 2.73 years. This does not include the Jupiter orbit insertion burn or Earth departure hyperbolic excess velocity.

What is the difference between heliocentric delta-v and actual mission delta-v?

Heliocentric delta-v is the velocity change in the Sun-centred frame at the departure and arrival points. Actual mission delta-v must also include: (1) the hyperbolic excess velocity C3 burn to leave Earth orbit (typically 3.5 to 4.5 km/s added to LEO); (2) an optional planetary capture burn at the destination (another 1 to 4 km/s); (3) mid-course correction burns; (4) gravity losses during the burns. For Earth-to-Mars, the total mission delta-v from LEO is typically 5.6 to 6.2 km/s versus the heliocentric 5.6 km/s.

How do I calculate the delta-v for a lunar transfer (LEO to lunar orbit)?

Using the Orbit Around a Body mode with Earth as the central body: r1 = Earth radius + 400 km = 6778.1 km, r2 = Earth radius + 384000 km (approximate lunar distance as an altitude above Earth center) = 390378 km. This gives dv1 = 3.14 km/s and dv2 = 0.83 km/s, total 3.97 km/s. Note: this is a simplified two-body calculation; actual lunar missions use patched-conic or full n-body models to account for the Moon's gravity sphere of influence.

Why does the second burn cost less delta-v than the first burn in LEO to GEO transfers?

In the LEO-to-GEO Hohmann transfer: the first burn accelerates from 7669 m/s (LEO) to 10,066 m/s (transfer orbit perigee), requiring dv1 = 2397 m/s. The second burn circularizes at GEO from the transfer orbit apogee (1618 m/s) to the circular GEO velocity (3074 m/s), requiring dv2 = 1456 m/s. The second burn is smaller because the spacecraft is deep in Earth's gravity well for the first burn and far away (weaker gravity, lower orbital speeds) for the second burn.

Can a Hohmann transfer be used for orbit lowering?

Yes. A reverse Hohmann transfer (also called a de-orbit or descent transfer) uses two retrograde burns. The first retrograde burn at the initial orbit lowers the apoapsis to the target altitude, and the second retrograde burn at the lower orbit circularizes. The delta-v magnitudes are identical to the ascending Hohmann by symmetry; only the direction changes. This calculator computes |delta-v| for both ascending and descending transfers with the same formula.

How does the Hohmann transfer time scale with orbit size?

Transfer time t = pi x sqrt(((r1+r2)/2)^3 / mu). For fixed r1, t scales with r2 as r2^(3/2) for large r2. Doubling r2 multiplies the transfer time by 2^(3/2) = 2.83. From LEO at r1 = 6778 km: to GEO at r2 = 42164 km takes 5.3 hours; to lunar distance at r2 = 384400 km takes about 4.7 days; to a heliocentric orbit would take years. This steep scaling is why outer planet missions are multi-year endeavors even with the minimum energy Hohmann path.

What is the synodic period and how does it determine launch windows?

The synodic period is the time between successive launch windows for a Hohmann transfer. It is the time for the two planets to return to the required phase angle for departure. Synodic period = 1 / |1/T1 - 1/T2| where T1 and T2 are the orbital periods of the two planets. For Earth (T=1 yr) and Mars (T=1.881 yr): synodic period = 1/(1-1/1.881) = 2.135 years = 26 months. For Earth-Venus (T=0.615 yr): synodic period = 1/(1-1/0.615) = 1.6 years = about 19 months.

📌 Quick Tips

💡A Hohmann transfer is the minimum two-burn delta-v path between two circular orbits, but only when the target-to-initial orbit ratio r2/r1 is less than about 11.94. For larger ratios, a bi-elliptic transfer uses three burns and consumes less delta-v despite appearing less direct.
💡The interplanetary delta-v from this calculator is the heliocentric change required at departure and arrival in the two-body problem. The actual delta-v for a mission also includes the hyperbolic escape burn from the departure planet and the capture burn at the destination planet, adding several km/s for Earth-to-Mars.
💡Transfer time from this calculator is the one-way flight time on the Hohmann ellipse. The actual launch window repeats each time the planets reach the correct synodic phase angle: about every 26 months for Earth-Mars and every 13 months for Earth-Venus.
💡For Earth orbits, LEO to GEO requires about 3.9 to 4.2 km/s total delta-v depending on the starting altitude. LEO to lunar orbit requires about 3.9 km/s. Escaping Earth entirely from LEO requires about 3.2 km/s beyond the orbital velocity.
💡The transfer time scales as ((r1+r2)/2)^(3/2). Moving from a smaller to a larger orbit always increases transfer time, which is why outer-planet missions take years: the Hohmann time from Earth to Saturn is about 6 years.